North American X-15, X-15A-2

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..._-------------------------_ ... Aerofax Datagraph 2

North American X·15/x·15A·2 by Ben Guenther, Jay Miller, and Terry Panopalis

ISBN 0-942548-34-5

©1985

Aerofax, Inc. P.O. Box 120127 Arlington, Texas 76012 ph. 214647-1105 U.S. Book Trade Distribution by:

Motorbooks International 729 Prospect Ave. Osceola, Wisconsin 54020 Ph. 715 294-2090 European Trade Distribution by:

Midland Counties Publications 24 The Hollow, Earl Shilton Leicester, LE9 7NA, England ph. (0455) 47256

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THE NORTH AMERICAN X·15/X·15A·2 STORY

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The first X-15, 56-6670, was officially rolled-out during ceremonies held at North American's Los Angeles facility on October 15, 1958. Publicity photos, however, had been taken several days earlier in preparation for the various press releases North American assembled for public distribution. As can be seen in this pre-roll-out view, the aircraft was painted semi·gloss black over-all and equipped with a rather large test boom equipped with pitch and yaw vanes.

CREDITS: Aerofax, Inc. and the authors are indebted to the many contributors who assisted in the gathering of information for this Datagraph. Particularly noteworthy were the contributions made by the following individuals: David Anderton; John Armstrong; William Arnold; Maj. William Austin of the USAF Office of Public Affairs; Ted Bear; Bill Beavers; Walter Boyne, Director, National Air & Space Museum; Robert Cooper; A. Scott Crossfield; Rene Francillon, Ph.D.; Richard Hallion, Ph.D.; Chuck Hansen; Wes Henry of the USAF Museum; Cheryl Hortel of the Office of History, Edwards AFB; John Kerr; Gayle Lawson; Dave Menard; Susan Miller; Myrl Morris of the USAF Museum; Susan Motley; Barbara Newton; Bill Pelzer; Ray Petrusch of the USAF Museum; Sue Seward; Lewis Shaw; Eric Simonsen of Rockwell International; Douglas Siowiak; John W. R. Taylor, Editor, Jane's All The World's Aircraft; Richard Uppstrom, Director, USAF Museum; Barbara Wasson; Vivian White of the USAF Museum; James Wogstad; Jim Young of the Office of History, Edwards AFB; and Lucille Zaccardi. Additionally, a special acknowledgement is due John Becker for the use of his paper entitled, "The X-15 Project, Part I-Origins and Research Background", that appeared in the February, 1964 issue of Astronautics & Aeronautics; and also to Gerald Balzer, for the loan of his exceptional X-15 photo collection. We also would like to acknowledge the following organizations: the A. F. Simpson Historical Research Center; Rockwell International; the Defense Audio Visual Agency; and the AF Systems Command.

higher altitudes falling under the categories of hypersonic and near-space, respectively. It was not until the late 1940's that it became apparent to the aerospace industry and various government agencies that it might be possible to build a flightworthy vehicle capable of hypersonic speeds using then-state-of-the-art technology. Until that time, propulsion systems capable of generating the thrust required for such missions had' not been considered possible. Large rocket engines, brought to an early state of maturity by German research during WWII, now permitted such concept studies to be initiated with some hope of success. The catalyst behind these programs, which eventually resulted in the birth of the U.S. and Soviet ICBM hardware programs and a large number of hypersonic boost glide bomber and reconnaissance vehicle paper studies (such as Beil's BOM/), was the prewar and wartime research conducted by German scientists Eugen Sanger and Irene Bredt (who later married Sanger). This pair, during 1944, concluded that a hypersonic rocket-powered aircraft could be propelled into orbit and then safely glided back to earth. Though paper studies exploring the Sanger and Bredt proposal were numerous, none bore fruit and no hardware construction was undertaken. After WWII, the Sanger and Bredt study surfaced in the U.S. It was considered exceptionally radical for its day, but its plausibility was undeniable. Much interest was generated within the U.S. aerospace industry, and many studies now surfaced that had as their origin, the Sanger and Bredt papers.

PROGRAM HISTORY:

During the late 1940's and early 1950's the U.S. aerospace industry assumed that the hypersonic flight· envelope should be reserved for the unmanned missile, and that manned aircraft capable of operating at such speeds (and altitudes) were technically unfeasible. Additionally, there remained in 1950 no great need for a hypersonic research aircraft, and it was assumed that no operational military or civil requirement for hypersonic speed capabilities would be forthcoming in the foreseeable future. Because of this supposed dubious utility, pre-1950's hypersonic research, either ground-based or inflight, was provided little support by the various U.S. government and private aerospace research facilities. The first substantial official support was not, in fact, forthcoming until June 24, 1952, when the powerful National Advisory Committee for Aeronautics (NACA) Committee on Aerodynamics passed a resolution calling for the NACA to "increase its program dealing with the problems of unmanned and manned flight in the upper stratosphere at altitudes between 12 and 50 miles, and at Mach numbers between 4 and 10." This resolution, which was effectively the result of the Sanger and Bredt research, later was ratified by the Executive Committee when it met the

North American Aviation's (now Rockwell International) seemingly inimitable X-15 high-speed, high-altitude' research aircraft remains perhaps the single most successful manned experimental flight test vehicle in U.S. history. During its nearly ten years of active flight test activity, it generated a legacy of research data and aeronautical firsts that remain unparalleled in the world of aerospace research. From materials to propulsion systems, it was a workhorse testbed that either laid or helped lay the foundations for everything from the Rockwell International Space Shuttle to the forthcoming National Aerospace Plane (now officially designated X-30A). The X-15, in turn, owes its origins to the mid-1940's birth of the dedicated NACNNASA, Air Force, and Navy research aircraft programs that led to the Bell X-1 family, the Bell X-2, the Douglas D-558-I, D-558-1I, and X-3, the Lockheed X-7, and numerous other experimental aircraft types. These aircraft, research vehicles, and an assortment of research facilities contributed significantly to the technology base that led inexorably to a program calling for the exploration of the greater speeds and

following month. NACA headquarters then asked its laboratories at Ames, Langley, and Lewis for comments and recofnmendations concerning implementation of this resolution. Interestingly, the resolution was the end product of work initiated by one of the founders of Bell Aircraft Corporation, Robert J. Woods, who had played a key role in the design and development of the Bell X-1, the Bell X-2, and the Bell X-5. Woods, with an obvious vested interest in any future advanced research aircraft programs that might reach fruition under the auspices of the NACA . or any other government agency, had spent several months during late 1951 placing on paper his thoughts concerning the need for a hypersonic research aircraft program. In a document sent to the NACA Committee on Aerodynamics dated January 8, 1952, Woods proposed establishing a special study group within the NACA which would be directed to evaluate and analyze the basic problems of hypersonic and space flight. As a final recommendation (not initially adopted by the Committee) Woods called for a manned hypersonic research aircraft and underscored its feasibility by attaching a preliminary proposal (dated January 18, 1952) by Bell engineer

The North American X-15 mock-up was primarily of wood construction and contained most aircraft systems in full-scale mock-up form.

H-SFRS RESEARCH AIRCRAFT BASIC PROPOSAL

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Launch Aircraft Approx. Dim. Wing: Span - 63.5' Root chord· 21' Tip chord· 10.5' MAC - 15.75' Thickness· 4% Area - 1,000 sq.'

Tail: 1/2 Span - 14' Root chord - 11.3' Tip chord - 5.8' MAC - 7' Thickness· 4%

Fuselage: Length· 85' Max. dia.• 7' Weight: Empty - 26,900 Ibs. Gross - 99,000 Ibs.

Walter Dornberger (a former commander of the Nazi's WWII Peenemunde flight research facility in Germanyand one of the key engineers behind the design and development of the V-2 rocket). The proposed Dornberger hypersonic research aircraft was to be powered by a rocket engine, have variable sweep wings, and be capable of flying at 6,000 feet per second at an altitude of between 50 and 75 miles (understandably, because of the time element involved with the presentation of Woods' report, Dornberger omitted analyses of feasibility in the critical areas of structural heating and stability). Woods and Dornberger were not the only aircraft engineers proposing a hypersonic research aircraft. In fact, during a meeting of the NACA Subcommittee on Stability and Control taking place on June 14/15, 1951, an aerodynamicist from Douglas Aircraft Company, Max Hunter, had suggested that research into speeds between Mach 5 and 10 be initiated to gather information for missile designers. The Woods and Dornberger documents directly led to the NACA releasing two early unsolicited research aircraft proposals. The first, submitted on May 21, 1952, was from Hubert Drake and Robert Carman of the NACA High-Speed Flight Research Station (HSFRS) at Edwards AFB calling for a two-stage system in which a large supersonic carrier aircraft would launch at Mach 3 a small manned second-stage aircraft. The Drake and Carman report stated that by "using presently available components and manufacturing techniques, an aircraft having a gross weight of 100,000 Ibs. could be built with an empty weight of 26,900 Ibs. Using water-alcohol and liqUid oxygen as propellants this aircraft would be capable of attaining Mach numbers of 6.4 and altitudes up to 660,000 feet. It would have a duration of one minute at a Mach number of 5.3. By using this aircraft, an aircraft of the size and weight of the Bell X-2 could be launched at Mach numbers as high as 3 and altitudes to 150,000 feet, attaining Mach numbers up to almost 10 and an altitude of about 1,000,000 feet. A duration of one minute at a Mach number of 8 would be possible." The second report released by the NACA, by David Stone of the Piloted Aircraft Research Division (PARD) and completed in late May, was of somewhat more conservative dimensions. Stone felt that the Bell X-2 could be used in its initial flights to reach speeds apprQaching Mach 4.5 and altitudes near 300,000 feet, if it were

equipped with two JPL-4 Sargeant solid fuel rocket engines-and low thrust rocket engines underneath the tail and at the wing tips for pitch and roll control for reduction of friction-generated heat build-up during reentry. Stone also recommended that a project group be formed that would work out the details of actual hardware research and development, the flight test program, and aircraft systems. In a further response to the Aerodynamics Committee, a study group consisting of C. E. Brown (chairman), W. J. O'Sullivari, Jr., and C. H. Zimmerman, was created on September 8, 1952, at the NACA's Langley facility. This group reviewed the then on-going ICBM-related work. of Convair, the Rand Corporation, and others, quickly endorsed the feasibility of hypersonic and reentry flight in general terms, and identified structural heating as the single most important technological problem remaining to be solved. It reviewed also the earlier proposals of Drake, Carman, and Stone and agreed to endorse the Stone X-2 modification proposal as the one with the most immediate promise. In a more conservative vein, it agreed also that research calling for speeds in excess of Mach 4.5 should be, for the time being, reserved for unmanned rocket-propelled missiles. It released its report on June 23, 1953. By 1954, the modest successes enjoyed by the second generation X-1 's and several other high-speed research aircraft programs had born fruit in the form of increasing political and philosophical support for a more advanced research aircraft program. It now was assumed unequivocally that manned, hypersonic flight was feasible, and that such speeds also would permit research in the void of near-space. The time was finally at hand for the launching of a bold new hypersonic and reentry flight research program. The events through mid-1954 had, in fact, already led to a number of policy decisions, with the majority of these occurring during a meeting of the Air Force's Scientific Advisory Board (SA B) Aircraft Panel during October, 1953. With Chairman Clarke Millikan at the helm, the Panel released a statement declaring that "the time was ripe" for an adva[lced manned research aircraft, and its feasibility "should be looked into". R. R. Gilruth, then an assistant director of Langley and a member of the Panel, played an important role both in establishing the NACA viewpoint and in developing a confluence of opinion between the SAB and the NACA groups.

PROPOSED NACA RESEARCH AIRCRAFT (GENERIC)

The actual orlgms of the NACA work specifically leading to the hypersonic research aircraft program occurred during a meeting of the NACA interlaboratory Research Airplane Panel held in Washington, D.C., on February 4/5, 1954, under the chairmanship of Hartiey Soule (who had directed NACA research aircraft activities in the cooperative Air Force/NACA program since 1946). In addition to Soule, the panel at the time consisted of C. J. Conlan from Langley; L. A. Clausing from Ames; Walter Williams from the Edwards AFB High-Speed Flight Station; W. Fleming from the Lewis Research Center; and C. Wood from NACA Headquarters. The panel concluded that a wholly new manned research vehicle (as opposed to the proposed modified X-2) was needed, and recommended that NACA Headquarters refer the matter to the four NACA research centers (then called laboratories) for detail study of goals and requirements for such a vehicle. Though support for hypersonic research had been minimal, the laboratories at Ames, Langley, and Lewis (the Edwards AFB HSFRC participated at a different level of involvement) had by this time already formed hypersonic aerodynamics groups. Langley, as early as 1945, had initiated preliminary research using its pioneering 13-inch hypersonic wind tunnel, and with this tool, had provided verification of newly developed hypersonic theories while bringing to light such important phenomenon as hypersonic shock-boundary-Iayer interaction. This tunnel later served to test preliminary design configu rations both for the NACA proposals and the manufacturer's proposals leading to the final hypersonic aircraft configuration. In the equally important area of materials and structures for hypersonic flight, Langley had organized a parallel exploratory program as early as the late 1940's. Via this effort, materials and structures optimized for hypersonic aircraft led to the development of processes and designs that would later playa major role in determining the hypersonic research aircraft structural considerations. In responding to the NACA Headquarters request for recommendations relating to a possible new research aircraft, all of the NACA laboratories set up small ad hoc study groups during March, 1954. The Langley group chose to deal with the problem in more depth than the other laboratories, and their initial study, in which a conceptual vehicle was proposed, was followed by a number of supporting research projects specifically aimed at

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establishing feasibility and solving some of the key problems of this vehicle. The members of the Langley team and their fields of specialization were M. A. Faget in propulsion; T. A. Toll in configuration, stability, and control; N. F. Dow in structures and materials; and J. B. Whitten in piloting. All four fell under the direction of John Becker who was then involved in the study's research trajectory and aerodynamic heating considerations. It was not difficult for these five men and their associates to reach a consensus of opinion concerning the objectives of the new aircraft program. Besides the almost mandatory elements of stability, control, and piloting, a fourth objective was outlined that would come to dominate virtually every other aspect of the aircraft's design; it would be optimized for research into the distinctly related fields of high-temperature aerodynamics and high-temperature structures. Thus it would become the first aircraft in which aero-thermo-structural considerations constituted the primary research problemas well as the primary research objective. During 1954, thermodynamics was considered a virtual unknown in the world of high-speed, high-altitude aircraft flight, primarily because ground-based simulation, in wind tunnels, had failed to provide realistic environments for accurate information gathering purposes. The proposed hypersonic research aircraft, it was assumed, would provide a bridge over the huge technological gulf that appeared to exist between laboratory experimentation and actual flight. With its research objectives now essentially determined, the Langley team members turned their attention to the questions of propulsion and launch techniques. Extant rocket propulsion systems had been examined in great detail with the most promising configuration, a collection of four General Electric Hermes rocket engines, being the one most favored (due primarily to the "thrust stepping" option it provided). An on-going and continuous review of this and other propulsion systems, however,

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Douglas had initiated work on a hypersonic research aircraft under the auspices of the Navy and with the chronologically appropriate designator, 0558-111. Data obtained from this program was then utilized as a basis for the Douglas Model 684 proposal (shown in wind tunnel model form).

eventually led to the conclusion that none was really safe enough for a manned research aircraft program. Concerning launch technique, it had become clear, during preliminary studies, that the only viable launch configuration achievable in the immediate future for a manned vehicle was the proven air-drop method developed by Bell for use with the X-1, X-2, and X-9. With this in mind, a number of possible carrier aircraft now were given serious consideration, these including the Boeing B-50, the Boeing B-52, the Convair B-58, and the Convair B-36. The B-50, B-58, and B-36 were eventually rejected, though the B-58 and its unique configuration and extraordinary performance remained in strong contention for quite some time. Eventually, however, the Boeing B-52 became the primary carrier aircraft of choice, this being the result of its performance, its extraordinary lifting power, its reliability, and its vulnerability to major modification. The B-52 provided an effective velocity increment of about 1,300 feet per second over a comparable ground launched single-stage vehicle. This speed, which would permit air-stagnation temperatures of some 4,000° F., and which was about twice the peak speed of the X-2, was accepted as an adequate hypersonic research aircraft design goal. The basic design of what was to become the X-15 was the end product of choosing a simple wing-body configuration on the grounds that it would minimize transonic and low-speed development and at the same time be entirely adequate for the contemplated aerodynamic and structural experiments in hypersonic flight. During mid-1954, the. tentative . design concept developed by the Langley team was distributed throughout the military services and aerospace industry in the course of numerous presentations and conferences during which the project was discussed. At this stage of the study, the vehicle concept itself was little more than an object of about the right general proportions and the right propulsive characteristics to achieve hypersonic flight.

The hypersonic stability and control properties of such an arrangement were unknown; and even more importantly, there remained unanswered the large question of whether a practical structural design could be found to survive successfully the 4,000° F. air temperatures in reentry research maneuvers at 7,000 feet per second. Because the Langley team felt that any NACA proposal for a major new project of this kind would have little chance for approval and implementation unless the questions concerning structural design and temperature barriers could be overcome based on thorough research and analyses, they organized an exploratory research program which was pushed vigorously to give real substance to recommendations for a conceptual vehicle. The difficult problems of hypersonic stabilization were the first obstacles of really major proportion encountered in the study. Stability difficulties had already been encountered with the X-1 and X-2 at Mach numbers substantially lower than those expected with the proposed hypersonic research aircraft, and it therefore was considered a major challenge to create a solution that would permit stable flight out to Mach 7. The solution finally adopted was devised by C. H. McLellan. His scheme, which was based on theoretical considerations of the influence of airfoil shape on normalforce characteristics, was to replace the time-honored thin supersonic-airfoil section of the tails with a 10° wedge shape. McLellan's calculations indicated that, at hypersonic speeds this shape, while allowing vertical surfaces of considerably smaller proportions than normal, should prove many times more effective than the conventional thin shapes considered optimum for lower Mach numbers-while eliminating the disastrous directional stability decay encountered by the X-1 and X-2. Besides the wedge shape, it also was discovered that the ability to vary the wedge angle was highly desirable. This would permit variation of the stability derivatives, which would provide increased research flexibility. Perhaps more important, it constituted a means of quick

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recovery from a divergent maneuver (such as those affecting the X-1 and X-2). Furthermore, the ability to reenter in a high-drag condition with a large wedge angle greatly extended the range of attitudes for reentry permissible from heating considerations. Orientation of the tail surfaces (Le., X vIs, +) also presented new problems. At hypersonic speeds the flow many chord lengths behind a lifting wing is characterized by sharply divided regions in which the dynamic pressure may be either several times larger or much smaller than that of the free flow, together with large changes in flow angularity. An analytical study supplemented by flowvisualization tests was organized to determine a suitable tail arrangement. Quoting directly from the initial Langley report: "At high supersonic Mach numbers interference effects may exist between the flow field of the wing and the tail surfaces which produce large changes in the stability derivatives". Basically, this stated that the X tail surface configuration initially chosen tended to develop increasing negative dihedral as the Mach number increased, this being caused by the high tail surfaces being in an area of low pressure and low downwash, and the low tail surfaces being in an area of high pressure and high downwash. The solution actually was quite conventional in appearance, though diametrically opposed to extant supersonic design theory. It was discovered that locating the horizontal tail in the plane of the wing, between the regions of high downwash, eliminated the stability difficulties; this contrasting with the then-accepted idea that the horizontal tail should be located far above or well below the wing chord plane in order to avoid, within reason, an undesirable shock wave interaction. In fact, on the final X-15 design, an acceptable location slightly below the wing plane was utilized by North American. Through use of differential operation of all-moving horizontal tails, North American also eliminated the need for ailerons-which were a potential source of shockinteraction difficulties, too. Research flexibility required that the hypersonic research aircraft have the widest possible angle-of-attack range during reentry. An inherent difficulty that threatened to limit the permissible hypersonic lift was discovered in the drastic change in relative directional effectiveness of the upper and lower tails. Quoting again

from the 1954 Langley report: "At high angles of attack, a marked difference in the eHectiveness of the upper and lower vertical tails develops. Effectiveness of the upper tail decreases to zero at about 20° angle of attack. The lower tail exhibits a marked increase in effectiveness as a consequence of its penetration into the region of high dynamic pressure produced by the compression side of the wing. If the wing is assumed to be a flat plate and the flow is taken to be two-dimensional, the dynamic pressure below the wing would increase with angle of attack. Since only a part of the lower tail is immersed in this region its gain in effectiveness is, of course, less rapid, but the gain more than offsets the loss in effectiveness of the upper tail." As first flown, North American's X-15 retained the symmetrical arrangement of upper and lower vertical tails as the most desirable configuration for the boost phase of its flights. The bulk of the remaining program, however, was flown with this lower surface (technically a rudder) removed. This was done in concert with the recommendations of the 1954 Langley paper. By the end of 1954, the majority of the aforementioned Langley studies had been completed. Detailed coverage of these exploratory stability and control investigations was made available to prospective hypersonic research aircraft contractors. A general summary of the various Langley studies was presented to these potential contractors at a bidder's briefing occurring at Langley during January, 1955. A wide variety of flight trajectories also had been, by now, studied and evaluated in the heating and structural parts of the Langley study. A skin temperature limitation of 1,200° F. had dictated many facets of this research, this stemming from the fact that it might be possible for the proposed hypersonic research aircraft to achieve Mach 7.6 if a B-58 were utilized as the carrier aircraft. The altitude mission profile also had been studied in great detail, this being the standard "space" trajectory in which the aircraft would leave the sensible atmosphere for a period of 2 to 3 minutes of weightless flight followed by a steep reentry. Aerodynamic control above 200,000 feet was found virtually to disappear. Hydrogen-peroxide rockets for control about all axes were consequently proposed and analyzed. They were found adequate in all respects, and the hydrogen-peroxide weight required for

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postulated control schedules was shown to be acceptable. Two basic structural design approaches had been debated since the initiation of the study-first, a conventional low-temperature design of aluminum or stainless steel protected from the high-temperature environment by a layer of assumed insulation; and second, an exposed "hot" structure in which no attempt would be made to provide protection, but in which the metal used and the design approach would permit high structural temperatures. It was found from analysis of the heating situation for various trajectories that insulation capable of accommodating temperatures of over 2,000° F. would be required. At that time, there was no known insulating technique that could meet this requirement. The Bell "double-wall" concept, in which a non-load-bearing metal sandwich acts as the basic insulator, would later undergo extensive development, but in 1954, it was in an embryonic state and not applicable to the critical nose and leading edge regions of the proposed research aircraft. Furthermore, it had to employ a supplemental liquid cooling system. The study group felt that the possibility of local failure of any insulation scheme constituted a serious hazard. Finally, the problem of accurately measuring heat-transfer rates-one of the prime objectives of the new research aircraft program-looked basically more difficult to accomplish with an insulated structure. At the start of the study it was by no means obvious that the "hot" structure approach would prove practical, either. The permissible design temperature for the best available material, Inconel-X, was about 1,200° F., which was far below the estimated equilibrium temperature peak of about 2,000° F. for the lower surface of the wing. Thus it was clear that some form of heat absorption would have to be employed-either direct internal cooling or heat absorption in the metal skin itself. It was felt that either solution would bring a heavy weight penalty. The availability of Inconel-X and its exceptional strength at extremely high temperatures, made it, almost by default, the preferred structural material for use in the proposed hypersonic research aircraft. It thus was decided, during mid-1954, that basic design analysis of an Inconel-X structure should be started (neglecting any

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heat-absorption requirements). Concurrently, the heat load would be analyzed and this would then be used to determine a modified structure. When the results of these preliminary analyses were examined, one of the major findings of the study came to light: the stress study indicated that the wing skin-thickness requirements ranged from about 0.05 to 0.10 inches-about the same as the values found necessary for heat absorption in the thermal analysis. Thus it was possible to solve the structural problem for the transient conditions of a Mach 7 aircraft with no serious weight penalty for heat absorption. This was an unexpected plus for the "hot" structure; and together with the fact that none of the aforementioned difficulties of the insulated-type structure were present, caused the study group to decide in its favor. A more detailed analysis of the "hot" structure was therefore undertaken. Unfortunately, it later proved that the "hot" structure had problems of its own, particularly in the area of nonuniform distributions of temperature. Langley thermal analysis studies revealed that large temperature differences between the upper and lower wing skin would develop during reentry pull-up portions of certain trajectories. This unequal heating would result in intolerable thermal stress in a conventional structural design. To solve this new problem, wing shear members were devised which did not offer any resistance to unequal expansion of the wing skins. The wing thus was essentially free to deform both spanwise and chordwise with asymmetrical heating. Large but not prohibitive tip deflections of as much as 15 inches (upward) and camber decreases as large as 2% of the chord were calculated. Although this technique did not induce any gross thermal stresses, local thermal-stress problems existed in the vicinity of the stringer attachments. The study indicated, however, that proper selection of stringer proportions and spacing would produce an acceptable design free from thermal buckling for the transient flight conditions of the proposed' aircraft. Similar difficulties were encountered with the design elements leading to the development of the wing corrugated-web beams and the wing leading edge. Differential heating of the latter was discovered to produce changes in the natural torsional frequency of the wing unless some sort of flexible expansion joint was incor-

porated in its design. The hot leading edge was discovered to expand faster than the remaining structure so that compression was induced in the leading edge. This compression destabilized the section as a whole and reduced its torsional stiffness. To negate this phenomenon, the leading edge later was segmented and flexibly mounted (in a further attempt to reduce thermallyinduced buckling and bending in the wing and elsewhere, North American Aviation, which would eventually become', the prime contractor for the new hypersonic aircraft, also later introduced the use of dissimilar metals to reduce local thermal stresses by taking advantage of their different rates of thermal expansion). While NACA headquarters was receiving the Becker! FageVTolI!DowlWhitten studies, a letter from the NACA was received by Lt. Gen. Donald Putt at Air Force head, quarters stating that the NACA was interested in the creation of a new manned research aircraft program that would explore hypersonic speeds and altitudes well in excess of those presently being achieved by extant research aircraft types. The letter also recommended that a meeting between the NACA, Air Force headquarters, and the Air Force Scientific Advisory Board be arranged to discuss the project. Gen. Putt responded favorably, and also recommended that the Navy be invited to participate. By late July, 1954, the NACA studies, although incomplete at the time, had reached the stage where a con-' vincing case could be made for the desirability and feasibility of a new Mach 7 research aircraft. The first of many presentations of the subject were made on July 9, 1954, at a meeting sponsored by the NACA and attended by representatives of the Air'Force, the Navy, the Wright Air Development Center, the Air Research and Development Command, and the Scientific Advisory Board. Hugh Dryden, Director of Research for the NACA, reported that the NACA believed a new research aircraft was desirable, outlined the reasons for this, and said the time had come to determine whether an area of agreement existed on the objectives and scope of such a project. During the meeting, the history of the research aircraft was reviewed by H. A. Soule and Walt Williams. Following this, the results of the Langley study were presented by J. E. Duberg and John Becker. The meeting conclud-

ed with agreement that a document setting forth the views of the NACA should be circulated and discussed !n detail by appropriate groups throughout the services and industry. During August, 1954, this suggested NACA document was released. It should also be noted that during the July 9 meeting, the Navy disclosed that they had sponsored a preliminary study by the Douglas Aircraft Co. calling for research into the feasibility of designing a piloted aircraft capable of returning safely from 1,000,000 foot altitudes. Douglas, in their report, had concluded that 700,000 foot altitudes would be possible from the reentry deceleration standpoint, but that the thermo-structural problem had not been thoroughly analyzed (interestingly, the Langley study had indicated a peak altitude of 400,000 feet for an aircraft of Inconel-X construction). The July 9 meeting and the resulting release of the NACA study papers served to awaken industry to the seriousness of the NACA's intention. Accordingly, many aircraft manufacturers sent representatives to Langley to become acquainted with the project in greater detail. One of the first to arrive was, in fact, Bell Aircraft Corporation's Robert Woods, who came not only with several staff members, but also with preliminary design proposals. With momentum now gathering, the NACA Aerodynamics Committee held a historic meeting on October 5, 1954, to consider the question of a hypersonic research aircraft. During the meeting, historic and technical data were reviewed by various committee members including D. Beeler, test pilot A. Scott Crossfield, and Walt Williams. Though at least one Committee member expressed strong opposition to the proposed hypersonic research aircraft as an extension to then extant on-going programs, the rest of the Committee unanimously endorsed the project.·Accordingly, Hugh Dryden, now with a formal technical endorsement in hand, initiated a joint effort with the Air Force and Navy to produce a suitable technical specification on which manufacturers could base proposals. As finally adopted, the specification called for a design speed of 6,600 fps and an altitude of 250,000 feet. The "Requirements" of the specification were written to conform to the philosophy developed in the original Langley study-but they were made sufficiently general to encourage fresh approaches. Appended to the specification under the heading of "Suggested

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Means of Meeting the General Requirements" was a section outlining the key results of the Langley study. A formal presentation of the proposal and the specification was made by the NACA team (Soule, Becker, and Dow) on December 14, 1954, for the Defense Department's Air Technical Advisory Panel. The specification was approved with the stipulation that the NACA should have technical control and that the Panel would have an opportunity to review the manufacturer's proposals. As a final step, a "Memorandum of Understanding" was signed on December 23, by the Air Force, the Navy, and the NACA. This agreement set up a "Research Airplane Committee" responsible for technical direction of the project, with Hugh Dryden as chairman. On December 30, steps to initiate a design competition were taken by the Air Materiel Command when invitation-tobid letters were sent to twelve prospective airframe contractors. On January 17, 1955, NACA representatives met with AF personnel at Wright-Patterson AFB and were informed that the research aircraft would carry the AF Project 1226 identifier and would be officially designated X-15. The Air Force also informed the NACA that the WADC Project Engineer would be Lt. Chet McCollough, Jr. The NACA subsequently informed the Air Force that its project engineer would be A. W. Vogeley. These two men would serve as the conduit for information transfer and problem rectification. A bidders' briefing was held on January 18, at which time the technical material was reviewed for contractor benefit. Representatives from Bell Aircraft Corp, Boeing Airplane Co., Chance Vought Aircraft, Inc., Consolidated Vultee Aircraft Corp. (Convair), Douglas Aircraft Co., McDonnell Aircraft Corp., North American Aviation, Inc., Northrop Aircraft Inc., and Republic Aviation Corp., met with the NACA and Air Force personnel to discuss the competition and the basic design requirements. Another briefing, for the four prospective propulsion system manufacturers (Aerojet, General Electric, North American, and Reaction Motors) was held on February 4. During the bidders' meeting, the airframe manufacturers were informed that one prime proposal and one alternate proposal (that might offer an unconventional but superior solution to the problems involved) would be ac-

\-.

An early hypersonic tunnel model of the North American X-IS proposal reveals the design's gently tapering chines which ran almost to the tip of the nose cone, and the hexagonal airfoil section of the vertical and ventral tail surfaces.

cepted from each company on or before the deadline of May 9, 1955. It also was noted that an engineering study, only, would be required for an alternate version in which an observer would be substituted for the research instrumentation (a Navy requirement); that a weight allowance of 800 lbs., a volume of 40 cubic feet, and a power requirement of 2.25 Kw be provided for the research instrumentation normally carried; and that the winning design be built in 30 months and be capable of meeting the baseline performance requirements of 6,600 feet per second speed and 250,000 feet altitude. The winning entry would be chosen by a joint NACAIService panel. Because of the four powerplant contenders, the airframe manufacturers were asked to design their aircraft around anyone of the four following rocket engines: the Bell XLR-81, the Aerojet XLR-73, the Reaction Motors, Inc. XLR-1 0 or XLR-30, or the North American NA 5400. Engine choice would be up to the airframe manufacturer and if the NACAIService panel picked another, this would not affect the final design evaluation. Following the preliminary statements concerning the bidding, NACA personnel briefed the various companies in attendance on new information that had resulted from late 1954 wind tunnel research that had taken place at Langley. Each company was given a pre-release copy of the following NACA Research Memorandums: RM L54L03b; RM L55A21; RM L55A21a; and RM L55B03. Between February and May five of the initial nine airframe contenders dropped out of the competition. Those remaining were Bell, Douglas, North American, and Republic. During this period, representatives from these companies met with NACA personnel on numerous occasions at the different laboratories and there reviewed technical information on various aspects of the forthcoming X-15, The NACA also provided these contractors with further information gained as a result of wind tunnel tests in the Ames 1Q-inch x 14-inch supersonic tunnel and the Langley Mach 4 blowdown jet tunnel. Concurrently, the Air Force and the NACA also were working on the procedures that would be used during the evaluation of the bids. Early in March, the NACA issued a Research Authorization (A73L179) that would cover

NORTH AMERICAN ESO 7487 (TWO-SEAT)

6

•

Langley's work on Project 1226 during the design competition and evaluation. Later during the same month, the NACA created the NACA Evaluation Group that would examine the received bids. The members of this Group included Soule (Chairman), A. W. Vogeley (Executive Secretary), John Becker, H. J. Goett, John L. Sloop, and Walt C. Williams. By May 9, 1955, all the Project 1226 bidding contractors had submitted their proposals. During the week that followed, copies of these various submissions were distributed to the WADC and NACA evaluation groups. On May 17, the Research Airplane Committee, composed of Hugh Dryden, Brig. Gen. Benjamin Kelsey, and Admiral R. S. Hatcher, met at the WADC for individual briefings by the four bidding contractors. Also present were the NACA, WADC, and Navy evaluation groups. Shortly thereafter, Soule, as the Chairman of the NACA evaluation group, sent to the NACA laboratories and headquarters offices summaries of the evaluation rules and processes. The evaluation would be based on the technical and manufacturing competency of each contractor, each contractor's time and cost estimates, and each contractor's design approach and aircraft research utility. In order to expedite the evaluation, each of the NACA laboratories involved were assigned specific items to consider with responses to be returned to Soule no later that June 13: Langley/Ames aerodynamics Langley structures Langley/Ames/HSFRS. .. flight control items Lewis/HSFRS , powerplantlpropulsion HSFRS crew provisions The actual evaluation process was broken down into eight steps: Step I: The WADC and the NACA laboratories would rate manufacturers on assigned phases. Numerical ratings were to be made on a percentage basis and to be presented with comments, explaining the ratings. Step II: The rating with comments would be turned over to the cognizant WADC sections for coordination. The sections would then submit final evaluations on factors to the project office. Step III: The project office would assign weights to the different factors and arrive at the final evaluation on technical competency. Step IV: The project office and the NACA evaluation group would consider all factors and arrive at an overall evaluation of the airframe contractor and prepare a recommendation with supporting date. Step V: A review and critique of the evaluation and recommendations would be made by the higher WADC echelons. Step VI: The evaluation/recommendations would then be submitted to the Air Force through the ARDC. Step VII: The evaluation/recommendation would be submitted to the Research Airplane Committee. Step VIII: The Research Airplane Committee's decision would be reviewed by the Coordinating Committee. The evaluation of the engine would be made at the same time, but would be conducted separate from that of the airframe contractor, with the possibility that the chosen engine might not be the one selected by the win-

ning airframe contractor. As events worked out, this is exactly what occurred. The engine evaluation, with the ultimate awarding of a contract to Reaction Motors, Inc. during September, 1956, caused some difficulty because it was not the powerplant chosen by the winning airframe contractor. Additionally, the three engines that were initially evaluated were found lacking in various degrees of mechanical technology and safety. Therefore, when the final airframe contract was awarded, the powerplant still had not been chosen. In the interim, it was recommended by Walt Williams that the proven Reaction Motors, Inc. LR8, earlier known as the XLR11 and used to power all of the X-1 's and the rocket-propelled versions of the D-558-1I, be used as a substitute propulsion unit until the final powerplant could be developed and tested. Early in June, 1955, both Ames and Langley laboratories completed their tentative evaluations of Project 1226. Ames ranked the submissions as follows: (1) Douglas; (2) North American; (3) Bell; and (4) Republic. The Douglas ranking resulted from "the completeness and soundness of (their) design study, (their) awareness of factors in speed and altitude regime, and (the) relative simplicity of (their) approach." Ames, however, expressed skepticism over the Douglas magnesium heat sink wing structure. North American's design had been ranked very close to that of Douglas, whereas those sumbitted by Bell and Republic were only marginally competitive, at best. The North American proposal had, in fact, been rated number one by the Langley study team, with Douglas number two, Bell number three, and Republic number four. According to the Langley assessment, the research utility of the North American "hot" structure approach was more representative of future aircraft wing construction and outweighed the advantages of the simplicity of the magnesium structure proposed by Douglas. On June 10, the final evaluations from Ames and the HSFRS were sent to Soule. The HSFRS evaluation was based on the design approach and research utility aspects of the airframe, flight control system, propulsion unit, crew provisions, handling and launching, and miscellaneous systems. Their conclusions resulted in the following rankings: (1) Douglas; (2) North American; (3) Bell; and (4) Republic. Again, the two proposals from Douglas and North American were almost equal on the basis of points. The evaluators for the HSFRS wrote that if the Douglas proposal was ruled out, the North American design would win by default. The Ames final evaluation ranked the contractors as (1) North American; (2) Douglas; (3) Bell; and (4) Republic. This represented a change from the earlier Ames evaluation, based largely on the thermal considerations which dictated the choice of North American over Douglas. The North American structure was considered to be more representative of future aircraft and thus superior in terms of "research utility". Douglas retained a simple and conventional structure, but in so doing, avoided the very problems at which the research aircraft was directed. Consideration was now given to recommending that the Douglas proposal be chosen with the provision that the first two aircraft be built as proposed, while the third have a wing based on the alternative "hot" approach. Serious support for this recommendation was not made beca.use sufficient funds to construct an alternate wing were' not considered likely.

On June 14, Langley's final evaluation was submitted. It was based on aerodynamics, structures, stability and control, and research considerations. The overall ranking was (1) North American; (2) Douglas; (3) Republic; and (4) Bell. Langley felt that while the magnesium wing structure of Douglas was feasible. it was feared that local hot spots caused by irregular aerodynamic heating could weaken the structure and be subject to failure. North American's use of Inconel-X was believed an advantage with regards to the 1,200° F. thermal limits that would be seen on the X-15 wing. A few days after having received all the final evaluations, Soule sent copies of each to the WADC project office. The results were summarized as follows: Design Approach Airframe Flight Control Propulsion Crew Provisions Handlingllaunching Miscellaneous

B 70 70 80 55 95 70

0 80 80 80 85 65 85

Research

Utility N 85 75 90 80 75 70

R 75 70 30 40 65 70

B 70 70 75 55 90 70

0 80 75 40 85 70 85

N 90 75 40 80 70 70

R 80 75 75 35 70 70

By the end of the month, Soule sent to the NACA the various laboratories' evaluations of the Project 1226 proposals. The final order representing the overall NACA evaluation was (1) North American; (2) Douglas; (3) Bell; and (4) RepUblic. All of the laboratories involved in this portion of the evaluation considered both the North American and Douglas proposals to be superior to those submitted by Bell and Republic. While North American was chosen because of its more realistic Inconel-X wing structure, the North American design was not without fault. The NACA, for instance, thought that the landing gear arrangement was basically undesirable, that the differentially-operated horizontal stabilator design in lieu of ailerons was an overly complicated arrangement, and that the replaceable fiberglass leading edge was unacceptable. During the first two weeks in July, the WADC evaluation teams sent their final reports to the WADC Project Office. As with the NACA evaluations, the WADC found Iittle difference between the Douglas and North American designs, point-wise, with both proposals significantly superior to those of Bell and Republic. At this time, George Spangenberg, then Bureau of Aeronautics project engineer for the Project 1226 evaluation, wrote a memorandum reviewing the BuAer evaluation. The order of merit with comments was: (1) Douglas-A clear choice for a research tool to accomplish aerodynamic research ai speeds of 6,600 feet per second and at altitudes up to 375,000 feet. (2) North American-A choice to accomplish research with skin temperatures on the order of 1,200° F. This choice was made primarily on the ability of the contractor, a good basic configuration, and the type of structure. (3) Republic-A second choice for high-skin temperature research. (4) Bell-The designer's knowledge of high temperature problems is evident, but the solutions are less than optimum. On July 26/28, the Air Force, Navy, and NACA evaluation teams met at the WADC to coordinate their separate results. The BuAer representative, Spangenberg, listed

Wind tunnel model of a late North American X-15 configuration illustrates how evoiutionary design processes refined the earlier configuration into a more practical aircraft. Noteworthy is the radical wedge shape airfoil of the vertical fin.

the BuAer's choices as (1) Douglas; (2) North American; (3) Bell; and (4) Republic, and also noted that there appeared to be some bias on North American's behalf, particularly when it was realized that the materials choice had been effectively narrowed to Inconel-X. The NACA concluded that the North American proposal accommodated their requirements, but that if it should be found t-y the Air Force to be unsuitable, a Douglas design, reconfigured to accommodate the Inconel-X requirement, would prove suitable. The Air Force assessment agreed completely with that of the NACA. Accordingly, the Navy decided not to be put in the position of casting the dissenting vote and after short deliberation, agreed to go along with the decision of the Air Force and NACA. As a result, the three agencies designated North American as the prime contractor in the Project 1226 competition. During the week of August 1/5,1955, the formal evaluation document and oral presentation was prepared by the WADC project office. The section concerning the evaluation stated, "the evaluation of the proposals submitted in competition was made in five areas: performance, technical design, research suitability, development capability, and cost". These evaluations were as follows: Performance: consisted of a check of the probability of the different designs, considering present uncertainties, and of meeting the specified speed and altitude requirements. The probabilities were calculated to be best for the North American proposal equal for the Bell and Douglas proposals, and least for the Republic proposal; but because of the assumptions of the analysis, all designs were judged able to meet the requirements. Technical Design: was judged on the awareness shown by the contractor of the problems of high-speed, high-altitude flight and of the means, as indicated by the airplane designs, the contractor proposed for exploring and studying these problems. The general design competency of the contractor also was judged from the designs submitted: North American 81.5%; Douglas 80.1 %; Bell 75.5%; and Republic 72.2%. No design, as submitted, was considered safe for the use intended. The Douglas design was considered best in this regard but did not include adequate margins for ignorance factors and operational errors. Research Suitability: The fundamental differences in the proposed structures were examined and rated because of their decisive importance in the research uses of this aircraft. North American was rated acceptable because of the "hot" structure of the Inconel-X heat-sink which was most suitable for research and which was potentially the simplest to make safe for the mission. Republic and Bell were considered unsatisfactory because of the hazardous aspects associated with the insulated structures used, and Douglas was considered unsatisfactory because. of the low safety margins available and because of the limited future usefulness of the "cool" magnesium heat-sink principle. Development Capability: was based on the physical equipment and manpower the contractor had available for pursuing the project. Evaluation of this factor resulted in the following ratings: (1) Douglas was acceptable; (2) North American was acceptable; (3) Bell was less acceptable; (4) Republic was less acceptable. North American, Republic, and Douglas promised that the first flight date would be within 30 months. Bell promised a first flight

The NACA X-15 tunnel test program was comprehensive, including, for instance, a series of tests of a 1120th scale X-15 free-fall model designed to explore X-15IS-52 separation characteristics during launch, and while mated.

7

COMMON PHYSICAL CHARACTERISTICS General Wing span (ft.) Overelll!. (ft.) Overall HI. (ft.) Diameter (ins.) Wings Airfoil section Section thickness RoolTipRool Chord (ft.) Tip Cftord (ft.) Angle 01 incidence Dihedral Sweepback (LE) MAC On5)

Bell (0171)

Douglas (Model 684)

NAA (ES07487)

Republic (AP·76)

25.66 44.66 12.33

19.5 46.7 13.7 62

22.36 49.33 11.4

27.75 53.3 14.9 60

Circular Arc

Clark "Y" mod

66005 mod

Hexagonal mod

5% 6% 13.16 3.86 0 0 40' 112.5

7'/0 4.5% lOA 2.75 0 0 40' 105.25

5% 1% 10.8

5% 7.5% 16 2.25 0 0 38°4(1' lJO.87

3 0 0 (25%) 25' 123.23

Vertical Tail 0

Airfoil section Tip Chord On5) Sweepback (lEl

Circular Arc 54.5 45'

Mod Diamond 21 40'

10 Wedge 29.3 (25%) 21'

12° Wedge 40 27°57'

Ventral Fin Airfo II section Tip Chord Ons) Sweepback (lE)

10° Diamond 84.25 45'

7' Wedge 45

50'

15° Wedge 65.76 (25%) 52'

10 0 Wedge 84 45'

Horizontal Tall Airfoil section Root Chord Ons) Tip Chord Ons) Sweepbacl< (LE) Span (ft.)

Circular Arc 84.5 25.35 35112 0 13.75

5° Wedge 92 20 40' 11.83

66005 mod 84.27 25.28 (25%)45' 17.64

10° Wedge 85 22 22'36' 15.7

date within 40 months, Costs: for three aircraft plus one static test article were: Bell, $36,3-million; Douglas, $36.4-million; Republic, $47-million; and North American, $56,1-million, In the last portion of the report, the Air Force presented its minority report justifying its choice of the XLR73 rocket engine, which had been rejected during the engine evaluation for the simple reason that none of the airframe contractors had called for the engine in their proposal. Also, in this section was the formal NACA recommendation that an interim powerplant, specifically the Reaction Motors LR8-RM-8, be installed in the X-15 for the initial portion of the flight test program until the actual powerplant was chosen and made ready for installation, A formal briefing, based on this report, was presented on August 8 to Gen. Estes, then Chief of the Weapons System Division under whose jurisdiction the project office then fell, and to a select group of Air Force officers, A second presentation was made by Capt. McCollough at Baltimore for Generals Sessums and Demler, who were then the heads of the WADC and ARDC, respectively. A combined meeting of USAF Headquarters, the Bureau of Aeronautics, NACA personnel, and the Research Airplane Committee was held at the NACA's Headquarters on August 12 for the final briefing on the evaluation, Following this, the Research Airplane Committee, composed of Hugh Dryden, Benjamin Kelsey, and R. E. Dixon, met and decided to accept the findings of the evaluation groups and to present the recommendation to the Defense Department. Because the estimated cost was well above the amount allocated for the projet, the Committee included in its final report a recommendation for a funding increase that would be approved before the actual contract was signed, A further recommendation called for the relaxation of the contract time of up to one-and-one-half years, These recommendations were then sent to the Assistant Secretary of Defense for Research and Development. Events now took a bizarre twist when on August 23, the North American Aviation office in Dayton verbally informed the Air Force that the company wished to withdraw its Project 1226 bid because the expenditure of engineering manhours on other projects would not allow the company to meet the desired deadline, McCollough notified Soule, the AF Headquarters, and the BuAer concerning this decision. On hearing of it, John Trenholm of the Weapons Systems Office at Wright-Patterson AFB, contacted Leo Devlin of Douglas to informally suggest to him that Douglas refigure its bid based on an InconelX structure, On August 30, North American sent a letter to the Air Force requesting that the company be allowed to withdraw from Project 1226 consideration, Shortly thereafter, Dryden informed Soule that he and Kelsey had decided, pending receipt of an official letter stating North American's desire to withdraw, to continue competition activities, Even when the North American letter had been received, the competition process continued, By September 7, Soule, now officially aware of North American's decision, contacted Dryden and recommended that the Research Airplane Committee give favorable consideration to the second-place bidder, Douglas, Dryden's response to this was negative, he feel-

8

Bell (0171)

Douglas (Model 684)

NAA (ES07487)

Republic (AP·76)

Areas Wing, total (sq. ft.) Vert. lail (sq. ft.) Ventral fin (sq. ft.) Horizontal (sq. ft.)

220 45.3 22.7 63

150.3 39.25 12.06 55.2

200 38.14 51.76

254 47.6 12.3 69.7

Engines Number Type Total thrust Obs.) Fuel Oxidizer

3 XlR-81 43,500 Jp·X RFNA

1 XlR·JO

1 XlR-30

~~goo

~~goo

lOX

LOX

Tank Capacities Fuel (g~~Fore ·Aft Oxidizer (gal)-Fore ·Aft

704 991 367

410 732 816

493 746 907.5

Weight Gross wi. Qbs.) Propel~nt Qbs.)

34,890 22,487

25,432 15,000

27,722 17,245

39,099 24,389

Performance Max. Speed (Ips) Max. Altitude (ft.)

6,850 400,000

6,655 375,000

6,950 800,000

6,619 300,000

Total Cost Three AlC and Static (millions)

$36.3

$36.4

556.1

547.0

Aut Right Estimate date

Jan. 1959

March 1958

Nov. 1957

Feb. 1958

11.42

4 XlR-81 57,600 Jp·X RFNA

710 900

630

ing that rather than award the contract to Douglas, the competition should be reopened, A presentation, prepared by Trenholm, of the design winner recommendations was made to the Technical Advisory Panel for Aeronautics on September 14, Because the Panel contained members from the aviation industry, and the information in the rejected proposals remained proprietary, the Panel was presented only with the North American proposal. The Piloted Aircraft Coordinating Committee attended the presentation on invitation, After the briefing the Panel endorsed the recommendations of the Research Aircraft Committee, thereby making North American Aviation the competition winner, The Defense Department's Piloted Aircraft Coordinating Committee took under consideration the funding of this project, looking for means of saving money since the North American Aviation cost estimates were the highest submitted, It was decided to ask Air Force headquarters for permission to make contact with North American Aviation for a more detailed cost estimate, Having learned that his company was the possible contract winner, R, H, Rice, then-Vice President and Chief Engineer for North American, wrote to the ARDC commander on September 23 and explained that the company had decided to withdraw from the competition because it had recently won Phase I awards for new fighter-bomber and long range interceptor competitions and also had increased activity relating to its on-going F-107 fighter. Having undertaken these more lucrative projects, it would not be possible for North American Aviation to accommodate the fast engineering man-hours build-up that would be required to support the desired Project 1226 time schedule, He went on to write that, "due to the apparent interest that has subsequently been expressed in the North American design, the Contractor wishes to extend two alternate courses which have been previously discussed with Air Force personnel: (1) The engineering man-power work load schedule has been reviewed and the contractor (North American) wishes to point out that Project 1226 could be handled if it were permissible to extend the schedule ... over an additional eight month period, (2) In the event the above time extension is not acceptable and in the best interest of the project, the Contractor is willing to release the proposal data to the Air Force at no cost," The decision to possibly extend the schedule had been discussed earlier on August 12 and had been approved, This allowed a way for North American to withdraw their previous letter of retraction once they had been told officially they had won the contract. Once the North American Aviation dilemma had been settled, the way was cleared for formal contract ratification and acceptance. Accordingly, on September 30, 1955, Col. Carl Damberg, Chief of the Aircraft Division, WPAFB, formally notified North American Aviation that its design had been selected as the winner. The other bidders were consequently notified on the same day of North American Aviation's selection and consequently thanked for their ready participation, On December 6, 1955, North American Aviation's Los Angeles, California division received a letter contract from the Air Force (AF 33[600]-31693), calling for the design,

construction, and development of three X-15 (Project 1226) aircraft, The quantity had been determined by experience. It had been noted during earlier research aircraft programs that two aircraft were enough to handle the anticipated workload, but three assured that the test pace could be maintained even with one aircraft down,

THE DEVELOPMENT AND FLIGHT TEST PROGRAMS: North American's X-15 design team, headed by Harrison "Stormy" Storms and Charles Feltz, had accepted an extraordinarily difficult task when they agreed to design, build, and initially flight test the new hypersonic research aircraft. Though giving the appearance of being a rather simple configuration, the X-15 was in truth, perhaps the most technologically complex single-seat aircraft of its day. Directly assisting Storms and Feltz was veteran test pilot A. Scott Crossfield, who had worked with (he NACA for some years prior to joining North American, and who had accumulated noteworthy flight test experience in a wide variety of aircraft types including the Bell X-1, the Convair XF-92A, and the Douglas D-558-11. Crossfield's X-15 input proved particularly noteworthy during the early days of the design and construction program as his experience permitted the generation of logical arguments that led to major improvements to the X-15's human factors design envelope, He played a key role, for instance, in convincing the Air Force that an encapsulated ejection system was both impractical and unnecessary, His arguments in favor of a highly developed ejection seat, capable of permitting safe emergency egress at speeds between 80 mph and Mach 4 and altitudes from sea level to 120,000 feet, saved significant money, significant weight, and significant development time. Though the first, and perhaps initially the most important pilot to contribute the X-15 program, Crossfield was not the only one to do so, In fact, all of the initially assigned X-15 pilots eventually participated in the program development phases, being called on to evaluate various operational systems and approaches, as well as such factors as cockpit layout, controllers, control systems, and guidance schemes. They worked jointly with research engineers in conducting the simulator programs designed to study the aspects of planned flight missions believed to present potential difficulties. Fixedbase ground simulator programs were initiated at North American's Los Angeles facility and at the various NACA (and later, NASA) research centers, The actual simulator had a working X-15 cockpit and control system that included actual hydraulic and control system hardware, It was operated via 3 Model 231 R analog computers, 90 diode function generators, computing sensors, electronic multipliers, and other associated computing equipment. Following initial tests at North American, where it was developed, it was moved to the NASA facility at Edwards AFB, Ground simulation of the dynamic environment was provided by use of the Navy centrifuge at Johnsville, Pennsylvania. A means for developing safe-landing techniques for X-15 flights was proVided by configuring a Lockheed F-104A to give a close presentation of X-15 dynamics, Similar work was conducted using a highly modified F-100F with a thrust reverser, a Convair TF-102A, and several other aircraft types, From the beginning, plans called for approaching the design speed and altitude through a carefUlly monitored incremental performance buildup. Because of the greater relative significance and cost of each flight in the program compared with tests of previous research aircraft, and in order to get as much out of each flight mission as possible, a six-degree-of-freedom X-15 fixed-base simulator was developed and used for flight training and planning, as well as concept and hardware development, Plans had been made at an early date for utilizing Edwards AFB, California, for all X-15 activities, Technical personnel from both the NACNNASA and the Air Force Flight Test Center were available, and had already cooperated on the planning and implementation of operational procedures. The performance and maneuvering capabilities of the X-15, communication and tracking requirements, and the need for suitable emergency landing areas suggested an essentially straight flight range with multiple tracking, telemetry, and communications capability, This led to construction by the Air Force of the X-15 High range, extending from Wendover, Utah, to Edwards AFB. Control stations were installed at Ely and Beatty, Nevada, as well

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the training program. This unit was later moved from North American to Edwards AFB where it eventually provided hundreds of simulated training flights for all X-15 pilots.

The X-15 program, once the contracts with North American were signed, was some three years away from actual flight test. Though most of the research into materials and structural science had been completed, there remained much work to be done in the field of human life sciences (including escape), the use of materials at extreme temperatures, and the establishment of high reliability in all flight sustaining systems. Of equal importance, was the development of fabrication and assembly techniques even though methods of doing so, at the time of contract signing, did not readily exist. North American met the challenge of each problem with a practical solution, and the eventual result was one of the most successful and productive flight test programs ever undertaken. Eventually some 2,000,000 engineering man-hours and 4,000 wind tunnel hours in 13 different wind tunnels were logged. On June 11, 1956, North American received a production go-ahead for the three X-15 airframes. The first metal was cut for the first aircraft the following September. The total construction process eventually consumed just over two years, and on October 15, 1958, the first aircraft, 56-6670, was rolled out of North American's Los Angeles, California plant doors for the first time. Following conclusion of the official roll-out ceremony, it was moved back inside and there prepared for shipment to Edwards AFB. Two days later, wrapped completely in a protective covering of heavy-duty wrapping paper, it was shipped by truck to Edwards AFB for initial ground test work. The static test program at Edwards consumed some five months. While this was on-going, North American also completed X-15 launch configuration modifications to B-52A, 52-003, and B-52B, 52-008, which were now redesignated NB-52A and NB-52B, respectively. B-52A, 52-003. had been moved to North American's Palmdale, California facility for modification on February 4, 1958, and B-52B, 52-008 had been moved to Palmdale on January 6, 1959. During November, 1958, the modification and preliminary flight test program for B-52A, 52-003, was completed and it was flown to Edwards AFB. B-52B, 52-008 followed several months later. Major modifications and/or equipment additions to the two B-52's consisted of the following: (1) An AN/APN-81 Doppler radar system to provide ground speed and drift angle information to the stable platform in the X-15. (2) A closed-circuit television system to monitor the prerelease and flight operations of the X-15. (3) An auxiliary UHF communications system to provide additional communications channels.

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as Edwards. Use of then-state-of-the-art instrumentation was specified to insure reliability in the recording of inflight measurements of aerodynamics, structural loads and heating, flight dynamics, and engine performance. The control stations also served to position the B-52 carrier aircraft over the desired launch point at the desired time by advising the B-52 pilot of course corrections and count-down-time corrections prior to launch; to advise the X-15 pilot of heading corrections, radar altitudes, and position during the flight, and provide energymanagement assistance; and to direct air search and rescue operations in an emergency.

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(4) A change in the AN/AIC-10 interphone system to provide an AUX UHF position. (5) Liquid oxygen top-off tanks (totaling at 1,500 gals.) in the bomb bay area for transferring to the X-15 LOX normally boiled off during flight. The B-52 forward body fuel tanks and mid-body fuel tanks were removed to accommodate this installation. (6) A rework of the upper fuel cell area to accommodate the installation of 27 storage cylinder tanks. These were used to supply system pressure for the liquid oxygen topoff tanks, space suit ventilation, cockpit windshield antifogging, the launch emergency release system, and the nitrogen cooling system for the X-15's engine. (7) The removal of the right main wing tank NO.3 to allow for carrier pylon tie fittings and supports in the front wing spar and rear wing spar. The fittings were used to fasten the carriage pylon to the B-52. The right and left inboard flaps were locked so that the flaps remained in a fixed retract position at all times. The inboard flap mechanism was disconnected, and the flaps were bolted to the flap tracks. (8) Bolting the carriage pylon to tie fittings between the right inboard engine nacelle and the B-52 fuselage. The pylon was a subsonic metal streamlined enclosure. The pylon was fastened to tie fittings at the front and rear wing spars between the right inboard engine nacelle and the fuselage. The upper structure of the pylon contained a rubber seal that contoured the pylon to the wing structure. The pylon structure, sway brace fillings, and hook mechanism absorbed the loads imposed by carrying the X-15. The release mechanism, in the pylon, consisted of a primary hydraulic and a secondary pneumatic release mechanism with a ground-operated, mechanically controlled release and locking mechanism used only for X-15 attaching purposes. The hydropneumatic, venting, antifogging, and liquid oxygen lines and electrical wiring from the right wing entered the rear section of the pylon and terminated at a quick-disconnect fitting at the front section. (9) Replacing the ECM compartment with the launch operator's station. The launch operator monitored the liquid oxygen top-off system control panel, radio and television equipment, and radar installation for the research program. Provisions were included for telemetering equipment. The escape hatch was changed to allow the operator to hold open the hatch during ground servicing operations. An astrodome-type viewing window was added to the right side for observation, and the portable oxygen bottle was relocated below the window. A defrosting system was provided for the window. Two steel straps across the window provided for safety of the observer if the window were to blowout. (10) Helium gas to pressurize the contents of the topoff tanks for transfer or jettison purposes. Nitrogen gas was used for pneumatic actuation of the liquid oxygen control valves. During the 2Q-second idle period, nitrogen gas was supplied to the engine for cooling the firing chamber. (11) Making breathing oxygen available to the crew members at all times. Also, oxygen was tapped from the B-52 oxygen system to supply the research pilot with breathing oxygen until flight release.

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(12) Making space suit ventilation available to the research pilot. A pressurized liquid nitrogen tank mounted in the carrier pylon supplied the system and provided ventilation gas at all times before launch. (13) Cockpit windshield anti-fogging, supplied by pressurized liquid nitrogen stored in the carrier pylon, that provided the system with gas flow at all times before launch. The system was controlled by the X-15 pilot according to his needs. (14) Electrical connections provided so that the B-52 power supply could be used for testing the X-15 electrical equipment and instruments. Connections also allowed the launch operator's panel to accomplish its intended function. (15) A mechanical linkage release system for the X-15 in the pylon, operated normally by hydraulic pressure and operated in case of an emergency by emergency pneumatic pressure augmented by an accumulator. (16) Propellant jettison lines, leading aft and overboard in the B-52, for dumping the liquid oxygen into the atmosphere in case of an emergency. (17) Loose equipment for the B-52 consisting of applicable Technical Manuals, the Modification Handbook, a spanner wrench, and covers for ground protection. For maintenance purposes, pressure liquid oxygen transfer line plates and a liquid oxygen transfer line pressu re test fixture were provided. (18) Installation of a high speed wheel, tire, and braking system on the B-52. (19) Removal of the B-52 bombing systems and tail turrets. Following completion of these modifications, both aircraft were tested by Air Force pilots Capt. C. C. Bock, Jr., and Capt. John Allavie. The launch panel operator during these test hops was William Berkowitz. Besides the two B-52's, a stable of chase aircraft was also scheduled for utilization. These were to be deployed along specific portions of each planned flight track. They were expected to prove most helpful in the launch and landing areas while performing various visual checks to ensure flight safety. By now, flight planning philosophy had been formulated, this calling for evaluation of vehicle characteristics and components in a noncritical environment before penetrating a possibly critical regime. Measurements of stability and control parameters were especially pertinent in establishing flight safety during the buildup phase. Assigned maneuvers during the flights yielded information on aircraft dynamics that could be analyzed in terms of stability and control-effectiveness derivatives. These were compared with predicted derivatives, and, when the comparison showed reason for concern, an additional flight would be planned before performance again could be advanced. Similarly, in the case of temperature measurements, analyses of thermal gradients and stresses were compared with information used by North American to establish the validity of the aircraft design. Following the five months of ground testing, on March 10,1959, X-15 #1, 56-6670. with North American company test pilot Scott Crossfield at the controls, completed its first captive flight under the wing of its B-52 carrier aircraft. Several additional captive flights followed, these

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culminating in the first glide flight on June 8. The first glide flight was successfuly completed, though not without incident. As it were, the X-15's control system proved significantly more sensitive in pitch during the landing approach than had been predicted, and as Crossfield began his flare, the aircraft began a pilotinduced pitch oscillation that proved difficult to control. Crossfield's piloting skills quickly got the upper hand, but not without significant effort. Development problems with the X-15's primary propulsion system, the Thiokol XLR99 rocket engine (which see), had led to a decision to complete initial powered flight trials with two Reaction Motors XLR11 rocket engines in its place. Accordingly, the first powered flight of the program, conducted on September 17, 1959, and utilizing the #2 aircraft, 56-6671, was flown with the stacked XLR11 's as the propulsion system. During the flight, with Crossfield again the pilot, the X-15 achieved a maximum speed of Mach 2.11 and a maximum altitude of 52,341 feet. The third flight, again with the #2 aircraft, also was successful, but the fourth flight was not. During the latter, which took place on November 5, an inflight explosion and fire in the engine compartment necessitated engine shutdown and an emergency landing on Rosamond Dry Lake (one of several designated emergency landing sites). Because propellants were jettisoned during a steep descent, jettisoning was incomplete and the landing weight was heavier than normal. When the nosewheel contacted the ground, the fuselage buckled just to the rear of the cockpit. The buckle was severe enough to cause the bottom of the fuselage to drag the ground. The post-accident investigation revealed that the principal problem existed in the nose-gear arresting system. In order to conserve space when the nose gear was retracted, the gear was stowed in a nearly compressed position. Upon rapid gear extension, the nitrogen gas which had been entrapped by the oil under high pressure was released and produced a gas-oil foam within the cylinder. Approximately the first one-third of the cylinder stroke was rendered ineffective by this foam; consequently, the loads built up to excessive values during the remainder of the stroke. A permanent solution was achieved by redesigning the internal mechanism of the strut to incorporate a floating piston which kept the gas and oil separated at all times. In order for repairs to be undertaken, the aircraft had to be returned to North American's Los Angeles facility. It remained there from November through early February. The damaged fuselage area was repaired and strengthened by dOUbling the number of fasteners and by adding a splice plate, both top and bottom, at the fuselage joint. Similar modifications were made later to the other two aircraft. By February, 1960, X-15 #2 was once again ready to resume flight testing. While X-15 #2 was undergoing repairs, the #1 aircraft, following modifications and the installation of an XLR11 propulsion system, became the primary flight test vehicle. On January 23, 1960, it made its first powered flight, again with Crossfield in the cockpit, setting a new speed record for the program along with a new altitude record of 1,669 mph and 66,844 feet respectively. The flight test program now began to accelerate. Joe Walker, the primary NASA X-15 test pilot, flew the X-15 on its first government sponsored mission on March 25, and Capt. Robert White became the first Air Force pilot to fly it when he successfUlly completed a mission on April 13. By now, a total of six government and military test pilots were assigned to the X-15 program. These included the NASA's Joe Walker, Neil Armstrong, and John McKay; the Air Force's Capt. Robert White and Maj. Robert Rushworth; and the Navy's Cmdr. Forrest Peterson. Later, the NASA team would be joined by Milton Thompson and William Dana; and the Air Force's team would be joined by Capt. Joe Engle, Capt. William Knight, and Maj_ Michael Adams. Each flight, by August, 1960, had expanded the X-15's speed and altitude envelopes to the point where new speed and altitude records were being set during virtually every mission. On August 4, Joe Walker had piloted the #1 aircraft to 2,196 mph, thus breaking the absolute speed record of 2,094 mph set some four years earlier by Capt. Mel Apt and the Bell X-2. This was followed by several altitude flights, including Robert White's August 12, 1960, mission in aircraft #1 to 136,500 feet-which broke Iven Kincheloe's four-year-old X-2 record of 126,200 feet. White followed with yet another speed record, reaching 2,275 mph on February 7,1961, also in X-15 #1. These speed and altitude records, with White

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at the controls, would remain the high water mark for the X-15 as powered by the XLR11 propulsion system. During January, 1961, NB-52A, 52-003,-was flown to Boeing's Wichita, Kansas facility for additional modifications. It was returned to Edwards on March 3D, bearing a new paint scheme and miscellaneous mission insignia. The X-15's flights, even at this early date in the flight test program, were already providing valuable stability, and control data. Coupled with preliminary aerodynamic heating data gathered during the first high speed (Mach 3) missions, it was apparent that the X-15 had extraordinary potential as a research tool. The remaining XLR11 flights served to familiarize new pilots with the X-15's unique flight characteristics. Additionally, important bio-medical base-line information was gathered that had a direct bearing on the special David Clark Company-developed MC-2 full-pressure suit (the standard MC-2 was a partial pressure suit-which did not provide pressurization for the feet) and its immediate successor, the NP22S-2. Though not often emphasized, the biomedical aspects of the X-15 program proved extraordinarily productive. Pilot biological signs were closely monitored via an electrocardiogram device, an oxygen flow sensor, a suit! cockpit pressure differential sensor, and a suit!helmet pressure differential sensor. This data was telemetered to the ground while the X-15 was in flight, and there recorded for reference. Additionally, significant research was conducted on the effects of weightlessness, as the X-15 was possibly the first aircraft to expose a pilot to this phenomenon for any significant length of time. These last XLR11 flights also served to allow installation and low-speed flight testing of the new Northropdeveloped "hot nose" (sometimes referred to as the "QBall" nose) which provided angle-of-attack, side-slip, and dynamic pressure information to the pilot via cockpit instrumentation. Some twenty-five XLR11 missions were eventually completed by the NASA, Air Force, and North American before the first flight-rated XLR99 arrived at Edwards for installation. By this time, the engine, unquestionably the most powerful, most complex, and safest man-rated throttleable rocket propulsion system in the world, had undergone an exhaustive static test program. Though long in gestation, it soon was to prove exceptionally reliable and extraordinarily safe. On March 28: 1960, the first flight-rated XLR99 was delivered by Thiokol to the Edwards AFB Flight Test Center and there prepared for installation in X-15 #3, 56-6672. Following successful mating of the engine to the airframe, the aircraft was moved to the static test area where an intensive ground test program was initiated. On June 8, however, a malfunctioning relief valve and pressurizing gas regulator caused a catastrophic explosion which totally destroyed the engine and severely damaged the aircraft. On June 17, while 56-6672 was being returned to North American for repairs, the second XLR99 arrived at Edwards AFB from Thiokol. This engine was shortly afterwards installed in X-15 #2, and following an intensive ground test program, prepared for the first XLR99 flight.

Cylindrical X-15 fuselage core section contained fuel and oxydizer tanks. Internal framing and supporting structure stiffened the unit's titanium skin and thus provided the strength required to accommodate the loads generated by the weight and mass of the propellants as the aircraft accelerated at hypersonic speeds.

Prior to the official roll-out ceremony for the first X-15 on October IS, 1958, the aircraft was extensively photographed by company photographers. Ground clearance of wheeled dolly prevented aNachment of the lower half of the ventral fin. A heat-resistant semi-gloss black paint was appfied over-all.

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A tractor tow was used to roll-out the X-15 during its unveiling on October 15, 1958. The aircraft's sinister all-black scheme was relieved somewhat by the predominantly all-white markings and the standard upper left wing insigne. Extended empennage fairing to accommodate twin-pack XLR11 is noteworthy.

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12

With Scott Crossfield at the controls, this took place without any major difficulties on November 15, 1960. Crossfield, whose contributions to the X-15 program had by now proven manifold and of great significance, was to fly the X-15 on only two more occasions. The first of these occurred on November 22, and the second on December 6, 1960. These missions were, in fact, the last of the obligatory contractor demonstration flights, and they served to effectively terminate North American's participation in the X-15 flight test program. Crossfield, shortly afterward was named officially to direct the testing of the Hound Dog cruise missile and North American's manned spaceflight systems. At this time, it proved opportune to return X-15 #1 to North American for installation of its XLR99. The aircraft departed Edwards on February 8, 1961. Four months later, on June 10, it was returned to Edwards to rejoin X-15 #2. With the completion of North American's participation, it was now possible for the Air Force, the Navy, and the NASA to embark on a systematic program that was designed to take the X-15 rapidly, but safely to the extremes of its performance envelope. North American, by prior contractual agreement, had been restricted to demonstrating the X-15's airworthiness at relatively low altitudes (100,000 feet) and reduced airspeeds (Mach 2). It now was time for the supporting agencies to undertake the more radical aspects of X-15' flight testing. The first NASAlAir ForcelNavy X-15 flight with the XLR99 installed took place on March 7, 1961. With Robert White at the controls, aircraft #2 accelerated out to 2,905 mph while ascending to an altitude of 75,000 feet. On March 30, with Joe Walker at the controls, X-15 #2 was used to set a new altitude record of 169,600 feet. Walker, during this flight, wore for the first time the new AlP22S-2 full pressure suit-which would later serve as the pre-production prototype for the suits worn by the Mercury, Gemini, and ApollO astronauts, and the basically similar S1 01 OB suits worn by Lockheed U-2 (second generation aircraft), A-12, F-12, and SR-71 pilots. By the fall of 1961, the X-15's had achieved speeds in excess of 3,500 mph and altitudes in excess of 215,000 feet. During October, the speed record was successfully improved with a flight to 3,920 mph, and in November, with one to 4,093 mph. These were the first flights out to the aircraft design Mach number. On October 29, 1961, the #3 X-15, almost completely rebuilt following the disastrous explosion that had occurred during the XLR99 ground tests on June 8, 1960, was returned to the NASA. During the rebuild, it had been fitted with a Minneapolis-Honeywell MH-96 self-adaptive flight control system. This device, originally tested aboard a Lockheed F-94C, a McDonnell F-1 01 A, a Cessna 310, and the X-15 simulator at North American, was designed to make the X-15 easier to control with either the aerodynamic or jet control system-or a combination of both. It was so capable, in fact, that it was possible for the MH-96 to control the X-15 without assistance from the pilot. 'While preliminary flight test activity was being undertaken with aircraft #3, 'on-going exploratory flights continued using aircraft #1. Design altitude was, in fact, achieved by Joe Walker in this aircraft during a flight to 246,700 feet on April 30. By now, aircraft #3 was progressing smoothly in its flight test program, having completed its first flight, with Neil Armstrong at the controls, on December 20. The MH-96 was proving to be of some benefit to pilots during high altitude missions, and underscoring this was a flight with Robert White at the controls, on June 21 (X-15 flights had been suspended during most of February and March because of rain and snow on the dry lakes-accordingly during this down period necessary modifications were undertaken on X-15 #1 and #2 .), to 246,700 feet (almost exactly the same altitude recorded by Joe Walker during his flight on April 30). On July 17 (the day before President Kennedy presented the prestigious Collier Trophy to Crossfield, Peterson, Walker, and White), the Walker and White record would fall, this time during a flight by White in X-15 #3 to 314,750 feet (this became the first flight in which astronauts wings were awarded to an X-15 pilot). On August 3, X-15 #1 was returned to North American for modification required by follow-on mission requirements. Changes included the installation of a nosemounted camera for optical degradation tests; installation of a new inertial guidance system (originally intended for the Boeing X-20); and installation of wing tip pods for transport of miscellaneous experiments. To this point, flight testing of the X-15 had been marred

by few serious accidents and even fewer injuries. On each occasion where damage had been sustained, repairs had proven possible. Even when aircraft #2 was very seriously damaged following a post-launch engine malfunction (which prevented fuel jettison) and an overgross weight emergency landing at Mud Lake, Nevada (seriously injuring test pilot John McKay-McKay jettisoned the canopy but could not escape and was distressed by ammonia leakage; a rescue helicopter blew the fumes away and McKay was cut free), it was set aside for a complete rebuild. This, in fact, eventually took place, giving birth to the high-performance X-15A-2. One accident would, unfortunately, mar forever the X-15's otherwise outstanding safety record. This occurred with aircraft #3 on November 15, 1967. The flight was Air Force pilot Michael Adams' seventh X-15 mission, and his third at the controls of the MH-96 equipped aircraft. Following an apparently normal launch, Adams had headed skyward and soon had achieved the mission's speed and altitude objectives of Mach 5.2 and 266,000 feet respectively. As with most of these later X-15 flights, a number of infligHt tests had been scheduled, including on this mission, wing rocking in order to facilitate optical tracking of the aircraft's exhaust plume. It was noted by ground personnel monitoring recording instrumentation, however, that when Adams initiated these control inputs, the X-15 exceeded by far the test's required bank angles. Less than thirty seconds later, the X-15 had skewed some ninety-degrees to its ballistic flight path and had initiated entry into a spin-stall while traveling at a speed in excess of Mach 5. The spin continued from 230,000 feet down to 125,000 feet, at which point a conventional dive was entered. With the onset of the dive, high frequency pitch oscillations were encountered, these leading to command saturation of the MH-96 system computer and a resulting increase in the severity of the pitch control problem. Structural loads by this time had increased to plus and minus 15g's-lar beyond the X-15's plus 7.33g and minus 3g design limits. Seconds later, the aircraft disintegrated. Adams did not survive. It was many months belore the official accident investiga1ion was completed. The final report indicated that the cause was a combination of human error and mechanical system failures, not the least of which were the pilot's unexplained laXity in permitting the aircraft to deviate from its flight path (thought possibly to have been caused by either instrument misinterpretation, cockpit distraction, or vertigo), and the MH-96 system's imposed limit cycle that sustained a longitudinal pitching oscillation allowing the control system to impose excessive flight loads on the airframe.

Following its arrival at Edwards, the first X-15 was mated to its NB-52A carrier aircraft and checked for launch pylon and associ~ted systems compatibility. Additionally, a structural test rig was built around the pylon and alfcraft and used to determine structural integrity of the mated configuration.

•

X·15A·2 During the latter part of 1962, X-15 flight tests were conducted to verify a predicted improvement in lateral directional handling qualities at high Mach numbers and high angles of attack with the lower ventral fin removed. The thirty-first flight of X-15 #2 on November 9, 1962, was planned to further investigate the aircraft's ventral-off handling qualities. Following launch, it was discovered by NASA pilot John McKay that only 30% thrust could be obtained from the XLR99 engine as a result of a throttle control failure. An emergency landing was then attempted at Mud Lake in accordance with preplan ned alternate procedures. Unfortunately, at touchdown, the left main gear collapsed, causing the X-15 to skid sideways.and turn over on its back (the gear struciural limit was exceeded primarily because the aircraft landing flaps had failed to extend; because 01 this, the landing speed proved greater than that which the gear was designed to Withstand). The X-15 suffered extensive damage to the landing gear, fuselage, wings, and tail surfaces; McKay, found to be suffering frOm three crushed vertabrae and lung damage caused by inhaling lE!aking ammonia gas, was seriously injured. McKay was able to return to test piloting less than six months after the accident, but the injuries sustained eventually contributed to his retirement from NASA. X-15 #2 had accumulated a total free flight time of 4 hours 40 minutes and 32.2 seconds at the time of the accident. Following the decision to rebuild it the aircraft was transported by truck back to North American's Los Angeles facility where the reconstruction, and as it would turn out, major modification program, was to take place. The X-15A-2 project, as it eventually became known, was a conscious attempt to increase, to its absolute maximum potential, the performance capabilities of the stan-

-

~------'

~ ~A dedicated ground test facility was built at Edwards to accommodate program propulsion system check-out

....

requirements. The facility permitted two aircraft to undergo testing at anyone time. Both the XLRll and XLR99 were ground tested using this special pad and its associated monitoring equipment and bunkers.

.-----

.-

~

'~~-

Prior to its first glide flight, which occurred on June 8, 1959, the first X-15 is seen being uploaded under the launch pylon attached to the right wing of the NB-52A. Two hydraulic jacks were ramp-mounted to accommodate the X-15's uploading requirements. To left, on NB-52A fuselage side, is unfaired view camera.

13

~ ~

~ ......., Following installation of the first flight-rated XLR99 powerplant, the third X-15 was almost completely destroyed when a malfunctioning relief valve and pressurizing gas regulator caused a catastrophic explosion on June 17, 1960.

dard X-15 aircraft. NASA, the Air Force, the Navy, and North American had, following successful completion of the X-15's performance envelope expansion program, concluded that the airframe was not being utilized to its fullest potential. The basic materials of which the X-15 was built would permit flight test work at speeds significantly greater than those being achieved by the aircraft in its stock configuration. Additionally, at the time of the decision to rebuild X-15 #2 into a more advanced aircraft, the U.S. aerospace industry was beginning to express strong interest in the potential of the supersonic combustion ramjet, known more popularly as the scramjet. This propulsion system, which was highly dependent upon supersonic and hypersonic velocities in order to achieve combustion, could be tested most effectively by the X-15. The McKay accident that seriously damaged X-15 #2 effectively set the stage for the birth of the advanced X-15 configuration that the various participating agencies had been contemplating. The aircraft was obviously repairable, the two flightworthy X-15's, #'s 1 and 3, were apparently capable of completing the remaining portions of the initial X-15 flight test program, and relatively little additional time would be involved in the modification over that required by North American simply to rebuild the aircraft. Approval for the modification was given to North American under contract AF33(657)-11614, on May 13, 1963. Contract funding was $5-million. North American finished the project some nine months later, and on February 19, 1964, the X-15A-2 was returned to Edwards AFB for initiation of four months of static ground testing. Following turnover to the Air Force/NASA on February. 24, and a first captive flight on June 15, the first postmodification flight, with Robert Rushworth at the controls, took place on June 25. A maximum speed of 3,104 mph at 83,300 feet was achieved. Additional envelope expansion flights took place throughout the rest of the year, these all being flown without the X-15A-2's distinctive external fuel tanks. In fact, flights without the external tanks continued into 1965, with the first tank-equipped flight not taking place until November 3.

The second X-15, 56-6671, with Scott Crossfield at the controls, was seriously damaged following an emergency landing at Rosamond Dry Lake on November 5, 1959. The aircraft was returned to North American for repairs.

The first flight with the external tanks attached and fully loaded with fuel occurred on July 1, 1966. Unfortunately, a fuel flow anomally led to premature engine shutdowninstrumention had indicated that the propellant was not transferring from one of the external tanks. This would have led to a lateral asymetric condition that could have exceeded the roll control power available. The ground rule in such situations was for the pilot to jettison the tanks and land at the launch lake. After the fact, the NASA and Air Force investigators found that the flow indicating system was, in fact, working as designed. Following a number of additional exploratory flights, on November 18,1966, the X-15A-2 was used to establish a world absolute speed record for conventional winged aircraft of Mach 6.33 (4,250 mph). This mission had been successfully accomplished due in part to the use of an experimental ablative coating developed by the Martin Marietta Corporation under the name of MA-25S. This material, consisting of a resin base, a catalyst, and glass bead powder, was designed to be sprayed on the external surface of .the X-15 (or any hypersonic vehicle) in order to protect it from the extreme heat generated by high speed flight in the atmosphere. MA-25S was expendable and following a mission, it could be removed and discarded and replaced with fresh ablator. Other ablators had been explored for use with the X-15A-2, these including Emerson Electric Thermo-Lag T-500, Dow Corning DC 325, Armstrong Cork #2755, NASA Purple Blend, Molded Refrasil Phenolics, and General Electric Century Series Materials. On several occasions these ablators were applied to various aerodynamic surfaces of the X-15 and tested for their heat protection qualities. Though Thermo-Lag T500 proved the most effective, it later was superceded by Martin Marietta's product. The success of the November 18 mission gave the various X-15 supporting agencies confidence in the aircraft's ability to perform safely at very high Mach numbers. A decision was therefore made calling for the use of the X-15A-2 in another very high speed flight out to the maximum speed potential of the aircraft in its modified configuration. On October 3,1967, again with Knight at the controls,

the X-15A-2 was launched on what was to become its most memorable, and historically most significant flight. This mission, wherein the aircraftwas to carry a dummy scramjet out to the X-15A-2's maximum Mach number, also was to serve as another test for the MA-25S ablative coating. Aircraft preparation for this flight, in fact, consisted mainly of refurbishing the ablative material wear resulting from the previous high Mach flight. This task required approximately 700 man-hours over a two week period. In the interest of flight safety it had been required, on all X-15 missions, that the major aircraft systems function normally for the flight to proceed to the planned speed. Malfunctions possible during the flight dictated that a preplanned alternate flight be flown. On the highMach flight that was planned for October 3, 1967, the maximum speed was limited to 5,400 feet per second in case of failure of any of the dampers, attitude indication, the ball nose andlor external propellant flow at certain times. The ground rule for a failure of the engine to light on the first attempt after launch on previous full external tank flights had been to immediately eject the external tanks and fly an alternate profile. The main concerns were that the maximum allowable dynamic pressure of 1,000 Ibs. per square foot not be exceeded during the rotation (due to altitude loss) and that the dynamic pressure at the planned tank ejection not be too high for good separation characteristics. The possibility that a good mission could be lost to these problems was very real because delayed engine lights, following the drop from the mother ship, had occurred on six X-15 flights. However, for this speed record flight, the ground rule was changed to allow one restart attempt. By the flight date, all the various alternate procedures were well known to the pilot after having practiced on the simulator for 35 hours. Fortunately, it did not become necessary to utilize an alternate profile since the flight proceeded basically according to plan. The takeoff under the wing of the B-52 went smoothly and within an hour, the launch altitude had been reached and the final countdown was underway. The launch occurred without problems, and following engine ignition, ~ ~

~

. . . --=--~-- :;.

Unquestionably the most serious accident of the X-15 program up to the date of its occurrence on November 9, 1962, was that suffered by NASA test pilot John McKay when a powerplant malfunction prevented fuel jettison. In an over-gross weight condition, McKay attempted an emergency landing on Mud Lake, Nevada, only to have the nose gear fail. Back injuries caused by the hard landing and lung injuries caused by leaking propellants, later caused McKay to retire from NASA.

14

Knight adjusted the angle-of-attack to 15° and began the mission profile. Within 38 seconds after launch, the planned pitch angle of 35° had been achieved and was maintained to within plus or minus 1° until pitch-over altitude was reached. The external tanks were ejected 67.4 seconds after launch at a speed of approximately Mach 2 and an altitude of approximately 70,000 feet. The external tank recovery system performed satisfactorily and the tanks were later recovered in repairable condition. As the X-15A-2 came level at an indicated altitude of 99,000 feet, Knight decreased the angle-of-attack to 6° to maintain a zero rate of climb. During this task, Knight reported that the pitch control was very sensitive and that it was difficult to hold a constant angle-of-attack. Knight now reached the maximum speed portion of the flight. Engine shutdown was reported to have occurred at 6,500 feet per second (however, the final radar data analysis revealed the maximum velocity to be 6,630 feet per second). The total engine burn time had been 141.4 seconds, which compared favorably with the 141 seconds planned (however, the X-15A-2 had achieved a velocity which was 130 feet per second faster than that of the simulator during this part of the flight). During the deceleration, Knight concentrated on performing stability and control maneuvers and as a result, the final flight profile was not exactly as planned. The ability of the MA-25S ablative material to protect the X-15A-2 structure from the high aerodynamic heating was considered good except in the area of the dummy ramjet where the heating rates were significantly higher than predicted. Considerable heat damage occurred on the dummy ramjet and the ramjet pylon. The ramjet instrumentation ceased to function approximately 25 seconds after the X-15A-2 engine shutdown-indicating that a burn-through of the ramjeUpylon structure had occurred and that the engine had possibly separated from the aircraft (in fact, it had, having fallen onto the Edwards AFB bombing range). Shortly thereafter the heat propagated upward into the lower aft fuselage area causing the XLR99 hydrogen peroxide hot light to illuminate in the cockpit. Ground control, assuming a genuine overheat condition, requested Knight to jettison the remaining engine peroxide. Unfortunately, the high heat in the aft fuselage area also had caused a failure of a helium control gas line allowing not only the normal helium source gas to escape, but also that of the emergency jettison control gas supply as well (because of the failure of a check valve). Thus, the remaining residual propellants could not be jettisoned. The aircraft was an estimated 1,500 Ibs. heavier than normal at landing. Fortunately, the landing was accomplished without incident. Post flight examination of the X-15A-2 revealed a number of serious problems with the ablative coating. The high temperature and extremely high speed airflows had combined to char and pit the MA-25S coating so badly that it was impossible to restore it for use during another flight. A completely new coating would be required for any future mission. Additionally, the X-15A-2's airframe had suffered serious damage, particularly in the ventral fin area and in various segments of the fuselage and associated chine/tunnels. In light of these problems, and in consideration of the fact that the X-15A-2 was obviously nearing the end of its service life, a decision was made to ground it permanently, and to turn it over to the Air Force Museum for permanent static display. Shortly thereafter, the X-15A-2 was stripped of its remaining ablator and restored to its standard bare metal condition. It later was returned to the Air Force which then placed it in the permanent possession of the United States Air Force Museum at Wright-Patterson AFB, Ohio. It now is on display in the Museum main display hall. With the demise of X-15 #3 and the modified (X-15A-2) aircraft making its last flight on October 3, 1967, only X-15 #1 remained airworthy by early 1968. This aircraft, too, was quickly reaching the end of its long flying career and though it continued to successfully explore unknown ground for the NASA and the Air Force, it finally was removed from active flight duty in late December (following the last airborne X-15 operation, an unsuccessful attempt to complete the 200lh X-15 flight, on December 12, 1968; the flight was aborted when weather conditions deteriorated prior to launch-the mated X-15 returned to Edwards in the middle of a snowstorm and though further attempts were made to consummate the flight on the 13th, 16th, 17th, 18th, and 20th, it eventually was cancelled and the X-15 was demated for the last time). Like the X-15A-2, X-15 #1 also was preserved for posterity, being shipped to the Smithsonian Institution's

The first X-15, 56-6670, is permanently displayed in the main hall of the Smithsonian Institution's National Air & Space Museum in Washington, D.C. It is cable suspended from structural girders and is in the same half with the original Wright 1903 "Flyer", the Belf X-I (46-062), the "Spirit of SI. Louis", and "Apollo XI".

The X-15A-2, 56-6671, is permanently displayed in the United States Air Force Museum at Wright-Patterson AFB, near Dayton, Ohio. Since this photo was taken, accurate markings, such as the AF serial number, have been painted on the aircraft and the display XLR99 powerplant has been placed in a glass case.

15

~

~ ~ ~

~ ~

_-........::.. 0

;~-

The full-scale X-15A-2 metal mock-up was donated to the Alabama Space and Rocket Center near Huntsville, Alabama. It is presently displayed outdoors mounted on a steel pole. When the Center first opened in 1970, the actual X-15A-2 was temporarily displayed there before going to the AF Museum.

~

Another X-15 full-scale mock-up was donated to the Tucson Air Museum Foundation of Pima County and their associated museum near Tucson, Arizona. It is currently displayed there under the wing pylon of N8-52A, 52-003, which also is on permanent display. Discernible on the 8-52 is spraylat preservative material.

16

National Air & Space Museum during December, 1958, and there being officially accepted on December 31. It was placed on static display for the first time during 1969 in the old Smithsonian facilities and later was moved to the new National Air & Space Museum when it opened its doors for the first time during 1976. It remains a primary NASM exhibit and presently is hanging next to the 1903 Wright Flyer and other historic aircraft in the main hall of the NASM in Washington, D.C. The X-15's legacy remains visible not only in the form of the two preserved airframes now on display at the Smithsonian Institution and the Air Force Museum, but also in all the many contributions it made to the aerospace community. John Becker, who was intimate with the X-15 program almost from the day of its birth, probably summarized the program best by assembling a modestly detailed listing of all the X-15's most important accomplishments: 'Development of the first large restartable "manrated" throttleable rocket engine. 'First application of hypersonic theory and windtunnel work to an actual flight vehicle. 'Development of the wedge tail as.a solution to hypersonic directional stability problems. 'First use of reaction controls for attitude control in space. 'First reusable superalloy structure capable of withstanding the temperatures and thermal gradients of hypersonic reentry. 'Development of new techniques for the machining, forming, welding, and heat treating of Inconel-X and titanium. 'Development of improved high-temperature seals and lubricants. 'Development of the NACA-sponsored "Q-Ball" hot nose flow direction sensor for operation over an extreme range of dynamic pressures and temperatures. 'Development of the first practical full-pressure suit for pilot protection in space. 'Development of nitrogen cabin conditioning. 'Development of inertia flight data systems capable of functioning in a high-dynamic pressure environment. 'Discovery that hypersonic boundary layer flow is turbulent and not laminar. 'Discovery that turbulent heating rates are significantly lower than had been predicted by theory. 'First direct measurement of hypersonic skin friction, and discovery that skin friction is lower than had been predicted. 'Discovery of "hot spots" generated by surface irregularities. 'Discovery of methods to correlate base drag measurements with tunnel test results so as to correct wind tunel data. 'Development of practical boost-guidance pilot displays. 'Demonstration of a pilot's ability to control a rocket-boosted aerospace vehicle during exoatmospheric flight. 'Development of large supersonic drop tanks. 'Successful transition from aerodynamic controls to reaction controls, and back again. 'Demonstration of a pilot's ability to function in a weightless environment. 'First demonstration of piloted, lifting atmospheric reentry. 'First application of energy-management techniques. 'Studies of hypersonic acoustic measurements used to define insulation and structural design requirements for the Mercury spacecraft. 'Use of the three X-15's as testbeds to carry a wide variety of experimental packages. The following is a complete listing of all X-15 pilots, their affiliation, and the number of X-15 flights each logged: Michael J. Adams, Maj. 7 Air Force Neil A. Armstrong NASA 7 North American A. Scott Crossfield 14 William H. Dana NASA 16 Air Force Joseph H. Engle, Capt. 16 William J. Knight, Capt. 16 Air Force John B. McKay NASA 29 Forrest S. Peterson, Cmdr. Navy 5 Air Force Robert A. Rushworth, Maj. 34 Milton O. Thompson NASA 14 Joseph A. Walker NASA 25 Robert M. White, Capt. Air Force 16

FLIGHT LOG: What follows is a complete listing of all X-15/X-15A-2 flights conducted between June 8,1959, and October 24,1968. Each flight is listed chronologically. Unnumbered flights (marked with an .) are aborts wherein the mated X-15 was carried aloft but not launched by the B-52; 'astronaut wings flight' indicates that the mission was above 50 miles, therefore entitling the pilot to astronaut wings: FLIGHT #

DATE

AlC#

3/10/59

2

3

4

5

PILOT

MACH/MPH

ALT. (FT.JMS)

Crossfield

4/1/59 4/10/59 5/21/59 6/8/59 7/24/59

1 1 1 1 2

Crossfield Crossfield Crossfield Crossfield Crossfield

9/4/59 9/17/59

2

Crossfield

10/10/59

2

Crossfield

10/14/59 10/17/59 10/22/59 10/31159 11/5/59

2 2 2 2 2

Crossfield Crossfield Crossfield Crossfield Crossfield

12/16/59

Crossfield

1/23/60

Crossfield

.79/522

37,550

Crossfield 2.11/1,393

52,341

2.15/1,419

61,781

1.00/660

45,462

2.53/1,669

66,844

2/4/60

2

Crossfield

6 7

2/11/60 2/17/60

2 2

Crossfield Crossfield

2.22/1,466 1.57/1,036

88,116 52,640

8

3/17/60 3/18/60 3/25/60 3/29/60 3/31/60 4/13/60

2 2 1 2 2 1

Crossfield Crossfield Walker Crossfield Crossfield White

2.15/1,419

52,640

2.00/1,320 1.96/1,293 2.03/1,340 1.94/1,254

48,630 49,982 51,356 48,000

13

4/19/60 5/5/60

1 2

Walker Crossfield

2.56/1,689

59,496

14

5/6/60

White

2.20/1,452

60,938

15 16 17

5/12/60 5/19/60 5/26/60

Walker White Crossfield

3.19/2,111 2.31/1,590 2.20/1,452

77,882 108,997 51,282

3.31/2,196

78,112

3.13/1,986

75,982

3.23/2,182

79,864

1.68/1,108

53,043

1.94/1,280 2.02/1,333

53,800 50,700

1.95/1,287 2.97/1,960 1.90/1,254 2.51/1,656

48,900 81,200 54,750 61,900

1.75/1,155 2.85/1,881 1.80/1,188

48,840 53,374 50,095

1.88/1,211 3.50/2,275

49,780 78,150

9 10 11 12

18

20 21 22

23 24 25 26 27 28 29 30 31

32 33

1 1 2

5/27/60 6/3/60

Walker Walker

6/8/60 8/4/60

Walker Walker

8/11/60 8/18/60 8/19/60 9/2/60 9/10/60 9/20/60 9/23/60 10/11/60 10/13/60 10/20/60 10/28/60 "\1/4/60 11/4/60 11/15/60 11/17/60 11/22/60

1 1 1 1 1 1 1 1 2 1 1 2 1 2 1 2

Walker Walker Walker White White Peterson Peterson Peterson Crossfield Peterson McKay Crossfield Rushworth Crossfield Rushworth Crossfield

11130/60 12/6/60 12/9/60 12/15/60 1/11/61 2/1/61 2/7/61

1 2 1 1 1 1 1

Armstrong Crossfield Armstrong White McKay McKay White

REMARKS scheduled captive flight; SAS malfunction; B-52 power supply problems; regulator frosted; generator failure H&V and radio failure; no. 2 APU cut off no. 1 and 2 APU failure; vertical stabilator upper panei crack APU failure; low liquid nitrogen source pressure planned glide flight scheduled full fuel captive flight; low liquid oxygen pressure; radio failure liquid oxygen tank pressure fluctuation due to regulator leakage first powered flight (XLR11 engines); turbopump case failure; fire in hydrogen peroxide compartment, engine compartment, and lower vertical fin liquid oxygen relief valve or regulator malfunction; could not top-off; intercom out cabin pressure failure; WALC jettison valve failure; could not top-off nose gear door faiied on landing oxygen selector handle stuck; canopy frosted weather cancel engine failure; fuselage structural failure upon landing at Rosamond dry lake; aircraft returned to Los Angeles for rebuild liquid oxygen tank pressure slow to build-up with excessive pressure drop during prime first flight following installation of XLR11 engines; telemetry failed during taxi WALC vent-relief valve failed; no. 2 APU failed; H&V failed; auto top-off failed nose gear bottomed on landing automatic shutdown of one chamber in upper XLR11; LH skid bungee and automatic top-off failed; high acceleration flight to test structure maneuver test (6g) WALC leak; windshield delaminated Walker's first flight; first NASA flight; roll damper malfunctioned SAS test; cold soak to simulate Wendover launch 3.5-5g pull-outs to simulate reentry White's first flight; no data till "High Key"; no landing film; hydraulic hose leak no gear data taken; hydraulic hose leak no 1 APU overspeed and shutdown; APU hydrogen peroxide failed to jettison normal ventral jettison system failed; roll damper malfunction; stability and control flight; SAS test; telemetered physiological data received and recorded on board an aircraft for the first time in USA Silver Lake launch; stable platform inoperative first flight over 100,000'; Silver Lake launch first ballistic control system flight; SAS 4-4-8 gain settings checked ok; control system vibration after landing; first test of Reaction Control System telemetry and power supply failed no. 2 hydraulic system 1,800 psi maximum-hydraulic stand coupling defective gaseous nitrogen pressure too low (1,700 psi) unofficial world speed record; Silver Lake launch; burned canopy seal gaseous nitrogen regulator runaway no start no. 1 APU Silver Lake launch telemetry system failed Silver Lake launch no start no. 2 APU Peterson's first flight; premature shut down of both XLR11's no engine hydrogen peroxide source pressure; regulator runaway first XLR99 flight attempt; no. 2 APU hydrogen peroxide valve leak McKay's first flight no. 2 APU hydrogen peroxide leak Rushworth's first flight first flight with XLR99; 50% power; hydraulic leak lower XLR11 shutdown and restarted; no. 2 APU start sluggish first inflight restart of XLR99; no. 2 ballistic control system nosedown rocket leak Armstrong's first flight; only seven of eight XLR11 chambers ignited Crossfield's last X-15 flight; two XLR99 air restart attempts first "hot nose" flight low pressure in no. 2 hydraulic system (1,500 psi) pressure decay in no. 2 hydraulic system when controls moved last XLR11 powered flight; Silver Lake launch; stability and control test

17

FLIGHT #

DATE

34

2/21/61 2/24/61 3/7/61

A/C # 2 2 2

PILOT White White White

3/21/61

2

Walker

35

3/30/61

2

36

4/21/61

MACH/MPH

ALT. (FT ./MS)

4.43/2,905

77,450

Walker

3.95/2,760

169,600

2

White

4.62/3,074

105,000

37

5/19/61 5/25/61 6/20/61

2 2 2

Walker Walker White

4.95/3,307

107,500

38 39

6/23/61 8/10/61

2 1

White Peterson

5.27/3,603 4.11/2,735

107,700 78,200

40 41

9/12/61 9/28/61

2 2

Walker Peterson

5.21/3,618 5.30/3,600

114,300 101,800

42

9/29/61 10/4/61

1 1

Rushworth Rushworth

4.30/2,830

78,000

43

10/11/61

2

White

5.21/3,647

217,000

44

10/17/61 10/27/61 11/2/61 11/3/61 11/9/61

1 1 1 1 2

Waiker White White White White

5.74/3,900

108,600

6.04/4,093

101,600

12/19/61 12/20/61 1/10/62

3 3 1

Armstrong Armstrong Peterson

3.76/2,502 .97/645

81,000 44,750

3/29/62

3

Armstrong

3/30/62

3

Armstrong

3/31/62 4/5/62 4/18/62 4/19/62

3 3 1 1

Armstrong Armstrong Walker Walker

4.12/2,850

180,000

5.69/3,866

154,000

3 2 2 1 1

Armstrong White White Walker Walker

5.31/3,789

207,500

52

4/20/62 4/25/62 4/26/62 4/27/62 4/30/62

4.94/3,489

246,700

53

5/8/62

2

Rushworth

5.34/3,524

70,400

54

5/22/62

1

Rushworth

5.03/3,450

100,400

55

5/25/62 5/29/62 6/1/62

2 2 2

White White White

5.42/3,675

132,600

56 57 58

6/7/62 6/12/62 6/21/62

1 3 3

Walker White White

5.39/3,672 5.02/3,517 5.08/3,641

103,600 184,600 246,700

59

6/27/62

1

Walker

5.92/4,104

123,700

60

6/29/62

2

McKay

4.95/3,280

83,200

7/10/62 7/11/62

3 3

White White

61 62

7/16/62 7/16/62 7/17/62

3 1 3

White Walker White

5.37/3,674 5.45/3,832

107,200 314,750

63 64

7/19/62 7/26/62

2 1

McKay Armstrong

5.18/3,474 5.74/3,989

82,250 98,900

65 66

8/1/62 8/2/62 8/8/62

3 3 2

Walker Walker Rushworth

5.07/3,438 4.40/2,943

144,500 90,877

45

46 47

49 50 51

18

REMARKS no cabin pressure; 20° in bank (inertial system malfunction) inertial attitude was off first NASA XLR99 mission; Silver Lake launch; temperature sensitive paint on nose no 8-52 AC power; heavyweight 8-52 ianding attempted; drag chute fitting failed; 8-52 locked wheels with resultant right front tire blow-out Hidden Hills launch; engine relight; first flight with new AlP-22S-2 suit Hidden Hills launch; engine relight; temperature sensitive paint on nose lack of APU source pressure for launch go-around Mud Lake launch; reentry investigation; side stick control used chase pilot thought liquid oxygen spillage was hydrogen peroxide coming from APU compartment drain line wing showed heat effect; iost cabin pressure; reentry investigation lost cabin pressure; Silver Lake launch; first flight of X-15 #1 with XLR99 Mud Lake launch; smoke in cockpit due to scorching paint Hidden Hills launch; 1,000° F. skin temperature; test reentry heat limit; smoke in cockpit due to scorching paint SAS roll mode pulsating first flight made with lower ventral fin removed; SilverLake launch; stability and control in level flight, portions of autostab and autopilot switched off several times in test outer panel of left windshield cracked; reentry and Reaction Control System test Mud Lake launch; SAS disconnected to study control problems overcast at Mud Lake cabin pressure failure no igniter idle pressure first flight in which X-15 design speed achieved; Mud Lake launch; aircraft showed heat effects; outer right windshield shattered inertia roll indication error X-15 #3 first flight; Silver Lake launch; Honeywell system checkout emergency landing at Mud Lake, NV after engine failed to light during two attempts; main chamber pressure switch malfunction; mission was to explore stability and control problems at high angle-of-attack stable platform heat exchanger iced; fire extinguisher detector faulty; light went on and off engine gaseous nitrogen cooling gas circuit breaker not closed on launch panel Honeywell system failed check no. 24 engine relight ok cloud cover over Mud Lake Mud Lake launch; SAS - primary control system twice shutdown to test emergency control system; reentry maneuver Mud Lake launch clouds over launch point valve malfunction and no igniter idle cloud cover over launch point first flight to design altitude; to obtain data on RCS at high altitude; data on aerodynamic heating during reentry; 900° F; rearward facing cameras show ice breaking away no 4 engine stopped on 8-52; 1,323° F. at wing lower 5% chord; max. q was 2,027 psf Hidden Hills launch; roller coaster flight with 3 peaks for local airflow investigation stable platform overheated 8-52 stable platform would not erect Delamar launch; steepest descent so far; highest angle-ot-attack (27°) so far in reentry; primary control system shutdown to test emergency control system Hidden Hills launch; structural test maneuver Delamar launch Delamar launch; "250,000 or no ice cream" inscribed on frosty fuselage by ground crewman unofficial world speed record; Mud Lake launch; test to determine safe reentry angle-of-attack (23° highest so far) Hidden Hills launch; mission to determine hot spots; white streaks shown on temperature sensing green paint 8-52 left aft landing gear would not retract no. 1 APU hydrogen peroxide jettisoned when tank pressurized; regulator overshot, rupturing blowout plug umbilical disconnect caused by short lanyard

Mud Lake launch FAI certified world altitude record for class; first astronaut wings flight; steepest climb 41°; reentry angle-of-attack 23~; severe shaking on reentry due to excessive yaw Hidden Hills launch; high speed heat tests Mud Lake launch; smoke in cockpit; lost lower vertical door; no. 1 hydraulic system leaked; no automatic top-off; auto control (yaw damper) disconnected; roller coaster descent made to simulate emergency reentry no fuel tank pressure Mud Lake launch; Honeywell system checkout Hidden Hills launch; high speed heat tests

FLIGHT #

DATE

A/C #

PILOT

MACH/MPH

ALT. (FT.lMS)

98

12/18/63 1/8/64

1 1

Rushworth Engle

5.32/3,616

139,900

99

1/16/64

3

Thompson

4.92/3,242

71,000

100

1/28/64

1

Rushworth

5.34/3,618

107,400

101 102

2/19/64 3/13/64

3 3

Thompson McKay

5.29/3,519 5.11/3,392

78,600 76,000

103

3/17/64 3/27/64

1 1

Rushworth Rushworth

5.63/3,827

101,500

104 105

3/31/64 4/8/64 4/29/64

3 1 1

McKay Engle Rushworth

5.01/3,468 5.72/3,906

175,000 101,600

106

5/11/64 5/12/64

3 3

McKay McKay

4.66/3,084

72,800

107

5/19/64

1

Engle

5.02/3,494

195,800

108

5/21/64

3

Thompson

2.90/1,865

64,200

6/11/64 6/15/64 6/23/64

1 2 2

Thompson Rushworth Rushworth

109

6/25/64

2

Rushworth

4.59/3,104

83,300

110

1 3 3

McKay Engle Engle

4.96/3,334

99,600

111

6/30/64 7/2/64 7/8/64

5.05/3,520

170,400

112

7/28/64 7/29/64

3 3

Engle Engle

5.38/3,623

78,000

113

8/12/64

3

Thompson

5.24/3,535

81,200

114

8/14/64

2

Rushworth

5.23/3,590

103,300

115

8/26/64

3

McKay

5.65/3,863

91,000

116

9/3/64

3

Thompson

5.35/3,615

78,600

117

9/23/64 9/28/64

3 3

Engle Engle

5.59/3,888

97,000

118

9/29/64

2

Rushworth

5.20/3,542

97,800

10/2/64 10/15/64 10/29/64 10/30/64

1 1 3 3

McKay McKay Thompson Thompson

4.56/3,048

84,900

4.66/3,113

84,600

2 2 2 1 3

McKay McKay McKay Engle Thompson

4.66/3,089

87,200

122

11/6/64 11/16/64 11/30/64 12/4/64 12/9/64

5.42/3,723

92,400

123

12/10/64

1

Engle

5.35/3,675

113,200

124

12/22/64

3

Rushworth

5.55/3,593

81,200

125

1/13/65

3

Thompson

5.48/3,712

99,400

1 3 2 2 2

McKay Engle Rushworth Rushworth Rushworth

5.71/3,886

98,200

127

1/26/65 2/2/65 2/15/65 2/15/65 2/17/65

5.27/3,511

95,100

128 129 130

2/19/65 2/25/65 2/26/65 3/26/65 4/23/65

1 1 1 1 3

McKay McKay McKay Rushworth Engle

5.40/3,750 5.17/3,580 5.48/3,580

153,600 101,900 79,700

119 120

121

126

20

REMARKS track (D'elamar to Edwards) and approximately same time, to compare photo distortion at high speed; test of new navigation instrument, cross range indicator camera system malfunctioned Mud Lake launch; inertial platform malfunction; shut-oft SAS and alternate SAS to evaluate basic stability at reentry Hidden Hills launch; heat transfer experiment with sharp leading edge upper vertical; damper off controllability mission Delamar launch; control and stability with upper speed brakes only; first test of hypersonic brake system; ablative test Hidden Hills launch; 3rd test with sharp fin in heat transfer studies Hidden Hills launch; 4th test with sharp fin in heat transfer studies; microphone embedded in skin to record turbulence noise in fatigue studies v/h computer malfunction Delamar launch; optical degradation test due to high temperature; secret recce camera tested (possibly linked to SR-71) mounted in forward fuselage stable platform malfunction Delamar launch; optical degradation test Delamar launch; inner panel of right windshield cracked; optical degradation test due to turbulence liquid oxygen pressure regulator malfunction Hidden Hills launch; 5th test with sharp fin in heat transfer studies; skin microphone to record BL noise Delamar launch; induced turbulence phase II; optical degradation test; altitude buildup for Engle premature engine shutdown at 41 seconds causing emergency landing at Cuddeback; Silver Lake launch communication, SAS, APU, and cabin pressure problems first captive flight of X-15A-2 no. 2 APU would not start first fiight of X-15 #2 following modification to X-15A-2 configuration; Hidden Hills launch; no tanks Delamar launch; stable platform failed to launch Honeywell system malfunction Delamar launch; Honeywell adaptive flight control system problems; first test of IR horizon scanner second stage igniter indicator malfunction Hidden Hills launch; flight test of heat resistant materials attached to fuselage Hidden Hills launch; structural fatigue studies; BL noise measured with skin microphone; measure airflow and heat transfer rates on two rippled panels on fuselage underside Delamar launch; nose landing gear extended at Mach 4.5 at 78,000 feet; both tires blew on landing; heat damage in nose landing gear well Hidden Hills launch; 4 instrumented panels (2 smooth, 2 rippled) carried on fuselage undersides to study heat transfer; BL noise studied with skin microphones Hidden Hills launch; airflow and heat transfer on instrumented rippled panels under fuselage; BL noise lost canopy pressurization due to canopy seal failure Delamar launch; ablative on lower tail tested for possible use on X-15A-2 Mud Lake launch; nose landing gear scoop door opened prematurely SAS malfunction Hidden Hills launch; first flight with wing tip pods, X-20 INS, etc. nose landing gear extension check Hidden Hills launch; fire warning light on after launch and off before landing nose landing extension check, 5.4 seconds nose landing gear extension check, 2.7 seconds Hidden Hills launch; main landing gear "toe in" noted after landing vented unknown quantity of ammonia Hidden Hills launch; boundary layer noise and shear layer measurement; airflow around nose; skin heating Delamar launch; Honeywell checkout and tip pod stability; instrument checks in wing tip pods, densitron, daylight sky experiments; sample upper atmosphere density; check of revised Honeywell INS Hidden Hills launch; "hot" experiments (ablative tests, airflow and heat studies); speed brake problems due to heat Hidden Hills launch; "hot" experiments and airflow data; immediately after burnout, aircraft gyrated about all 3 axes for 6 to 8 seconds IFDS malfunction Delamar launch; "hot" experiments; ablative tests captive gear check (no gear coverage) captive gear check Mud Lake launch; right half of main landing gear extended at 2,880 mph at 88,000 feet; normal landing lost no. 2 APU gear case pressure uprange weather Delamar launch; check-out revised LG; IFDS check out Delamar launch; IR scanning radiometer; check Honeywell INS Hidden Hills launch; heat transfer experiments with distortion panel ablative test

FLIGHT #

DATE

AlC#

ALT. (FT .IMS)

REMARKS

4/28/65

2

PILOT McKay

MACH/MPH

131

4.80/3,273

92,600

5/11/65

1 2 2

Thompson McKay McKay

5.17/3,541

102,100

Hidden Hiils launch; star tracker check out; stability and control; test MLG mods SAS roll drop-out lost cabin pressurization Mud Lake iaunch; stability and control; star tracker check; landing gear mods check; advanced X-15 landing dynamics; no ventral

Thompson

4.87/3,418

179,800

5/13/65

132

5/18/65

133

5/25/65

134

5/28/65

3

Engle

5.17/3,754

209,600

6/4/65

2 2 2 3

McKay McKay McKay Engle

4.69/3,404

244,700

Thompson

5.14/3,541

108,500

6/8/65 6/11/65

135

6/16/65

136

6/17/65

137

6/22/65

2

McKay

5.64/3,938

155,900

6/29/65

3 2 2

Engle McKay McKay

4.94/3,432

280,600

5.19/3,659

212,600

3 3 1 1 1 2

Rushworth Rushworth Thompson Thompson Thompson Rushworth

5.40/3,760

105,400

5.16/3,602

208,700

1 3

Thompson Engle

5.15/3,534 5.20/3,550

103,200 271,000

Thompson Thompson Thompson

5.11/3,604

214,100

3 2

Rushworth McKay

4.79/3,372 5.16/3,570

239,600 239,800

1 3 1 3

Rushworth McKay Rushworth McKay

5.25/3,534

5.33/3,732

97,200 239,000 100,300 295,600

1 1 3

Knight Engle Knight

4.06/2,718

76,600

4.6213,108

94,400

Engle

5.0813,554

266,500

138

7/2/65

139

7/8/65

7/13/65

140

7/20/65 7/23/65 7/27/65 7/28/65

141

8/3165

142 143

8/6/65 8/10/65 8/20/65 8/24/65

144

8/25/65

145 146

8/26/65

147 148 149 150

9/9/65

151

9/2/65

9/14/65 9/22/65 9/28/65

9/30/65 10/8/65

5.03/3,519 5.18/3,550

152

10/12/65

153

10/14/65

154

10/27/65

3

McKay

5.0613,519

236,900

11/2/65

1 2

Dana Rushworth

2.31/1,500

70,600

Dana

4.22/2,765

80,200

Rushworth Rushworth Rushworth McKay

2.21/1,434

68,400

5.43/3,689

99,000

155

11/3165

156

11/4/65

157

5/6/66

2 2 2 1

158

5/18/66

2

Rushworth

6/2/66

6/27/66

1 1 3 2

McKay McKay Dana Rushworth

159

7/1/66

2

Rushworth

1.54/1,023

45,000

160

7/12/66

Knight

5.34/3,652

130,000

4.71/3,217

96,100

4/13/66 4/20/66 5/5/66

6/10/66 6/20/66

7/18/66

3 3

Dana Dana

7/20/66

2

Knight

7/13/66

161

Mud Lake launch; MIT horizon photometer (UV scanner) checkout; high altitude pilot checkout; INS checkout; RAS Mod check; Pace transducer on left wing to measure air density Delamar launch; 2nd test of horizon scanner; UV and IR experiments (NSL radiometer & Langley scanner); BL noise lost cabin pressurization low source pressure low source pressure Delamar launch; IR and boundary layer experiments with microphone on vertical fin Delamar launch; IR and cross track vernier; further checks on new INS Delamar launch; star tracker (4 x 35mm cameras mounted on top of X-15); test landing gear mods; stability and control maneuvers; advanced landing dynamics; ventral off Delamar launch; horizon scanner; astronaut wings flight stable platform malfunction Delamar launch; star tracker mission to photograph Gamma Cassiopeia; 3rd brightest star in constellation Cassiopeia, film sent to University of Wisconsin lost cabin pressurization Delamar launch; boundary layer noise; IR scanner suit/faceplate leak no uprang'e X-15-to-ground station communication "Q ball" ~ose malfunction Delamar launch; RAS (Reaction Control Augment System) checkout; UV experiments; advanced landing dynamics Delamar launch; IR, stability and control, and ablative experiments Delamar launch; NSL scanner; BL noise; reentry maneuver techniques; astronaut wings flight lost cabin pressurization computer failure Delamar launch; MIT scanner; stability and control; Pace transducer Delamar launch; NSL radiometer; BL noise Delamar launch; star tracker (obtain info on lighting conditions); RAS check; advanced landing dynamics Delamar launch; IR and ablative experiments Delamar launch; NSL radiometer Delamar launch; IR experiments Delamar launch; NSL scanner; astronaut wings flight; upgraded XLR99 used; aerodynamiclstructural loads test on horizontal tail surface; BL noise Knight's first flight; Hidden Hills launch; IR scanner pitchlyaw RAS valve faiied Hidden Hills launch; no. 2 APU quit after launch and was restarted after burnout Delamar launch; MIT horizon scanner; astronaut wings flight; measure microscopic ATM pressure Delamar launch; boundary layer noise; NSL radiometer; horizontal tail loads cabin pressurization and telemetry lost first flight with external tanks (empty); Cuddeback launch; Ammonia (rl.) tank recovered; LOX (11.) tank destroyed Dana's first flight; Hidden Hills launch; engine restart; modified MIT scanner; measure microscopic ATM pressure stable platform failed SAS/yaw channel failure SAS/yaw channel failure premature engine shutdown at 32 seconds caused by engine turbopump failure; emergency landing at Delamar Lake no tanks; liquid hydrogen jettison valve leak; stability and control flight with ventral fin (probed) stable platform malfunction stable platform malfunction inertial system computer full tank fuel flow; SAS problems; B-52 right outrigger would not retract first heavy tank flight for X-15A-2 (to test fuel flow transfer and tank separation); engine shutdown at 34.6 seconds due to fuel flow anomalies; at 22.3 seconds after launch, throttled back; 29.5 seconds, jettisoned external tanks; 34.6 seconds, shut down engine; engine light at 1.2 seconds after launch; Mud Lake landing; ablatives (pink spots) tested for later use; ventral on landed shorter than usual; check electrical loads, non-glare glass, stick pusher, shade window telemetry system problems Hidden Hills launch; stick pusher and third skid; first flight of energy management system (Litton displays-round insl. replaced by vertical); Bell Aerosystems was system integrator; computer by Honeywell weather at Delamar

21

FLIGHT #

DATE

A/C #

PILOT

MACH/MPH

ALT. (FT.IMS)

REMARKS

162

7/21/66

2

Knight

5.12/3,568

192,300

163

7/28/66

1

McKay

5.19/3,702

241,800

164

8/3/66

2

Knight

5.03/3,440

249,000

165

3 1 1 1

Dana McKay McKay McKay

5.34/3,693

132,700

166

8/4/66 8/9/66 8/10/66 8/11/66

5.21/3,590

251,000

167

8/12/66

2

Knight

5.02/3,472

231,100

168

8/19/66

3

Dana

5.20/3,607

178,000

169

8/25/66

1

McKay

5.11/3,543

257,500

170

8/30/66

2

Knight

5.21/3,543

100,200

171

9/8/66

1

McKay

2.44/1,602

73,200

172

9/13/66 9/14/66

3 3

Dana Dana

5.12/3,586

254,200

173

9/28/66 10/4/66 10/6/66

1 1 1

Adams Adams Adams

3.00/2,900

75,400

174

10/7/66 10/19/66 11/1/66

2 2 3

Knight Knight Dana

5.46/3,750

306,900

175

11/18/66 11/18/66

3 2

Dana Knight

6.33/4,250

98,900

11/23/66 11/29/66 12/22/66 3/15/67 3/21/67 3/22/67

3 3 2 1 1 1

Dana Adams Knight Adams Adams Adams

4.65/3,120

92,000

5.95/3,822

133,100

4/20/67

1

Adams

Delamar launch; no external tanks; checkout 01 star photo equipment (doors opened at 172,000 ft.); base drag studies no H-Dot; Delamar launch; MIT scanner micro-meteroite; Pace transducer check; non-glare glass check; stick pusher Delamar launch; no external tanks; UV photos 01 stars in Auriga Constellation Mud Lake launch; BL noise (sharp lin) inertial platform and leaky BCS valve no helicopter at launch lake Delamar launch; collect micrometeorites and extra·terrestrial dust; maneuvers to check horizon scanner; check electrical loads; wing pod IIutter Delamar launch; no external tanks; UV photos 01 star Alpha Gemini in Constellation Gemini; base drag studies; test alternate pitot static Delamar launch; altitude buildup; BL noise (sharp lin); horizontal tail loads; heat transler panels T/M dropped out Irom 2 seconds until 5 minutes after launch; Delamar launch; collect micrometeorites and extra-terrestrial dust; maneuvers to check horizon scanner; check electrical loads; wing pod flutter no external tanks; lower ventral chute lost in turn towards Edwards AFB; Delamar launch; heat tests; base drag studies; stability and control; ablatives; wing tip acce!. meters low pressure luel indication; engine shutdown at 38 seconds; landed at Smith Lake; Smith Ranch launch; purpose 01 mission was to scan horizon radio malfunction Delamar launch; JPL spectrometer to measure solar flux and broadband light distribution radiometer to measure exhaust plume UV characteristics (required for surveillance satellite sensors); micrometeorite collection uprange weather loss 01 cabin pressurization Adams lirst flight; XLR99 shut down at 90 seconds; landed at Cuddeback telemetry malfunction nitrogen pressurization problem Smith Ranch launch; astronaut wings flight; micrometeorite collection; maneuver perlormed to checkout dual channel radiometer tip pod acce!. meter (precision attitude) attitude indicator malfunction unofficial world speed record lor class; Mud Lake launch; maximum q was 857 ppl; ventral on (probed) hot APU bearing temperature Hidden Hills launch; stability and control maneuvers captive inspection flight weather inertial platform mallunction inertial system dropped out after peak altitude was reached; loss 01 cabin pressurization during descent; high altitude flight experience lor Adams; record stabilator angle-ol·attack flow; test electrical loads; check·out 3rd landing skid; check-out pressure attitude indicator; record X-15 sonic boom; check-out ablatives on stab.; Mud Lake launch up-range weather

178

4/26/67

3

Dana

1.80/1,163

53,400

179

4/28/67

1

Adams

5.44/3,720

167,000

180

5/5/67 5/8/67

2 2

Knight Knight

4.75/3,193

97,600

181

5/12/67 5/17/67

3 3

Dana Dana

4.80/3,177

71,100

182

5/26/67 6/1/67 6/14/67 6/15/67

1 1 1 1

Adams Adams Adams Adams

5.12/3,606

229,300

183

6/22/67

3

Dana

5.44/3,611

82,200

184

6/29/67

1

Knight

4.17/2,870

173,000

185

7/20/67

3

Dana

5.44/3,693

84,300

8/7/67 8/11/67 8/16/67

2 2 2

Knight Knight Knight

176

177

22

low inlet pump pressure; launched and ianded at Silver Lake; mission to conduct heat tests Delamar launch; high alt. experience lor Adams; record stabilator angle-ol-attack Ilow; test electrical loads; check-out 3rd landing skid; check·out horizon scanner no external tanks; up-range weather no external tanks; dummy chute; Hidden Hills launch; "eyelid" in place; check-out stability and control with ramjet; check-out thermocouple system; check airflow around ramjet and support ass'y; ramjet recovery chute cable broke ball nose and radio malfunction Silver Lake launch; check-out PCM system, boost guidance, energy management system and tip pod accelerometer; measurement of cold wall heat transler and stepped panel heat transler; study 01 sonic boom and horiz. tail loads 3-axis gyro malfunction stable platform malfunction X-15 radio transmitter malfunction PMR experiment; Delamar launch; check-out horizon scanner; check-out horizon stabilator blow-off panel; Hidden Hills Ranch launch; check·out 01 PCM system, boost guidance, energy management system and tip pod accelerometer; measurement 01 cold wall heat transler and stepped panel heat transler; study 01 sonic boom and horiz. tail loads electrical lailure while climbing through 107,000 feet; lost engine and APU's at 69 seconds; restarted APU; landed at Mud Lake with no flaps; Smith Ranch launch; pilot injured during egress; mission was to check-out horiz. scanner Hidden Hills launch; check·out 01 PCM system, boost guidance, energy management system and tip pod accelerometer; measurement 01 cold wall heat transler and stepped panel heat transler; study 01 sonic boom, NLG loads and horiz. tail loads X-15 LOX vent valve stuck open; hard over servo signal liquid nitrogen leak helium leak caused by decompression 01 hydrogen peroxide tanks

FLIGHT #

DATE

A/e #

PILOT

MACH/MPH

ALT. (FT.lMS)

REMARKS

186

8/21/67

2

Knight

4.94/3,368

91,000

187

8/25/67

3

Adams

4.63/3,115

84,400

9f22/67

3 2

Dana Knight

6.70/4,520

102,100

first flight with full MA-25S ablative coating; second engine light; Hidden Hills launch; investigate stability and control with ablative (pink), sealant (white), and ramjet before speed record attempt; investigate ramjet local flow and ramjet separation characteristics; flew Hycon phase 2 camera (KA-51 A); wingtip accelerometer; ramjet with ablative and 20° cone Hidden Hills launch; phase 2 tail loads-two types of loadsthermal & aerodynamic (the horiz. tail was instrumented to record thermal stresses and loads); boost guidance check-out; traversing probe; tip pod accelerometer; BL noise up-range weather unofficial world absolute speed record; aircraft flown with full ablative coating; external tanks; dummy ramjet mounted on ventral; mechanical eyelid over windscreen; investigate ramjet local flow; check stability and control with ramjet, ext. tank separation; wingtip accelerometer; conduct fluidic temp. probe (first flight); ramjet with ablatives and 40° cone probes; Mud Lake launch Smith Ranch launch; camera in lower bug-eye for wingtip pod deflection measurements; test Ames boost guidance technique; to check UV exhaust plume (tail box); JPL solar spectrum measurement (air density, x-ray air density and tip pod camera); micrometeorite collection astronaut wings flight; camera in lower bug eye for wingtip pod deflection measurements; test Ames boost guidance technique; to check UV exhaust plume (tail box); JPL solar spectrum measurement (air density, x-ray air density, and tip pod camera); micrometeorite collection astronaut wings flight; fatal accident; aircraft destroyed; 65th flight of X-15 #3; camera in lower bug-eye for wingtip pod deflection measurements; test Ames boost guidance technique; to check UV exhaust plume (tail box); JPL solar spectrum measurement (air density, x-ray air density, tip pod camera); micrometeorite collection; first use of traversing probe on .right wingtip pod cone to measure stand-off distance of the bow shock wave ahead of pod; test adhesive and insulating properties of a cryogenic insulating material for Safurn 5 2nd stage, mounted on left upper speed brake; camera in upper bug-eye test new spray foam cryogenic insulation for Saturn 5 2nd stage, beginning with vehicle 8 (mounted on speed brakes); check out elect. system test new spray foam cryogenic insulation for Saturn 5 2nd stage, beginning with vehicle 8 (mounted on speed brakes); first flight with second set of tip pods repeat Saturn 5 insulation test; horiz. scanner; fixed ball nose WTR exp.; horizon scanner exposure and retraction of WTR experiment not accomplished due to abnormally low hydraulic pressure and severe vibrations preventing required altitude from being reached astronaut wings flight; fixed angle-of-attack ball; conduct WTR experiment; check horiz. scanner; sky brightness; fluidic probe WTR experiment; horiz. scanner; fixed angle-of-attack ball nose; fluidic probe last X-15 flight; WTR experiment; fixed angle-of-attack (ball nose) cone; fluidic probe

188

~

115Jo y \\5; _ c,,?

1'1)

(,0

.

189

10/4/67

3

Dana

5.53/3,897

251,100

190

10/17/67

3

Knight

5.5313,856

280,500

191

11/15/67

3

Adams

5.2013,570

266,000

192

3/1/68

Dana

4.36/2,878

104,500

193

4/4/68

Dana

5.27/3,610

187,500

194 195 196

4/26/68 6/12/68 7f16f68

Knight Dana Knight

5.00/3,545 4.96/3,545 4.74f3,409

207,000 214,000 218,500

197

8f21f68

Dana

4.71/3,443

264,000

198

9/13/68

Knight

5.09/3,716

250,000

199

10/24f68

Dana

5.04f3,682

250,000

~

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Prior to its first glide flight, X-15, 56-6670, is seen shonfy before being uploaded under the NB-52A wing pylon. At this stage, the aircraft was equipped with the twin-pack XLRll propulsion system and the long nose sensor boom. Visible is a calibration device for refining the slab stabilatar control input values.

23

CONSTRUCTION AND SYSTEMS: The external appearance presented by the X-15 was that of a rather conventional mid-wing monoplane equipped with low-aspect-ratio wings, large dorsal and ventral vertical tail surfaces, and a tubular fuselage to which were attached chine-like tunnel fairings. Unlike more conventional aircraft, however, the materials used in the X-15's construction were selected on the basis of their compatibility with strength, temperature environment, corrosion resistance, and processing requirements. It was obvious from the start to North American chief structures engineer R. L. Schleicher, Ph.D., that much of the wetted surface of the aircraft would be subjected to temperatures of from 800° to 1,200° F. and would require high strength at these temperatures. Exotic materials utilizing the rare elements had not advanced to quantity production or usage and consequently the list of possible materials candidates narrowed to the corrosion resistant steels, titanium, and nickel base alloys. Though steel and titanium had higher temperature efficiencies over a wide temperature range, each fell off rapidly above 800° F-while Inconel-X showed only a gradual drop in strength up to 1,200° F. Because of this stability, Inconel-X was selected for the skin covering for the entire aircraft. The internal structure was made from a variety of materials. High strength aluminum alloy (2024-T4) was used to form the inner pressure shell of the cockpit and part of the equipment bay. As a relief from high thermal stresses in the internal structure of the lifting surfaces and fuselage, titanium alloys were used there, too. These originally were of two compositions, mainly 8 Mn titanium which, though not suitable for welding, was the highest strength alloy then available, and 5A1-2.5Sn which had good strength and was weldable. Later, a high strength and weldable titanium alloy (6A1-4V) was introduced. Titanium framing was used almost exclusively in the aft fuselage structure where high concentrated loads were expected. Fusion welding was used predominately throughout the construction but always in a controlled atmosphere. Resistance welding also was used. All critical welds were radiographically inspected to assure high quality and soundness. Most of the internal volume of the X-15 consisted of tanks, or vessels, to accommodate fuel, oxydizer, and miscellaneous liquid and gaseous requirements. The materials used in many of these pressure vessels had to be selected not only on the basis of strength, but also ductility. Martensitic alloys, such as heat-treated 4130 low alloy and 350 precipitation hardening corrosion resistant steels, exhibited a considerable loss in ductility as the temperatures they were expected to be exposed to (liquid oxygen) continued to decrease to large negative values. Titanium alloys showed more favorable characteristics in this respect, but were used sparingly, depending on the environmental requirements of the vessel. The various vessels for accommodating helium and hydrogen peroxide ranged in size from 6 to 32 inches in diameter for spherical shapes. Two other vessels, one a 14 x 96 inch cylinder and the other an elastomer bladder-lined prolate spheroid measuring 15 x 28.5 inches, also served to accommodate these gases. Temperatures varied from -300° F. for the cryogenic oxidizer to + 160° F. ambient in the engine compartment. In the design of the vessels, considerable importance was placed on the relation of working pressure to the yield strength of the materials used, and to design proof pressure. Proof, cyclic, vibration and burst pressure tests were conducted on each. Weight and balance were considered extremely critical for the X-15. It originally was estimated that a vehicle weighing approximately 30,000 Ibs. fully loaded and 12,000 Ibs. without fuel would be required to perform the NACA specified missions. After the basic configuration was agreed upon, a vehicle design estimated to have a fully loaded weight of 31,2751bs. (wing 1,406Ibs.; empennage 1,078 Ibs.; fuselage 3,812 Ibs.; landing gear 447 Ibs.; and surface controls 937Ibs.) emerged, this eventually being only 400 Ibs. less than the fUlly loaded weight of the actual X-15. Maintaining the X-15's center of gravity requirements throughout its flight envelope proved exceptionally challenging. Expending nine tons of propellants in a matter of seconds, and maintaining a nearly constant center of gravity location required serious consideration during design. The annular-type liquid oxygen tank with torus shaped bulkheads and three compartments, containing 1,034 gallons (10,400 Ibs.) of oxidizer, was located for-

24

ward, and the similarly constructed anhydrous ammonia tank with 1,445 gallons (8,400 Ibs.) of fuel, was located aft of the center of gravity. Each tank was divided by semitorus frames into three compartments. In the center of the liquid oxygen tank's annular ring was another tank containing high pressure gaseous helium. Other helium tanks, spherical in shape, were located throughout the rest of the aircraft for use as source pressure, engine purge, control gas, and emergency jettison. The seven cubic feet of helium gas at 3,600 psi pressure used to force the LOX into the oxydizer lines had almost negligible weight. Center of gravity control was established by expelling the LOX and ammonia towards the center of gravity location with the LOX expelled aft through the tank compartments and the ammonia expelled forward in the same manner. In this way, the movement of the c.g. during powered flight was held to within 3.5%. No other weight or balance problem developed which required special considerations. The ammonia tank combined an annular tank and a core tank. To the rear of the ammonia tank was a spherical tank for hydrogen peroxide, this having a total capacity of 77.5 gallons. The tank incorporated a swiveltype pickup feed line that permitted positive feeding regardless of aircraft attitude. A combination vent pressure relief and tank pressurization valve was mounted on top of this tank. When the engine was not in operation, the system was vented into the atmosphere, and when the engine starting sequence was begun, the system was pressurized by helium gas, forcing the peroxide to gas generators which provided steam power for the turbopump operation. There also was a jettison valve which permitted the peroxide to be forcibly expelled overboard by the helium gas pressure. Just forward of the liquid oxygen tank was the auxiliary power unit compartment and equipment bay which contained flight test instrumentation. Also just forward of the LOX tank was a supply of liquid nitrogen for the cooling system. Since considerable liquid oxygen boiled off during taxi, climb, and cruise to the launch point, it was necessary to carry additional liquid oxygen in the carrier aircraft (B-52). This tankage was used to top off the liquid oxygen tanks in the X-15 just before launch. Two liquid oxygen tanks were utilized, one called the "cruise tank" and the other, the "climb out tank". Again, spherical tanks of helium furnished gas to force the liquid oxygen into the X-15 under pressure. Float level valves in the X-15 automatically shut off supply when the tanks were full. The 1,000 gallons carried in the "cruise tank" and 500 gallons carried in the "climb out tank" were sufficient for a prelaunch cruise of about 2 hours. Since the mission of the X-15 was manned flight at extremely high speeds, altitudes, and temperatures, the design criteria centered around basic mission profiles. The first mission was based on flight at 250,000 feet and a velocity of 6,300 feet per second. Two types of pUll-out were considered, each maintaining a zero lift trajectory until pull-out. One type of reentry called for a maximum angle-of-attack entry wherein the speed brakes remained closed closed and pull-out was initiated at a predetermined altitude (the highest where available lift and control power permitted a 7.33g maneuver). In other types of reentry, the speed brakes were to be deployed at the peak altitude and a 7.33g recovery was initiated at a point so as not to exceed a limit dynamic pressure of 2,500 Ibs. per square foot. To attain a true airspeed of 6,600 feet per second, the pull-up after launch was to be made to a lower climb angle than for the altitude mission. The design speed was to be reached at burn-out, from which a zero lift coast was to be made to approximately 130,000 feet. From this altitude, recoveries similar to the high altitude missions were to be made. To provide the X-15 with a reasonable strength level, North American engineers designed the aircraft to accommodate limit maneuver factors of + 4.0g and - 2.0g before burn-out and 7.33g and - 3g after burn-out. The limiting dynamic pressure of 2,500 Ibs. per square foot was chosen as representing reentry at the lowest altitude for recovery consistent with safety for terrain clearance. Additionally; in order to avoid an unnecessary weight penalty dictated by increased structural requirements, the pUll-out at 7.33g at maximum dynamic pressures were assumed to be attainable only once during a particular recovery. During this maneuver it was determined that the X-15 would slow down appreciably and heat up rapidly. If another pull-up was required following the first, it would have to be made at a lower acceleration to avoid overloading the heated structure.

It was also concluded that an apparent air load problem could be expected during the launch phase with the X-15 suspended under the wing of the B-52. Performance and environment limitations allowed this anomaly to be controlled without further changes or aircraft modifications. The X-15's wings were trapezoidally configured surfaces with extremely thin (5%) sections and an aspect ratio of 2.5. The 25° (at quarter chord) swept leading edge was thin, but not extremely sharp, and the trailing edge was blunt (with a thickness of apprOXimately 2-inches at the root and .375-inches at the tip). There were no ailerons fitted (roll control being accommodated by the horizontal slab stabilators), but there was one small, hydraulically-actuated flap located at the inboard trailing edge of each wing. The wing outer panel configuration was a conventional multispar box beam design. The skins were reinforced Inconel-X sheet chosen for strength and favorable creep characteristics at 1,200° F. The skin thickness varied from .090 inches at the root to .040 inches at the tip of the upper surfaces and .065 to .040 respectively for the lower surface. The internal structure was built entirely of titanium alloy sheet and extrusions. Both the front and rear spars consisted of flat web channel sections. The intermediate spars had corrugated webs which were attached to the skin through separate angles. The three ribs used in the design were of the same construction as the front and rear spars. Attachment of the above was by means of A-286 rivets and screws. The leading and trailing edges noted earlier were of multi-rib design where panel size was determined by stiffness requirements. The extreme leading edge itself consisted of a solid bar of Inconel-X which acted as a heat sink. The Inconel-X leading edge was originally divided into five segments to minimize thermal stresses. Later, because of hot flight anomalies, this figure was increased to nine. The wing stress analysis involved both simple beam theory as well as the solution of the redundant root structure. Outboard of Station 89 (approximately mid-span), a simple cantilever eight cell beam solution was sufficient to determine load stresses. Inboard of Station 89 the wing was partitioned into nine individual box beams and a single cell torque box to which were applied the redundant shears and moments from the structure surrounding the wing root rib. The fuselage/wing attachment assembly was bimetallic with Inconel-X wing and side fairing covers and titanium alloy (5A1-2.5Sn) forgings and extrusions forming the "A" frame details. Throughout the design of the wing, many artifices such as beads, lightening holes, scallops, and corrugated webs greatly minimized internal loads and thermal strains. The two slab stabilators, having the same airfoil section and 5% thickness/chord ratio as the wings, were mounted separately at the aft end of each fuselage tunnel/chine with 15° of anhedral and could thus be differentially moved in order to provide roll inputs as well as symmetrically moved for pitch control (though comonplace today, this system was innovative when the X-15 was designed, possibly representing the first use of allmoving slab stabilators in a high-performance aircraft). Their structure consisted of an Inconel-X main spar, an A-286 front spar, a titanium trailing edge, Inconel-X ribs ahead of and 8 Mn titanium ribs aft of the main spar, and .050 inch Inconel-X skin. Each surface was mounted on its own spindle which was an integral part of the main spar and which attached to the fuselage in the region of the side tunnels. For the most efficient design, the main spar was used to carry all normal bending along the entire span of the stabilator. The front spar effectively closed out the torque box which terminated at the root rib. Actuation of each stabilator was by a hydraUlic cylinder attached to an arm located in line with the outboard bearing. During work at North American, it was discovered that for even small temperature differences, it was necessary to increase the stabilator skin gauge and decrease the stiffener spacing in order to eliminate skin buckling. This would impose severe weight penalties. For example, to increase and stiffen the skin to prevent buckling up to limit conditions would add 195 Ibs. Consequently, thermal stress buckles were permitted to exist during the brief period of heating during reentry, but no permanent buckles were condoned. A skin thicknes$ of .050 was finally selected with no stiffeners. The vertical tail surfaces consisted of a fixed box structure with a mixture of Inconel-X and titanium alloy with Inconel-X skin covering. The speed brakes, located at the

root section of each surface, were hinged from the fixed portions and each pair was actuated by a hydraulic strut. The brakes utilized Inconel-X skin covering reinforced by a corrugated inner skin and ribs made of the same material. Above the upper fixed structure was an allmovable section made entirely of Inconel-X and employing front, main, and rear beams. The main and rear beams plus skin formed a sixth beam for attachment to the fixed portion. A spindle support having two bearings spaced 18-inches on center completed the attachment. A hydraulic strut was used to actuate the all-movable portion of the surface. A section of the lower vertical surface below the fixed portion was jettisonable. While attached, it was all-movable as was the case of its dorsally mounted counterpart. Release of the lower section, triggered either by landing gear extension or by pilot command, was by means of explosive bolts. After release, a parachute was deployed to lower it to the ground for reuse. The fuselage was divided into three distinct structural sections. The forward section extending from the nose to the forward end of the LOX tank was semi-monocoque. The regions surrounding the cockpit and parts of the equipment bay utilized a double wall construction. The outer skin was Inconel-X and the inner wall was 2024 T4 aluminium alloy. Between the two was spun glass matting for insulation. The inner wall was used only as a pressure seal. The intercostals connecting the inner and outer skins were titanium alloy. The center section of the fuselage formed the propellant and oxidizer tanks. Since no insulation was used, a pure monocoque structure was utilized. This portion of the fuselage was designed by considering critical conditions during both flight and landing. The monocoque construction obviously simplified the design to a very large extent and eliminated many of the weight and thermal stress problems that might have resulted from a more complex configuration. Provisions were made, however, to relieve thermal stresses in the side tunnel/chine areas by using partial circumferential beads in the skin. The tunnel/chines were at first designed as contiguous surfaces throughout their length. Actual flight experience wherein residual buckles developed, however, later led to the incorporation of expansion joints on all three aircraft. These fuselage fairings served not only to house all control cables, electrical leads, plumbing, and hydraulic lines, but also to increase the aircraft's total aerodynamic lift. The airflow near these fairings provided well over half of the X-15's total lifting force at hypersonic speeds. The oxidizer, or LOX tank section of the fuselage was first analyzed for an internal pressure of 111 psi ultimate. Next it was analyzed for critical external loads which included the normal shear, axial, bending, and torsional loadings. Finally, it was examined for negative or collapsing pressure, this reading an ultimate of 6 psi. The high load carrying sectors, namely the upper and lower Inconel-X segments, had a thickness of .063 inches which was stiffened by light welded-on "J" section ring frames with an area of .0355 inches and an average spacing of 6 inches. The maximum temperature in the LOX tank at the inner torus and cylinder was 307 0 F. and the minimum temperature when filled with LOX was -230 0 F. The LOX tank also was SUbjected to fore and aft inertia loading from its contents. These loads were transferred to the outer shell wall by means of three radial flat panels acting as baffles. The baffles, in turn, were supported from the inner shell wall. This inner shell supported also the cylinderical helium storage bottle as well as the tori bulkheads, thus adding to the general stiffness of the tank assembly. The design and analysis of the ammonia tank portion of the fuselage followed closely that for the LOX tank. The maximum pressure inside the ammonia tank was the same as for the LOX tank, but the minimum temperature was -24 0 F. The inner shell, however, was filled with fuel instead of serving as a container for a gas storage bottle. The aft section of the fuselage was designed as a semimonocoque structure along more conventional lines. The outer skin made from Inconel-X was riveted to heavy titanium framing. The latter was arranged to accommodate high concentrated loads introduced by the engine mount, landing gear, and empennage attachments. The landing gear for the X-15, based on experience acquired during the X-2 program, was an unconventional configuration consisting of a conventional nose wheel assembly and two boat-shaped steel skid-type main gear. The nose gear was stored in a conventional nose gear well, and the skid gear were stowed externally, under the aft portions of each tunnel/chine. The skids were

mounted on inflexible struts with an air/oil shock absorber attached to the upper end permitting some outboard rotation when the weight of the aircraft was on the gear. The extremely far aft location of the main gear was made possible by the fact that since the X-15 was to be air-launched, the usual nose wheel lift-off requiring rotation about the main gear was not an important consideration. All landing gear were retracted manually while the aircraft was suspended. Extension was via gravity freefall and aerodynamic down-loads. Landing gear design requirements included a sinking velocity of 9 feet per second, landing speeds at touchdown of between 164 and 200 knots, and an aircraft attitude of 6 0 (ground angle). The main gear skids were pinned in two planes to permit pitch and roll, but were restrained in yaw. The nose wheels were used to prevent shimmy and to offer less castoring torque resistance than a hydraulically damped system. Three landing gear test programs, other than dynamic drop tests of the shock absorbers, were undertaken during landing gear development. These included a dynamic model test for stability, a nose wheel shimmy test, and full-scale main gear skid tests on the Edwards AFB dry lake bed that would be used for landing. Full-scale nose wheel shimmy tests were later conducted at the NASA landing test facility at Langley Field, Virginia. These explored a velocity range from 20 to 125 mph and cond itions representing wet pavement, sand, uneven tire pressure, flat tires, and unbalanced wheels. Blocks placed in the path of the nose wheel were used to induce shimmy. These tests proved that neither shimmy dampers nor torque links were needed. Full-scale main skid tests also were conducted at Edwards. The skids were mounted on a retractable carriage unit fastened to a trailer and drawn by truck. Speeds up to 70 mph were attained with this test rig. When the test speed was reached, an electrically operated release dropped the skids on to the lake bed and high speed cameras recorded the motions. Other instrumentation recorded the vertical and drag loads, strut stroke, and other parameters. From the data gathered, the coefficient of friction of the landing surface was determined. The windshield was composed of a single outer pane of 3/8ths inch thick alumino-silicate plate glass (With a temper of 25,000 psi) and a double inner pane of sodalime plate glass and a silicone type "K" interlayer (with a temper of 14,500 psi). These were asbestos fabric cushion mounted in a retaining frame of Inconel-X with a titanium (6A1-4V) outer glass retaining strip and an aluminum alloy inner flange element. The maximum air pressure on the outer pane was 7.8 psi above 35,000 feet and 9.3 psi below that altitude, while the cabin presure was maintained at 3.5 psi. The outer surface of the glass was designed to reach a temperature of 800 0 F. while the inner surface of the outer glass was to reach a temperature of 550 0 F. The inner surface temperature, in actual practice, however, lagged behind the outer surface temperature. During the rapid heat up, a maximum temperature differential of 480 0 F. occurred at a time when the outer surface temperature was only 570 0 F. The mutual restraint of the glass panels and the supporting frame further complicated the differential, causing severe stress problems. Several panels failed during initial high speed flights and accordingly, a decision was made, following significant research into the problem, to change the outer glass panel to alumino-silicate and to change the outer retaining strip to titanium alloy. The X-15's cockpit was relatively unconventional, even by research aircraft standards. Among the many unique items were three control sticks. One was a conventional center-mounted stick which was used to control the aircraft in pitch and roll during low-speed atmospheric flight; the second was directly linked to the first through pitch and roll hydraulic boost actuators, but was mounted on the right side of the cockpit (thus making it a side-stick controller; it was tested aboard a North American F-107A [55-120) during May, 1958) and operated by wrist movement only; and the third was mounted on the left side of the cockpit and was used to control the aircraft during exo-atmospheric flight by metering, through valves, the peroxide utilized for thrust in the nose and wingtip mounted reaction jets. Interestingly, the central stick became somewhat of a redundancy and was removedas most pilots preferred the side-stick controller. The aerodynamic flight control system incorporated hydraulically actuated yaw, pitCh, and roll control cylinders. The system was irreversible with artifical feel furnished by bungee springs. Attached to this was the

stability augmentation system which provided damping inputs to the aerodynamic flight control Sy1em about all three axes. Major components of the system were a three-axis gyro, servo cylinders, and pilot-controlled gain selector switches. Cockpit air-conditioning and pressurization were furnished by a ram-air sy1em during takeoff to pre-launch, or by a liqUid-nitrogen system during X-15 flight. Helium gas was used to force the liquid nitrogen out of a segmented container and into a system of injectors where it became gaseous. A mixing chamber and blower were used to mix and recirculate the gaseous nitrogen continuously. Thermostats were used to control cockpit temperature by regulating the flow of nitrogen vapor. The pilot's full pressure suits were ventilated and pressurized by gaseous nitrogen. If cockpit pressure failed, the nitrogen supply automatically pressurized the suit to maintain a 35,OOO-foot environment. The temperature of the gaseous flow to the suit could also be warmed by a small electric heater. The suits were modified for the X-15 and had the restraining straps and parachute harnesses designed as an integral part. A neck seal was used to keep the suit pressurization of nitrogen and breathing oxygen separated. The nitrogen flow through the suit also served to cool the pilot's body. The oxygen regulator, suit pressure regulator, anti-g valve, and emergency oxygen supply were attached to the restraining layer. The helmet consisted of a fiberglass shell with a molded full head liner. All helmet oxygen and electrical services were internal with the helmet, and therefore, were not affected by high-speed, high-altitude bail-out blast effects. The entire X-15 supply of breathing oxygen was contained on the seat. The high pressure (1,800 lbs. per square inch) gaseous system had a capacity of 192 cubic inches, equally divided into two bottles located beneath the seat bucket. The ejection seat was rocket-propelled and was designed to permit safe pilot ejection from 90 kts up to Mach 4 in virtually any attitude and at any altitude up to 120,000 feet. For ejection, the ejection handgrips were pulled up and inboard, the end position being the pilot's lap. Simultaneously, the armrests were rotated inboard from the normal position, these laterally restricting the elbows. Foot manacles, which locked to a closed position when the pilot retracted his feet full aft, restrained the lower legs. Once the ejection levers had been squeezed, a system of mechanical linkages and an initiator fired the canopy remover which forcibly separated the canopy from the cockpit. Displacement of the canopy, in turn, fired another initiator which actuated the ballistic rocket catapult. This resulted in separation of the seat and pilot from the aircraft. The rocket motor exhaust nozzle was canted at an angle of 34° to direct the centerline of thrust through, or very close to, the center of gravity of the ejected mass. The burning rate of the two charges was designed to limit the magnitude and rate of onset of ejection forces to 20g and 250g/sec., respectively. During seat ejection along the guiderails, an interference tripper on the airframe engaged a seat-attached bell crank to arm the aneroid timer control for parachute deployment. Also actuated during this phase of ejection were the extendable boom-mounted stabilizing fins which became effective simultaneously with seat-rail separation. During the final phase of guided travel, the rocket engine was ignited to contribute an additional altitude increment, ensuring clearance over the X-15's vertical tail under any flight condition within the seat's envelope. Following ejection, the pilot remained with the seat during free fall to 15,000 feet-or if the ejection occurred below that altitude, for 3 seconds. The aneroid timer would then automatically release the pilot's restraint system and initiate parachute deployment by jettisoning the seat headrest. Transportation from altitude was via the 22-foot diameter main parachute canopy which had a terminal sea-level rate of descent of 22 feet per second for a 200 lb. man. The Bell Aerospace Textron designed reaction control jets used a modern form of superheated steam, created by the decomposition of hydrogen peroxide, for fuel. This steam (technically generated by the peroxide portion of the APU circuit) was exhausted through a series of 12 nozzles (four in the wingtips and eight in the nose) and created the pitch, roll, and yaw moments necessary to maneuver the aircraft when it was above the usable atmosphere. Each jet chamber was capable of producing from 40 to 100 Ibs. thrust. The X-15 was eqUipped with two auxiliary power units that were primarily for providing hydraulic and electric

25

power for control and flight systems following the termination of the powered portion of a flight. These were built by the General Electric Company and each one provided 40 hp and operated independently of the XLR99. They were driven by hydrogen peroxide which was decomposed into steam and oxygen. Electrical power provided by the two auxiliary power units through two alternatortype generators (400 cycle, 115 volts; two transformer rectifiers also were used to provide a 24-volt d.c. electrical system for other essential equipment) energized the X-15's instrumentation, powered the heating elements in the pilot's pressure suit and throughout the aircraf1, powered the inertial guidance system and computers, powered communications equipment, and powered various pieces of recording and telemetry equipment. The hydraulic system consisted of two 3,000 psi units operating independently of each other, but operating simultaneously. Dual, tandem hydraulic actuators were utilized so that failure of one hydraulic system would still permit the other to operate the various units. X·15 #3 was equipped with a special Honeywell MH-96 adaptive control system. This unit had the ability to sense the external atmospheric and aerodynamic conditions and provide the pilot with the required type of control, either aerodynamic or ballistic. A liquid nitrogen system designed and manufactured by the Garrett Corporation's AiResearch Division, provided cooling and pressurization for the cockpit and various instrument bays. liqUid nitrogen was used as the coolant. The Sperry Gyroscope Company provided the X-15's inertial guidance system. This provided information to the pilot for precise aircraf1 control. Among the instrument presentations furnished data by this system were critical attitude, velocity, distance, and altitude. A small computer digested and interpreted external cues. Later in the program, surplus inertial systems built for the ill-fated Boeing X-20A Dyna-Soar were utilized in the X-15. Not long af1er the X-15 began to explore its full performance envelope, a ball nose, or air direction indicator (sometimes referred to as the "a-Ball"), developed by the Nortronics Division of the Northrop Corporation, was installed in each of the three aircraft. This device, a hypersonic, spherical f1ow-direction sensor, was a null-seeking, hydraulically operated, electronically controlled servo mechanism. Pressure measurements, which were the sensor's sole inputs, consisted of measuring the differential pressure of each of two pairs of static-pressure ports located 42° from the reference line in the vertical and horizontal planes to determine flow-direction in terms of angle-of-attack and angle of sideslip, and measuring total pressure along the reference line to determine dynamic pressure. The ball nose, developed in response to a NASA requirement, was a 6.5 inch diameter sphere made of Inconel-X to resist the high temperatures encountered during X-15 flights. The overall length of the system package was 16.7 inches and the rear portion contained the mechanical and electrical components. These, along with the sphere, were cooled by vaporized liquid nitrogen as needed. When mounted in the X-15, the sensor was operated by utilizing the aircraf1's electrical, hydraulic, and coolant inputs. Electrically, the sensor operated on 28-volt d.c. and 115-voil, 400 cycle a.c. current; hydraulically, it used Oronite 8515 fluid at 3000 Ibs.lsq. in. Power consumption was 75 watts and 30 watts on the 28-volt and 115-volt circuits, respectively. The X-15A-2 incorporated numerous major and minor structural and mechanical changes beyond the stock X-15 configuration notes given above. The most significant external changes were the accommodations provided for the two externally mounted tanks. These jettisonable containers were designed to increase the X-15A-2's engine burn time by approximately 70%, thereby increasing the performance capability of the aircraft considerably. Each external tank was approximately 23.5-feet long and 38-inches in diameter. The left-hand tank, weighing 1,150 Ibs. empty, contained three helium bottles required for propellant tank pressurization in addition to a capacity of approXimately 793 gallons of liquid oxygen. The right-hand tank weighed 648 Ibs. empty and contained apprOXimately 1,080 gallons of anhydrous ammonia. The total weight of additional propellant to be carried in the external tanks was approximately 13,500 Ibs. Because of the difference in empty wieught and in propellant volumes, the lef1 hand tank was approximately 2,000 Ibs. heavier than the right at launch. The external tank jettison system contained two sets of fore and aft gas cartridges to eject the tanks from the

26

aircraft. In addition, the design included a solid propellant sustainer rocket on the nose of each tank to input a nosedown moment upon jettison to improve separation characteristics at supersonic speeds. For a normal empty tank jellison, both sets of gas cartridges were fired and the nose rocket was ignited. In the case of a requirement to make an emergency tank ejection while the tanks were still full, only one set of the gas cartridge ejectors were fired and the nose rocket was not activated. The high cost of the tanks dictated that they be reusable, hence each tank contained its own recovery system consisting of a drogue and descent chute. The drogue chute was deployed immediately after separation and the main descent chute deployment was initiated by a barometric sensor normally set for 8,000 feet. Other modifications to the basic X-15 airframe included a 29-inch extension added to the fuselage in the area of the center of gravity between the liquid oxygen tank and the anydrous ammonia tank. Tanks containing 48 Ibs. of liquid hydrogen for the scramjet engine were to have been installed in this area at a later date, but this was never undertaken. Additional hydrogen peroxide required for the extended engine propellant pump operating time was stored in tanks on the extended aft side fairings. A helium tank for additonal propellant pressurization gas was installed on the aft fuselage above the engine. The X-15A-2 also had a longer landing gear that provided ground clearance for landing with a scramjet engine installed. Since a functioning scramjet engine was not to be made available until much later in the flight test program, it was decided to take advantage of the increased landing load margin that could result from a shorter main gear during the initial portion of the flight test program. The strut of this interim gear was 6.75-inches longer than the standard X-15 gear. Drawing from the experience of the initial X-15 performance envelope expansion program when the standard windshield design suffered several glass fractures caused by thermal stress near the corners of the rectangular glass retainer, the X-15A-2 windshield was designed with an ellipitcal shape. In addition, three panes of glass were installed in the new design, instead of two as in the original X-15. Following a decision to proceed with an ablative coating on the aircraft, a single mechanical eyelid was devised that fit over the left windshield glass panes. This device could be opened at the pilot's discretion and was provided to give at least one external visual clue during landing. The eyelid itself was necessitated by the fact that the ablative material used on the rest of the X-15 outgassed and deposited ablatives on the windscreen during reentry. The eyelid covered the left windscreen and protected it from this unplanned coating. The eyelid was usually closed from launch through reentry. Other X-15A-2 modifications included a removable wingtip on the right wing that could be replaced with wingtips containing sensor equipment, or with wingtips constructed of unusual materials (a variety of experiments were in fact possible with this modification); a special retractable "dog leg" pitot tube that could be extended after reentry in order to avoid clogging by the off-gassing MA-25S ablator; and a "sky hatch" just behind the cockpit windshield/canopy fairing (this was a small compartment enclosed by two small doors that could be opened during flight in order to expose an experiment or sensor such as a camera). The X-15's primary mission, as a research aircraft, was to carry instrumentation and sensors for exploring and documenting, for future reference, the high speed and high altitude segments of the hypersonic envelope. The instrumentation built into the X-15 during construction (weighing some 1,300 Ibs. and consisting of some 160 pressure sensors, 700 temperature sensors, 110 strain gauges, and 96 acceleration, velocity, control position, angle, and physiological data points) was intended to provide measurements pertaining to what has come to be called the "basic program". This included the conditions imposed by flight of the aircraft itself, on the pilot, and on the system composed of the pilot/aircraft combination. The basic program concerned primarily the more familiar objectives outlined during the conceptual stages: (1) Verification of the predicted hypersonic aerodynamics, including heating rates. (2) StUdy of aircraft structures in a high heating and loading environment. (3) Investigation of stability and control problems encountered while leaving and reentering the earth's atmosphere. (4) Investigation of problems associated with weightlessness.

The advantage of using the X-15 to provide the environment desired in many scientific experiments was appreciated early in the flight program. The ability to escape the earth's atmosphere, for most practical purposes, and to return to earth with all the apparatus intact was especially attractive to researchers preparing for participation in the space program. Among the experiements eventually flown aboard the X-15 were the following: Ultraviolet stellar photography-cameras mounted on X-15 #2 were carried aloft to measure the energy emissions of various stars. Ultraviolet exhaust plume characteristics. Horizon seeking stabilization systemsallowed the investigation of the resolution or sharpness of the earth's horizon at different wave lengths. The nature of the visible spectrum in sunlight also was studied for application to the Apollo program. Optical degradation measurements. Infrared exhaust signatures. High-temperature windows. Atmospheric-density measuremer.ts at high altitude-developed and tested techniques of measuring ambient density for high-speed vehicles. Micrometeorite collection-equipment was mounted in a wing tip pod to collect samples of micrometeorite and extraterrestrial dust at high altitudes. Advanced integrated data sy1ems and energy management (including a landing computer)investigated the feasibility and compatability of combined air data probe, inertial guidance system, and pilot display information for possible use in future aerospace vehicles. Energy management also was studied. Vapor-cycle cooling. Rarefied-gas experiment. High-altitude sky brightness-an apparatus was mounted in a wing tip of one X-15 to measure the intensity, polarization, and spectral distortion of day1ime sky light at high altitude. Heat exchanger-though never actually flown, this test was to have evaluated performance analysis techniques for evaporators and condensors under zero-gravity conditions. Air breathing propulsion systems-a small hypersonic air-breathing engine, more often referred to as a scramjet (supersonic combustion ramjet) was scheduled to be tested while attached externally to the X-15A-2. Thrust and drag measurements were to be recorded from flight tests. Problems with the installation of the mockup units, coupled with the termination of the X-15 program, led to cancellation. Supersonic decelerators-investigated the deployment, stability, and drag characteristics of various bailon-decelerator configurations. High-temperature leading edges-various types and concepts concerning the dissipation of extreme temperature effects were studied utilizing the jellisonable lower ventral fin. Ablative materials-various types of ablators were tested on the X-15, including the previously discussed MA-25S, and the ablative tiles later utilized during the Apollo program (this research led directly to the heat-soak tiles now found on the Rockwell International Space Shutt/e). The latter were attached to the base portion of the dorsally-mounted vertical fin. High-speed, high altitude photography-three major and several minor test programs were conducted. These included an experiment employing two small aerial cameras installed in the center-of-gravity compartment of X-15 #2, which provided data permitting the investigation of contrast attenuation at high altitudes and showed the feasibility of obtaining aerial photography from supersonic vehicles; the installation in the instrument bay of the X-15 #1 of a KC-1 camera, fitted with the GEOCON I low-distortion mapping lens to provide mapping photography for the determination of geometric distortions in aerial mapping from advanced vehicles; the installation of a KS-25 camera with a 24-inch high resolution lens in X-15 #1 to determine whether extreme flight environments could significantly limit the optical performance of high-resolution camera systems; and two experiments, using X-15 #2, designed to provide data for the evaluation of

special color films in high-altitude photography. In addition to the cameras, special instrumentation was employed to monitor the mechanical and o"ptical performance of equipment and the conditions present during a test flight. On at least one occasion an unidentified high resolution optical sensor possibly a digitized optical transmission system of the type used in the Lockheed SR-71 and/or various types of reconnaissance satellite, was tested aboard the X-15. Because of the aircraft's unique performance envelope, it was found to be ideal for such test work. The X-15 program, by today's standards, was an exceptional bargain, financially speaking. Total program costs, including design, development, flight testing, and miscellaneous expenses amounted to just over $300-million.

SPECIAL DESIGN STUDIES: There were a number of advanced X-15 configurations proposed by North American during the program's lengthy history. Among these was a delta-wing modification calling for a maximum speed of between Mach 6 and Mach 8. The delta wing was considered to be the next logical step in the X-15 program. To this end, North American spent four years researching the delta X-15 configuration. During this time, some 300 hours of wind tunnel test data were gathered by using a 1/15th scale model for low-speed research (Mach 5) and a 1/5Oth scale model for high-speed (Mach 8) research. The wind tunnel tests covered both heating and aerodynamic loading, as well as the integrity, of such components as tanks and wing sections made of new super alloys. It was eventually concluded that the wing would use columbium alloy in the leading edges and Rene 41 steel on the surfaces. North American's research showed that X-15 #3 could be lengthened from 50 to 62.5 feet for the delta wing modification. Its wingspan would remain the same at 22 feet, but the stub wing would be replaced with a sharply raked delta wing sweeping back at an angle of 76°. North American also determined that lengthening the aircraft's nose into a more slender, bullet shape and increasing the aircraft's weight by 21,500 Ibs. (thus permitting a gross weight of 55,000 Ibs.) would be possible. These maximum limits were imposed to permit continued launching of the vehicle from the B-52 carrier aircraft without modifications to the B-52. An improved XLR99 was to be developed permitting slightly higher thrust at altitude and an increased burn time of 150 seconds. Additional propellants would be carried in a single external fuel tank. Benefits to be expected from such a delta-wing X-15 program were as follows: (1) Aerodynamic research-The flight tests would provide realistic aerodynamic data under fully developed turbulent flow conditions to supplement ground-based research where such conditions could not be achieved. Answers would be obtained to key questions relating to hypersonic aerodynamics of delta wings, large-scale behavior of flap-type controls, tip-fin interference effects, and handling qualities of a configuration typical of thenpresent thinking for a future hypersonic air breathing vehicle. Aerodynamic research on this vehicle would be unclouded by propulsion effects, inasmuch as most of the data would be taken under gliding flight conditions. (2) Structural research-The delta wing proposal would permit the evaluation in a practical flight application of a hot radiation-cooled structure designed for repeated flight at temperatures between 1500° F and 2200° F. It also would focus technical effort on a refurbishable, hotwing leading-edge design. In general, a delta-wing X-15 program could establish a baseline of confidence and technology from which decisions regarding the feasibility and design of advanced air-breathing vehicles could be realistically made. Additional advanced design studies included an X-15 to be used to launch Scout satellite launch vehicles; an X-15 study, conducted during November, 1964, at the Jet Propu lsion Laboratory under the auspices of the NASA Flight Research Center, calling for the use of canards to permit higher angles of attack during reentry; and an orbital X-15 that would be launched by a cluster booster of four Martin Titan liquid fuel heavy lift rockets-early schemes outlining an orbital X-15 program called for the use of a North American Navaho booster and an expendable X-15 (the pilot would have parachuted back to earth following the reentry portion of the mission). Recently coming to light is an illustration of a highly

modified X-15-type aircraft powered by what appear to be two ventrally-mounted scramjets. Little has surfaced concerning this study, but it appears to utiIize portions of the X-15 nose, canopy shell, and windscreen configuration. In almost all other respects it appears to be a totally new design.

POWERPLANT: During 1954 and early 1955, Reaction Motors, Inc. (soon to become the Reaction Motors Division of Thiokol Chemical Corp.) responded to inquiries by a number of airframe contractors calling for a rocket engine to propel the proposed X-15 high-speed, high-altitude research aircraft. A large number of configuration studies were undertaken in response to these inquiries, and a final design, based on Reaction Motors earlier XLR30, was eventually submitted to North American and the other airframe contract contenders. Following the award of the airframe contract to North American, the Air Force, which had jurisdiction over the project, decided to contract for the engine directly with the manufacturer. A formal bid request for the engine was received during December, 1955. As it were, the technical requirements were virtually identical to those in the earlier proposal to North American. The request did not leave any doubt that the Air Force wanted a rocket engine that would meet every facet of the safety requirement which stated that "any single malfunction in either engine or propulsion system should not create a condition which would be hazardous to the pilot". While it was possible to satisfy this requirement in general terms by improved reliability, the letter of the requirement could only be met by a rigorous design philosophy which would eventually leave its mark on virtually every detail of the propulsion package. Eventually, it was determined that the two most important requirements from a safety standpoint concerned the propellant combination, and the means of achieving combustion safety during starting and shut-down. Seven propellant combinations were explored in depth, these eventually being narrowed to liquid oxygen as the oxydizer and anhydrous ammonia as the fuel. The choice was based primarily on the fact that Reaction Motors had significant experience with liquid oxygen/ammonia propellant systems, and also on the fact that this propellant combination had much less critical starting characteristics. Additionally, the liquid oxygen/ammonia combination was an ideal coolant for the regenerative cooling of the proposed engine's thrust chamber. Another difficulty was unburned fuel accumulation in the thrust chamber. To solve this problem, a large, continuously operating igniter which would initiate the combustion process at low level, run continuously during operation, and over-run the main combustion during shutdown, was developed. Safety was further increased by the inclusion of a basic control sequence consisting of a series of successive steps using electrical interlocks. At each step, means would be provided for the electrical circuit to determine the status and operate a "decision function" to proceed to the next step or shut down safely. The basic XLR99 specification was as follows: Thrust-50,OOO Ibs. at sea level; 57,850 lbs. at 100,000 feet Throttling range-30% to 100% Engine specific impulse-276 lb.-sec./lb. at 100,000 feet Oxydizer/fuel ration-l.25 Chamber pressure-600 psi Duration-limited by propellant supply Dry weight-915 Ibs. Service Iife-1 hour (TBO) Early test versions of the XLR99 were designed to have thrust ranges of from 50% to 100%, but later versions were able to meet the 30% requirement. The throttling control of the engine was provided solely by varying the rotational speed of the turbopump, the main propellant valves being used only for shut-off purposes and the control of propellants in the priming cycle. Control gas for the engine pneumatic controls was provided by an aircraft helium supply at a regulated pressure of 550 psi; hydrogen peroxide for the turbine drive was similarly provided from a pressurized tank regulated to 600 psi. The gas in this tank was provided by decomposing monopropellant fuel, 90% hydrogen peroxide, in a gas generator. . The XLR99 operating sequence was as follows: (1) While the carrier aircraft and X-15 were climbing to launch altitude, the priming cycle was started, passing propellants through the pumps and then overboard to

bring the metal parts of the turbopump, lines, and valves down to full cryogenic conditions. The final priming consisted of approximately 75 seconds at relatively high flow rates. (2) Immediately before drop, an engine "idle" cycle was initiated for 5 to 30 seconds and consisted of a thrust chamber purge cycle followed by pump run-up to 6,000 rpm and the establishment of igniter operation. The pilot was able to monitor these functions during idle running and had a short period to determine from his condition lights in the aircraft that all was well before committing to flight. (3) Immediately after the X-15 dropped from the carrier aircraft, the main chamber firing was initiated by opening the main propellant valves. Ignition of the propellants in the main chamber was assured from the already operating igniter. The engine was now under the direct control of the pilot and could be throttled up, shut down, or restarted at will. (4) Manual shut-down was effected by closing the pilot's throttle or by switching off. The engine responded by closing the main propellant valves and automatically applying a helium gas purge. (5) Restarting operations could be initiated by a single switch or by opening the throttle without other control operations by the pilot. Purge gas was available in the aircraft system for five restarts in flight. (6) If the propellants were run to complete exhaustion, shut down was accomplished automatically by the onset of cavitation in the appropriate pump. Upon sensing cavitation, the engine control system shut down the main thrust chamber and proceeded through the purge cycle. Mechanically, the XLR99 was an exceptionally complex and sophisticated powerplant. The thrust chamber, which went through many configuration studies before final qualification, had a throat diameter of 8.64 inches and a nozzle area ratio of 9.8. The nozzle diameter was 39.3 inches. The chamber was regeneratively cooled by the fuel and was constructed by the well known tubebundle method. The cooling arrangement was of the "two pass" type in which the coolant was passed down one tube and returned in the adjacent tube. The tubes, of which there were 196 furnace brazed in the chamber bundle, were made of AISI 347 stainless steel. They were 3/8ths inch in outside diameter and they had a wall thickness of .033 inches. At the head end of the thrust chamber, inlet and outlet manifolds were formed from age-hardened Inconel-X alloy. They were attached by welding to the bundle head rings. Though various types of injectors were studied, the multi-spud type injector was chosen because of ease of construction and favorable ejector characteristics. The injector also was of age-hardened Inconel-X. The igniter, which was, in effect, a small rocket thrust chamber, provided two states of combustion to assure dependable ignition with maximum safety. In the first stage, the ignition of a mixture of gaseous oxygen and liquid ammonia was effected by the discharge from three surface-gap spark plugs. The products of combustion were discharged through a throat into the second stage combustion cavity where they ignited the second stage flow. In the second stage igniter, the propellants were introduced into the combustion cavity through an impinging jet injector with the fuel flow being utilized to regeneratively cool the second stage chamber. To maintain the mixture ratio within the flammability range, a cavitating venturi was located in each of the second stage propellant lines to limit flow under transient conditions. The turbopump was perhaps the most difficult part of the XLR99 to develop. During the early engine design stages, areas of uncertainty existed regarding the drive gas for the turbopump turbine and the basic hydraulic characteristics of the two pumps it was required to power. The turbine drive system eventually chosen was based on the use of 90% hydrogen peroxide in monopropellant decomposition using a silver screen catalyst bed. This generated gas at 1,360° F. upstream of the nozzles and lent itself to a most simple system of speed control. The gas impinged upon a turbine wheel which then transmitted its power to the two pumps via a single shaft. The turbine drive powered two separate pumps, one for fuel and one for oxydizer. The oxydizer pump operated at approximately 13,000 rpm and was based on the oxydizer pump developed by the Reaction Motors engineering team for the XLR30. The fuel system pump, mounted on the same shaft driving the oxydizer pump, operated at 20,790 rpm. Outlet pressures for these units were in the order of 1,200 psi and the horsepower required to power them was nearly 1,500. The combined oxydizer/

27

fuel flow rate at maximum thrust was 13,000 Ibs. per minute. At maximum thrust the standard X-15's 18,000 lb. propellant supply was exhausted in 85 seconds. Static testing of prototype XLR99's and associated assemblies and systems took place at Reaction Motors' Lake Denmark, New Jersey facility. There, four test stands were utilized, inclUding the Picatinny Arsenal Stand El. The largest stand, R2, was set up to test a complete aircraft system in all attitudes. Stands R2W and R3 were capable of horizontal firing only. The former was utilized for durability testing and environmental testing and the laller for delivery acceptance tests (it was equipped with an elaborate thrust vector mount). The entire test area was provided with self-contained support facilities with a storage capacity of 30,000 gallons of liquid oxygen, 18,000 gallons of liqUid ammonia, and 4,000 gallons of hydrogen peroxide. Actual engine system testing was initiated during the fall of 1958. Beginning in 1959, eight engines of flightworthy design were constructed for development purposes. From this phase, engine running time progressed consistently and, in all, some 340 minutes of engine operation were accumulated. Much greater running times were logged on various sub-systems with total thrust chamber time being in the order of 1,800 minutes and turbopump running time being in the order of 4,200 minutes. Four engine assemblies were used in the performance of the Preliminary Flight Rating Tests (PFRn. Based on the specification MIL-E-6626, the completion of this series formed the basis for approval of the engine for use in experimental aircraft. Engine durability was later demonstrated using two engines, each accumulating more than 60 minutes of operating time. One engine was fired 108 times without having any more than routine maintenance checks. A series of 93 tests was made to demonstrate that the engine would react safely under imposed malfunction conditions. In all, 234 engine tests were made to demonstrate the operational ratings and the achievement of manned vehicle safety. Of these, 192 were full engine firing demonstrations and the remaining 42 were safety limit tests in which thrust chamber operation was not necessary. The entire PFRT program was accomplished during a period of 4 months without a major incident. Shortly after the completion of the tests during December, 1960, official approval of the engine for flight was received from the USAF. Due to its complexity and numerous development problems, the XLR99 was many months late in being delivered to North American for installation in an X-15. In order to get flight testing underway, the various X-15 operating agencies elected to have North American complete the first two aircraft with interim Reaction Motors,

Inc. XLRll (RMI 6000 C4) rocket engines. Two XLR11's were installed in each of the first two X-15's, each engine having four thrust chambers and capable of producing a total of 8,000 Ibs. thrust. These engines were quite familiar to personnel working in the experimental rocket aircraft programs at Edwards as they were the same powerplant type used in the original Bell X-l's, the Douglas D558-II's, and the Republic XF-91. The basic XLRll configuration was as follows: Construction-four thrust chambers producing 2,000 Ibs. thrust each with a turbopump unit, valves, regUlators, and controls mounted forward of the chambers; the four chambers were mounted about a control support beam assembly which was the main structural member of the engine; welded stainless steel construction and aluminum tubing were used throughout. Propellants-ethyl alcohol-water for fuel and liquid oxygen for the oxydizer; both were injected under pressure supplied by a turbopump; valves in the fuel and oxydizer lines controlled the flow of the propellants to the chambers. Ignition-each thrust chamber contained an igniter which fired into it to initiate combustion; the chambers could be fired individually in any sequence. Cooling-regenerative; fuel, before injection into a cylinder, was circulated through coolant passages in each exhaust nozzle and around each combustion chamber individually. Mounting-the engine mount was conventional with four mounting points on a bracket. Weight-dry (including pump) was 345 Ibs.; the thrust to weight ratio was 17 Ibs. per lb. of thrust. As noted previously, the XLR99 was the end product of a lengthy Reaction Motors powerplant development program that had been on-going for nearly ten years. Reaction Motors (which, during the course of XLR99 development, became a division of Bristol, Pennsylvaniabased Thiokol Chemical Corporation) had been one of the four contenders for the X-15 powerplant contract when invitations to bid were opened on February 4, 1955. Of the four proposals submitted, Reaction Motors' design was determined to be the best, and most importantly, the safest. The Reaction Motors contract, AF33(600)-32248, signed on September 7, 1956, called for the first flightrated engine to be ready for installation two years later. As is often the case, however, the powerplant lagged behind the airframe by a considerable margin, and thus the XLR11's were used in the X-15 during its first powered flights. Because of the high thrust levels of the XLR99, the noise levels in the immediate vicinity of the X-15 were severe. A maximum of 163 decibels was noted 25 feet behind the aircraft. It was necessary to design the vehicle

structure to withstand these high sound pressure levels, such as the 156 decibels at the horizontal stabilator position.

XLR99 DATA Continuous Operating Limit Limited by airframe tankage (est 87 sec at max thrust: 156 sec at min thrust) Rated thr Chamber Pressure 335-600 psia - 40 to 160°F Operational Temp Range. -40 to - - - OF Storage Temp Range. . Pressure Feed. . . . . . . . . . . . . . . .. . None . . . . . . . . . . . . . •. . Centrifugal Pump Feed. Pump Drive. ..... ................. . ... Turbine Chamber Arrangement. . . . . . . .. Single chamber Variable Thr Control. . . . . . • . . .. Variable 25.000 to 50,000 Ib thr at SL Utilizes 90% H,O, Gas Generator. Cooling: ................. Regen with fuel Thr Chamber. ............... None Gas Generator. .2-stage combustion device Ignition. Automatic controls: Mixture Ratio. . . . . . . ..... Fixed orifices Pump Speed Variable, controlled by throttle setting Starting Pressurized H,O, feed to gas generator Fuel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anhydrous ammonia Oxidizer. . .. Liquid oxygen Gas Generator Propellant. . 90% H,O, Required Pump Inlet Pressures: 37.2 psia Fuel (- 28°F at SL) . .50.0 psia Oxidizer (- 285°F at SL) . . None Accessory Drive Provisions .

AVAILABLE SCALE MODELS AND DECALS: Kits Aurora (NASA): 1/48th Aurora (USAF): 1I48th Busch: 1I175th Eldon: 1/98th Heller: 1/65th Hobby Miniatures: 1/175th Marusan: 1I98th Mlyauchl: 1/70th Revell: 1I65th Revell (B-521X-15 comblnstlon): 1I175th Sankol: 1I125th SanwalTokyo Plamo: 1/100th Starflx: 1I111th UPC: 1/98th ZA: 1I70th Decals Other than decals known to be provided with the aforementioned kits, there are none available.

SPECIFICATIONS AND PERFORMANCE: Fuselage length (w/XLRll; no pitot) length (w/XLR99) height (skid to vertical fin tip) maximum diameter maximum width (chine to chine) ground line angle at rest Wing span root chord (at chine intersect) tip chord loading area (total, including 94.98 sq.' blanketed by fuselage and side fairings) aspect ratio sweepback (25% element) airfoil (root and tip) incidence dihedral flap area flap travel Horizontal stabllator span area (exposed) aspect ratio root chord (at chine intersect) tip chord sweepback (25% element) airfoil (root and tip)

anhedral travel (at leading edge)

28

i

X-15

X-15A-2

49'10" 50'3" 139.6" 56" 88.50" 2° 21'

n.8. 52'5" 145.6" 56" 88.50" 2° 21'

22'4" 10'3" 35.78" 166.5 Ibs.lsq.'

22'4" 10'3" 35.78" 280.65 Ibs.lsq.'

200 sq.' 2.5 25° 38' 27.6" 66005 (Mod) 1% blunt TE 0° 0° 15.48 sq.' 30° down

200 sq.' 2.5 25° 38' 27.6" 66005 (Mod) 1% blunt TE 0° 0° 15.48 sq.' 30° down

216.93" 51.76 sq.'. 2.83 60.07" 25.28" 45° 66005 (Mod) 1% biunt TE 15° 15° up 35° down

216.93" 51.76sq.' 2.83 60.07" 25.28" 45° 66005 (Mod) 1% blunt TE 15° 15° up 35° down

Vertical stabllator (dorsal) height above fuselage area (exposed) aspect ratio chord (at fuselage intersect) tip chord sweepback (25% eiement) airfoil movable tail area (slab) travel Vertical stabllator (ventral) ventral dimension below fuselage tip chord area (exposed) jettison area jettison trim line (below fuselage) fixed area remaining after jettison travel Speed brakes area (~pper surface each side) area (lower surface each side) travel (measured from centerline of aircraft) travel (measured from edge of stabilator) Landing gear skid track nose gear to skid tread Empty weight Launch weight Maximum altitude Maximum speed Range

X·15

X-15A-2

55" 40.9 sq.' 0.516 107.4" 90.75" 23° 24" 10° wedge 26.5 sq.' 7.5° left 7.5° right

55" 40.9 sq.' 0.516 107.4" 90.75" 23° 24" 10° wedge 26.5 sq.' 7.5° left 7.5° right

44" 96" 34.4 sq.' 19.9 sq.' 17.5" 14.5 sq.' 7.5° left 7.5° right

44" 96" 34.4 sq.' 19.9 sq.' 17.5" 14.5 sq.' 7.5° left 7.5° right

5.6 sq.' 5.6 sq.'

5.6 sq.' 5.6 sq.'

41° 36°

41° 36°

115.85" 476.01" 11,374Ibs. 31,275Ibs. 314,750' 4.104 mph 275 miles avg.

115.85" 502.01" 18,340Ibs. 56,130Ibs. 354,200' 4,534 mph 275 miles avg.

X-15, 56·6670

X-15, 56-6670, following upload. Photo calibration markings are visible on the left slab stabilator and a protective cover is visible on the left wing. Noteworthy is the missing lower half of the ventral tail surface.

X-15, 56-6670, is seen following uploading aboard NB-52A, 52-003. Noteworthy is the NB-52A's black painted wing root section leading edge which was done to attenuate sun glare. The boarding ladder was the standard cockpit entry device. Q

X-15, 56-6670, being dropped from NB-52A, 52-003 on May 12, 1960, with Joe Walker as pilot. Condensation from liquid oxygen can be seen streaming behind the aircraft. This ffight resulted in the first Mach 3 ffight of the X-15 program.

X-15, 56-6670, at the moment of flare during landing. Fairly shallow angle of attack is apparent. Noteworthy is the extreme rearward placement of the main gear skids and the nominal displacement of the slab stabilators for pitch control.

!ttl. Post-landing activity surrounds X-15, 56-6670, in 1959. Visible are the extended airbrakes, the raised canopy, and some servicing equipment details. Transport dolly facilitated towing the aircraft back to NASA's facility for refurbishment.

Post-landing view of X-15, 56-6670, in 1961, following modification to "ball" nose configuration. As shown, the "ball" nose had been covered with a special plastic cap for protection. Static X-15's low stance is particularly noticeable in this view.

X-15, 56-6670, on Rogers Dry Lake immediately following test ffight. Visible underneath fuselage is liquid oxygen tank external condensation. Single or double support rod could be used to prop open the canopy.

Late in its career, X-15, 56-6670, is seen on the dry lakebed at Edwards. USAF/NASA markings are noteworthy and symbolic of the close cooperation maintained on the project between the two main sponsoring government entities.

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Pre-flight servicing of the X-15, as seen in this view of 56-6671, was a complicated and somewhat dangerous affair. The various propellant and systems tanks were usually filled prior to takeoff, and then topped-off immediately prior to launch.

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While attached to the pylon, the X-15 was almost totally dependent upon B-52 systems for support. X-15, 56-6671 is seen from the forward NB-52B "bug eye" shortly before being launched. Distinctive upward flex of NB-52B's wing is noteworthy.

Moments after launch, X-15, 56-6671, begins powerplant ignition procedures. The lower half of the ventral tail surface is attached in this view, and the aircraft has been equipped with a "ball" nose and an XLR99 powerplant.

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Post-flight view of X-15, 56-6671, as it is towed by tractor back to the NASA facility at Edwards AFB following the first powered flight of the X-15 program on September 17, 1959. Visible are the exhaust nozzles of the twin-pack XLRll propulsion unit.

Post-release view of X-15, 56-6671, as powered by the twin-pack XLRll propulsion unit. Visible are seven ignited chambers. The eighth chamber has not yet been activated. Also visible is the external skin condensation from the liquid oxygen tank.

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X-15, 56-6671, flares for landing at Edwards AFB. Condensation from the liquid oxygen tank is visible underneath the fuselage. Extended nose gear strut is noteworthy. Lower half of ventral tail surface has been jettisoned.

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X-t5, 56-6671, after its August 29, 1962, flight with AF pilot Robert Rushworth at the controls. The extended nose landing gear is uncharacteristic of the aircraft following landing. Note, too, the plastic nose cap placed over the "ball" nose.

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':r '.", :!, .' Following its arrival at Edwards AFB, X-15, 56-6672, is run through a weight and balance check on May 19, 1960, while suspended from a crane-supported sling. Calibration was done manually, using finely calibrated field surveyor's levels.

With its XLR99 covered e,nd plugs placed in its anyhdrous ammonia and LOX dump tubes, X-15, 56-6672, is prepared for upload and attachment to the launch pylon on July 9, 1962. This mission, aborted three times, was finally completed on July 17,

X-15, 56-6672, with AF pilot Robert White at the controls, is seen departing Edwards AFB on July 17, 1962, under the wing of NB-52A, 52-003, on what was to become the first flight wherein an X-15 pilot earned the right to astronaut's wings.

Filled with propellants, X-15, 56-6672, is seen departing Edwards AFB under the wing of the NB-52A, on October 27, 1965. Not visible is an experimental sharp leading edge attached to the upper vertical tail surface for the exploration of boundary layer noise. I~

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Just prior to launch, X-15, 56-6672 is seen dumping LOX and anhydrous ammonia during the top-off process. Also visible on the nose is "tempfac" paint-which is temperature sensitive and can provide post-flight temperature data based on its color

Equipped with podded wingtip sensors, tail mounted radiometers, and a vertical tail surface with a special sharp leading edge, X-15, 56-6672, is seen being carried to launch altitude under the wing of the NB-52A on September 14, 1966.

X-15, 56-6672, at Edwards AFB during July, 1962. This aircraft was ill-fated and, some five years after this photo was taken, would be involved in the only fatal accident of the X-15 program. The exact accident cause was never fully determined.

X-15, 56-6672, was the only one of the three X·15's to be equipped with the MH-96 adaptive control system. This unit automatically transitioned control of the aircraft from aerodynamic to ballistic systems and back again, as the envelope required.

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Following its emergency taneJing attempt on November 9, 1962, at Mud Lake, Nevada, with NASA pilot John McKay at the controls, X-15, 56-6671, was transported back to North American's Los Angeles facility for rebuilding into the X-15A-2.

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Following completion of the modification effort that included an extended fuselage, a revised windscreen, propellant drop tanks, and modified landing gear, X-15A-2, 56-6671, was rolled·out and prepared for transport back to Edwards AFB.

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The first flight of the X-15A-2 with its jettisonable propellant tanks attached took place on November 3, 1965. For the flight, the tanks were left empty; they were successfully jettisoned as the X-15A-2 accelerated through Mach 2.

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Though often published, this photo has rarely been identified as depicting the launch of the X-15A-2 at the beginning of its October 3, 1967, flight wherein an unofficial world's absolute speed record of 4,520 mph was set by AF pilot Pete Knight.

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The first X-15, 56-6670, immediately prior to the official roll-out ceremonies held at the company's Los Angeles facility on October 15, 1958. Markings consisted of an over-all blue-black paint scheme and high-visibility white reference and identification markings. All but the over-all blue-black paint would change dramatically during the course of the aircraft's flight test program-with markings almost never exactly the same from flight to flight.

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each of the NB-52's was painted somewhat differently in defference to identification requirements. NB-52A, 52-003, seen on April 13, 1960, had a bright orange vertical fin with distinctive, white-trimmed, red scallop-like markings on the engine nacelles and red sections on the nose and outer wing panels. Distinctive also was the aircraft nose radome, which was painted flat black. The national insigne and USAF on the fuselage sides and wings were standard for type.

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NB-52B, 52-008, on September 20, 1960 (an aborted flight), with X-15, 56-6670. The NB-52B's markings were similar to those of the NB-52A, but eliminated the orange vertical fin while adding a wraparound orange band on the empennage section. Noteworthy are the white wing tips and trailing edge markings, and the white bomb bay doors. Barely visible on the fuselage side are the first of what were to become many X-15 launch/abort symbols (painted black with red exhaust plumes).

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SELECT MARKINGS

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The X-15A-2, 56-6671, was originally flown with a full ventral fin assembly, including the lower, "rudder" ponion. The ventral later would be highly modified to accommodate the attachment and aerodynamic requirements of the scramjet engine. The over-all scheme of the aircraft was semi-gloss blue/black with a large, yellow-orange horizontal NASA bar on the venical tail surface assembly. A white bar was. painted on the wing undersurfaces and stretched from the root to the wing tip at about two-thirds the wing chord. The external tanks had da-glow orange noses, black photo reference marks, and gray main bodies. The left external tank had a white aft body section and a black end cover. The ventral tail surface (but not the rudder portion) was covered in a dark orange ablative.

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Initial X-15A-2 flights with a dummy scramjet in place were conducted without the external propellant tanks. The primary physical difference at this point was the highly modified ventral fin which had a totally reconfigured leading edge and numerous internal changes. The various scramjets tested were usually painted white over-all with a black horizontal photo reference marking. The steel helium tank which protruded from the companment immediately aft of the venical tail was left unpainted. All other markings for the X-15 were standard.

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When the first X-15, 56-6670, was rolled out at North American's Los Angeles facility on October 15, 1958, it was painted in semi·gloss· blue/black paint over-all and given a varied selection of miscellaneous warning markings in gloss white paint. The nose boom was painted in white and black stripes (later changed to white and red) and the rescue logo just below the windscreen was in yellow with black letters. Only the national insigne was colorful, being standard red and white. The nose landing gear wheel was painted silver. The USAF serial number on the vertical fin, and the "U.S. Air Force" on the fuselage were painted in semi-gloss white.

The X-15A-2, 56-6671, was originally flown with a full ventral fin assembly, including the lower, "rudder" portion. The ventral later would be highly modified to accommodate the attachment and aerodynamic requirements of the scramjet engine. The over-all scheme of the aircraft was semi-gloss blue/black with a large, yellow-orange horizontal NASA bar on the vertical tail surface assembly. A white bar was. painted on the wing undersurfaces and stretched from the root to the wing tip at about two-thirds the wing chord. The external tanks had da-glow orange noses, black photo reference marks, and gray main bodies. The left external tank had a white aft body section and a black end cover. The ventral tail surface (but not the rudder portion) was covered in a dark orange ablative.

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Initial X-15A-2 flights were made with the lower part of the ventral off, and without the external propellant tanks in place. The markings remained essentially unchanged, with the aircraft being painted semi-gloss blue/black over-all, and the vertical tail surface carrying the large NASA logo on a yellow·orange horizontal bar.

Initial X-15A-2 flights with a dummy scramjet in place were conducted without the external propellant tanks. The primary physical difference at this point was the highly modified ventral fin which had a totally reconfigured leading edge and numerous internal changes. The various scramjets tested were usually painted white over-all with a black horizontal photo reference marking. The steel helium tank which protruded from the compartment immediately aft of the vertical tail was left unpainted. All other markings for the X-15 were standard.

With its MA-25S ablative in place, the X-15A-2 gave the appearance of having weathered a rather serious snow storm. The aircraft was quite literally white over-all, with only the "ball" nose left unpainted. Various warning markings were visible on the nose and empennage sections, but all national insigne and related items were covered over. Even the windscreen was given an ablative covering, this dictating that the pilot fly almost the entire mission on instruments.

Scale: 1/72nd _____."iII Drawn By Douglas Siowiak 35

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.... When the first X-IS, 56-6670, was rolled out at North American's Los Angeles facility on October IS, 1958, it was painted in semi-gloss' blue/black paint over-all and given a varied selection of miscellaneous warning markings in gloss white paint. The nose boom was painted in white and black stripes (later changed to white and red) and the rescue logo just below the windscreen was in yellow with black letters. Only the national insigne was colorful, being standard red and white. The nose landing gear wheel was painted silver. The USAF serial number on the vertical fin, and the "U.S. Air Force" on the fuselage were painted in semi-gloss white.

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.... Initial X-15A-2 flights were made with the lower part of the ventrai off, and without the external propellant tanks in place. The markings remained essentially unchanged, with the aircraft being painted semi-gloss blue/black over-all, and the vertical tail surface carrying the large NASA logo on a yellow-orange horizontal bar.

.... With its MA-25S ablative in place, the X-15A-2 gave the appearance of having weathered a rather serious snow storm. The aircraft was quite literally white over-all, with only the "ball" nose left unpainted. Various warning markings were visible on the nose and empennage sections, but all national insigne and related items were covered over. Even the windscreen was given an ablative covering, this dictating that the pilot fly almost the entire mission on instruments.

Scale: 1/72nd ........----. .. Drawn By Douglas Siowiak 5

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NORTH AMERICAN X-iS, 56-6671

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The second X-IS, 56-6671, as it was painted during Robert White's November 9, 1961, flight in which an unofficial world absolute speed record of 4,093 mph was set. In order to attenuate friction generated heat, the canopy/hatch area was given a coating of white, heat-resistant paint and several fuselage side panels were similarly covered. All other markings remained essentially unchanged, with systems markings in white, most warning labels in yellow, and the NASA logo in yellow/orange visible on a large horizontal bar on the vertical tail surface. For this mission, the aircraft was flown with the complete ventral tail surface set in place.

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The second X-IS, 56-6671, also was flown with a large nose test boom in place (painted in red and white stripes), prior to having this removed and replaced by a "ball" nose. The aircraft was painted blue/black over-all and except for the serial number on the vertical tail surface, was virtually identical from a markings standpoint to 56-6670 in all respects during the early stages of the X-IS flight test program.

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The second X-15, 56-6671, as It was painted during Robert White's November 9, 1961, flight in which an unofficial world absolute speed record of 4,093 mph was set. In order to attenuate fnction generated heat, the canopy/hatch area was gIVen a coatmg of white, heat-resistant paint and several fuselage side panels were similarly covered. All other markings remained essentially unchanged, with systems markings in white, most warning labels in yellow, and the NASA logo in yellow/orange Visible on a large horizontal bar on the vertical tail surface. For this miSSion, the alfcraft was flown with the complete ventral tall surface set m place.

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The second X-IS, 56-6671, also was flown With a large nose test boom in place (painted in red and white stripes), prior to having this removed and replaced by a "ball" nose. The aircraft was painted blue/black over-all and except for the serial number on the vertical tall surface, was virtually Identical from a markings standpoint to 56-6670 in all respects during the early stages of the X-15 flight test program.

X-15A-2 CONFIGURATION CHANGES ......- - - - - - .

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Late in the X·1S program the NB-S2B was given a new coat of paint and significantly revised markings of a less conspicuous nature. Noteworlhy were the off-white to light gray engine nacelles, the white wing tips, and the reflective white paneling on the top of the fuselage to the rear of the cockpit area. Barely discernible on the right fuselage side was the X-IS launch/aborl symbols panel. This parlicular photo was taken on August 7, 1967.

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The X-1SA-2, 56-6671, seen on November 2, 1965, in preparation for its first flight with its external propellant tank system in place, was to become the most colorful of the three X-IS's thanks to its varied markings and (later) coating of MA-2SS ablative. Visible in this view is the open Q-bay-like sensor bay, foeated just behind the cockpit area. This served to accommodate various test instruments and research tools, such as cameras. Tank ground clearance is notewOrlhy.

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Cockpit of X-15A-2, 56-6671, as it looks today while on display at the USAF Museum, Wright-Patterson AFB, Ohio. The cockpit was purposefully painted in subdued gray. Flight attitude and performance instrumentation dominated center panel.

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North American developed a dedicated ejection seat for the X-15. The X-15A-2, as displayed at the USAF Museum, still contains an original seat though one that was never utilized in the actual flight test program. Seat cushions were pilot customized.

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View of the full-scale X-IS simulator cockpit depicting it as it appeared in aircraft 56-6670 upon roll-out. Of particular note is the fact that the instrument panel board is painted black and divided by white lines that segregated flight instrumentation from propulsion and other instrumentation. This helped pilots prioritize panel references without unduly burdening the decision making process. All three control sticks are visible in this view.

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X-15, 56-6672, was the only one of the three X-15's to be equipped with the MH-96 adaptive control system. This unit dictated significant changes in the design of the main instrument panel. In this view, the MH-96 display system is indicated by the three main centrally-mounted analogue indicators. The vertical display on the right was associated with yaw, while that on the left was associated with angle of attack. The MH-96 was highly automated, requiring little pilot input.

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The X-15A-2's instrument panel was essentially unchanged from that of the original X-15 configuration. Noteworthy in this view, showing the aircraft as it now appears at the AF Museum, is the relatively uncluttered panel layout, the obvious segregation of propulsion instrumentation from flight instrumentation, and the missing Mach meter. Also noteworthy is the total lack of digital instrumentation; only analogue was available at the time of the X-15's flight test program.

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The original X-IS panel layout was typical of research aircraft of late 1950's vintage. All instrumentation was analogue, and the prevailing philosophy for instrument panel coloring was black-which was assumed to provide better contrast and thus improved legibility. The panel was essentially divided into four instrument groups including (from left to right) propulsion, flight, secondary propulsion/electrical, and environmental (temperature/pressurization for cockpit, instrument bay, etc.).

The instrument panel and associated instrumentation were mounted under a titanium cowl that was rigidly attached to the main airframe and ventilated by a large' number of holes. With the canopy closed, the panel protruded into the cockpit area and over the pilot's knees and legs. For ejection to be accomplished without injury, it was necessary for an automatic system to retract the pilot's appendages to provide the necessary clearances.

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The left side console of the X-15A-2 served as the mounting point for the throttle quadrant, the radio communication system, the upper and lower air brake actuation handle, the internal propellant tanks jettison handle, and the external propellant tanks jettison handle. Visible, but with its grip removed, is the ballistic control system control handle, which is seen protrudin.a from· the rectanaular hole above the left console.

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The left side console of X-IS, 56·6670, illustrating its similarity to the same panel in the X-15A-2. Noteworthy in this view, however, is the complete grip assembly for the ballistic control system which activated the reaction control rockets in the nose and wings. The intercom control panel box is the black rectangle mounted on the inner aircraft wall. Visible also are the throttle quadrant, the air brake lever, and one of the retracted ejection seat hand grips.

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The right side console of the X-15a-2 served almost solely as a circuit breaker panel. Protruding downward from the canopy combing was the canopy release lever, and protruding from the venical sub-panel on the left was a T-handle that was for emergency canopy release. Visible too, is one of two ejection seat hand grips. This device served not only to initiate seat ejection, but also to restrain arm and hand movement as the seat moved upward into the free airstream.

There were few differences between the X-15A-2 left side console and that of the standard X-IS. X-IS, 55·6670's left side console was, like that of its sister ships, almost totally dedicated to a large circuit breaker panel. This was placed so that it was easily accessed by the pilot, who was usually highly restrained in the X-IS cockpit and ejection seat-and subject to severe movement restrictions. The T-handle visible on the ejection seat arm was an emergency harness release.

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The X-15A-2's canopy was modified during conversion to accommodate a new elliptical windscreen configuration. This design was optimized to withstand the temperatures generated by the X-15A-2's improved performance capabilities:

The X-15A-2's ablator "outgassed" during reentry and thus covered the windscreen with a transluscent coating. To overcome this difficulty, a mechanical eyelid was developed and attached to the left windscreen.

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X-15, 56-6670's ejection seat headrest and inside canopy details. To prevent glare difficulties, the inside of the windscreen area was painted black, whereas the rest of the canopy remained off-white, or gray. Various sensors were attached to the canopy.

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Inside canopy details of the X-15A-2. Like that for X-15, 56-6670, the inside of the windscreen area was painted black to prevent glare difficulties. Noteworthy is the pilot upper helmet restraint built into the canopy/windscreen.

An anthropomorpntc dummy sits in a test sample of North American's dedicated X-15 ejection seat. Noteworthy are the deployed stabilization wings.

A test-sled mounted X-15 ejection seat is seen shortly before launch. The seat was optimized for safe ejection at speeds as slow as 80 knots or as high as Mach 4. Unlike contemporary ejection seats, this unit had a pronounced robust image that has never been duplicated.

A large number of sled tests were undertaken to verify the X-15 ejection seat's capabilities and performance envelope. As can be seen, during 1957 a mock-up X-15 forward nose section was built which provided a full-scale aerodynamic environment to test both canopy separation and ejection seat effectivity. This rail-mounted unit could be used to test the seat out to speeds of approximately Mach 1. Photo markings provided reference data for camera generated imagery.

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The ejection seat also was tested for aerodynamic stability by mounting it on a rocketpropelled sled. As can be seen, large retractable wings for seat stabilization were deployed from each side of the seat after it exited the X-15 cockpit.

The X-15 forWard nose section rocket sled is seen during a 1957 test at Edwards AFB in which a full-scale development X-15 seat was ejected at high subsonic speeds. As can be seen, canopy clearance was substantial by the time the seat exited.

47

Two views showing the North American ejection seat installed in X-15, 56-6670 (left), and the X-15A-2. As can be seen, the seat installation maintained tight tolerances inside the X-15's already cramped cockpit. Restraining lanyards were built into the seat and attached directly to the pilot's suit. Each pilot had his own buttock pad fitted for comfort. This could be changed to accommodate each pilot. Of note is the neat fit of the retractable stabilization wings inside the X-15's canopy. _.,3"

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The North American-developed ejection seat was equipped with extendable stabilization arms as well as the stabilization wings. The arms were mounted underneath the seat and resembled a large telescoping automobile radio antenna.

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Following initial flight test work with a less accurate system, all three X-15's were modilied to incorporate an inertial guidance unit that originally had been developed for the still-born Boeing X-20A "Oyna-Soar".

The X-15 was equipped with a ballistic control system consisting of eight nose-mounted and four wing-mounted reaction control rockets. Once the X-15 was in a ballistic mode the RCS could be utilized to orient the aircraft for the proper reentry angle.

All three X-15's were litted with the Northrop-developed "ball" nose following initial flight trials with a conventional nose boom. The "ball" nose, contrary to some reports, was only partially articulated and was not a complete ball.

This view of the "ball" nose as installed on X-15, 56-6670, also provides details of the pitch and yaw controlling exhaust ports for the ballistic control system RCS rockets. Two exhaust ports, somewhat offset, were provided for each quadrant.

,., Localized airflow tests were sometimes accomplished by installing a rake, such as this one seen under the nose of X-15, 56-6670, immediately prior to a flight on May 22, 1962. Noteworthy are the exhaust ports for the RCS rockets.

The slight offset of the reaction control system nozzles is particufarly noticeable in this view of the nose of the X-15A-2. The offset was dictated by internal space constraints which forced the rocket chambers to be offset in pairs.

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A fiberglass nose cover often was utilized to protect the X-15's "ball" nose from damage while the aircraft was on the ground. Noteworthy is "templac" paint which is utilized to visually register temperature highs through color change.

GENERAL ARRANGEMENT

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BALLISTIC CONTROL SYSTEM Front view of X-15A-2 provides details of ablative coating application, elliptical and eyelid-equipped windscreens and access panels for ReS. Noteworthy are protected "ball" nose, exposed APU exhaust nozzles, and paneling details.

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A static pitot type sensor (right) and an extendable probe are seen on the nose of the X-15A-2. The extendable probe, which represents just one of many that were eventually

used during various X-15 research efforts, is seen in its extended position. These probes were envelope sensitive and were designed to sample atmospheric and aerodynamic conditions only at certain times. They were mechanically actuated and tended to be exposed for only short periods during a flight.

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The nose landing gear well door was itself equipped with a smaller door that was, in fact, a small airbrake. This unit popped down into the slipstream and created the aerodynamic drag required to pull the nose gear down and into its locked position.

The free-castoring nose landing gear was kept relatively simple in terms of its mechanical design and had only one hinge point. It also was provided with a long travel to accommodate the high landing loads it was expected to encounter.

The X-15A-2 nose gear assembly was a robust unit designed to accommodate the exceptional momentary loads generated during an X-15 landing. The X-15's unique gear placement and type increased these loads considerably.

The nose gear well for the X-15A-2 was heavily insulated to protect the tire and nose gear assembly from friction generated heat. Nose gear extension depended on aerodynamic drag, and gravity.

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To accommodate the X-IS, the ground transport dolly could be lowered by relieving the pressure in each of its two hydraulic jacks. The wheels and tires, which were mounted on articulated parallel arms, would then allow the main fuselage jacks to be moved into place and locked into special fuselage sockets. Once this was done, the hydraulic jacks were then pumped up and the aircraft raised off its skids. A tow bar attached to the nose landing gear accomplished the rest.

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The skid main gear was manually ground-retractable. It was extended by the effects of gravity and aerodynamic download. Each skid shoe was equipped with a separate shock absorber for stabilization, and the entire strut assembly was stabilized by a two-part foldable drag link. The gear strut was part of the fairing assembly when the gear was retracted, and the entire unit remained exposed during flight.

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The main gear skid was of titanium construction, similar to that of other major X-IS assemblies. It was a boat-shaped unit with a small hinge point inside its hull to which was attached the main gear strut. Shock absorption was minimal, this being accommodated by an internally mounted energy absorption assembly and the natural outward play of the gear as it made ground contact. In service, the main gear assembly provided very few difficulties. F59

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-u s .. \R.: The chine-like tunnels that accommodated much of the X-IS's plumbing and control system assemblies were faired onto each of the fuselage sides beginning just aft of the cockpit and continuing all the way to the end of the empennage section.

The chine-like tunnels were affected by the high temperatures generated by high-Mach flight like all other parts of the X-IS and accordingly, expansion joints were designed as an integral element to allow the skin to expand without distortion .

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Because of its unique propulsion system and the difficulty in bleeding power off to accommodate the needs of various subsystems such as the hydraulic system and electrical power supply, an auxiliary power unit which ran off its own propellant source was installed in the X-IS's fuselage, just behind the cockpit. This unit generated enough power to meet subsystem requirements for any standard or emergency situation. The exhaust dumps for this unit were permitted to protrude above the aircraft skin.

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f' '{~ The X-IS's aft fuselage section and empennage served to support the horizontal and ventral/dorsal tail surfaces, the skid main landing gear, and the engine compartment. Not often noticed was the flare in the chine-like tunnel assembly that occurred where the slab stabilators intersected the aft fuselage. This was done to accommodate the stabilator actuation ram and associated support assemblies. Visible in this view is the retracted skid main gear.

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iX-15 CUTAWAY Internsl Equipment Compartment Ballistic Control Rockets

Nose Instrument Compartment

Tanks

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Side Fairing

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X-15 IN-FLIGHT TEMPERATURES Unprotected Inconel-X

·2400° Vmax =8,000 fps Alt =100,000 ft Note: Temperatures In parenthesis are from basic X·15 high All temperatures are shown In degrees Fahrenheit

temperature flights

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rX-15 ST::URAL PROBLEMS

TYPICAL WING STRUCTURE

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The X-15 was equipped with a conventional single-piece flap on each wing trailing edge. This unit, which had a blunt trailing edge like the rest of the wing, was utilized during final approach and landing to increase the aircraft drag coefficient and decrease the landing roll-out distances. The flap was hydraulically actuated by a single ram mounte? inside the fuselage chine/tunnel assembly. It was attached to the wing by two large hinges.

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Special temperature sensitive paints, generically known as "templac", were utilized many times during X-15 flights to determine heat patterns and hot spots on the aircraft skin and structure. This view illustrates the visuai results of "templac's" application; as can be seen, the wing's internal structure, which absorbed heat during the course of the flight, has become visible through the skin. The leading edge is obviously the hottest part of the wing.

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The X-IS's airfoil section was not particularly radical. Even more interesting was the fact that a relatively blunt leading edge, dictated by wind tunnel work conducted at the NACA's Langley facility, had been determined optimum for the X-IS's performance envelope. Wing tip cap, as can be seen in this view, was flat.

Because the wing leading edge was expected to expand more than other parts of the wing during flight, an expansion slot was left between it and the fuselage.

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View of the right wing tip cap gives excellent impression of conventional airfoil section and blunt leading edge. Noteworthy was the wing's blunt trailing edge. End cap was Inconel-X channel riveted to the upper and lower wing skins with clearances to permit heat expansion.

The X-IS was equipped with all-moving slab-stabilators, which were possibly the first of their kind ever installed on a supersonic aircraft. Aerodynamic forces were inputted by either differential or symmetrical operation.

The airfoil section and construction technique used in the X-IS's slab stabilators was similar to that utilized in its wings. The leading edge was a solid bar of Inconel-X. Like the wings, the slab stabilators also had a blunt trailing edge.

The slab stabilators were attached to the aft end of the chine-like tunnel fairings at a single mounting point. Just ahead of the hinge point, the chines acquired a flat side profile to provide a nominal aerodynamic interface for the root section.

Where the slab stabilator hinge points and carry-through assemblies were located was very tightly faired into the chine-like side tunnels. The stabilator spar was a bar-like structure of Inconel-X construction which protruded into the tunnel.

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A distinctive sharp leading edge, made of Inconel-X, was attached to the upper component of the vertical tail assembly of 56-6672 to study heating and accoustical anomalies that might be created by this conliguration.

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The dorsal tail assembly consisted of a stub lixed surface and a single-piece upper surface which served as the rudder assembly. Attached to the stub section was a two-surface airbrake which was hinged at the leading edge of each half.

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The corrugated blunt trailing edge of the rudder section was designed to accommodate heat-generated expansion requirements. The airbrake was mounted below.

The ventral tail surface and its associated airbrake assembly were quite similar to those mounted dorsally. Construction was primarily of Inconel-X alloy. Noteworthy in this view is the support assembly of the ground dolly which attached to the aircraft under the ventral tail surface stub. "\t,)'i

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The airbrakes were hydraulically actuated by a single ram unit attached to a large, hinged yoke with two spreader arms. As the piston for the ram moved forwards or backwards, it forced the brake surfaces to move into and out of the airstream.

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When the airbrakes were in their closed position, the actuation unit filled much 01 the brake housing's internal volume. Miscellaneous dump lines for various propellant units also were routed through this area for venting purposes.

The hydraulically-actuated slab-type airbrakes were normally faired-in as the rear component of each of the tail surface stub sections. Each airbrake surface, which consisted of an external plate backed with a corrugated inner panel, was equipped with two hinges at its respective forward edge. A total of four brake surfaces provided aerodynamic drag factors that could be utilized throughout almost all of the X-15's flight envelope, including reentry.

X-15 RESEARCH SYSTEMS & S E N S O R S - - - - - - - - - - - - - - - - - - - - - - - - - ,

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Research sensors were usually controlleeJ Oy the pilot from the cockpit during flight. A control panel for an ultra-violet sensing device, one of many such experiment panels instal/ed in the X-15 can be seen in this view of 56-6670.

There were a variety of pitot-type sensors permanently attached to the X-15 to project into the high-speed airstream. Typical of these sensors was this ventral/v-mounted unit which was raked toward the rear of the aircraft.

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Optical ports were instal/ed in the X-15 in several locations to provide external sensor capabilities for various photographic systems and related sensors. X-15, 56-6670, was equipped with an elaborate ventral port to accommodate special high-altitude reconnaissance system optical sensors for operational reconnaissance platforms such as the Lockheed SR-71A. The X-15 proved ideal for such test projects because of its extraordinary speed and altitude capabilities.

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Though similar in objective to the optical accommodation seen on X-15, 56-6670, this sensor bay, on the X-15A-2, was provided with a less-pronounced fairing that retained its optical glass port through the end of this aircraft's flight test program.

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There were several equipment bays available in the X-15 to accommodate various cameras and atmospheric samplers. Several of these bays were equipped with snapaction doors which could be opened in flight to accommodate test objectives.

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DUMMY SCRAMJET DIMENSIONS

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Several scramjet configurations were tested while suspended under the aircraft's ventral tail surface. The remains of the dummy scramjet unit unsuccessfully tested during AF test pilot Pete Knight's world speed record flight of October 3, 1967, are seen in this view.

An infrared horizon-scanner, with cover plate removed, was mounted behind the upper airbrakes of 56-6672. It measured space background noise. "" I'"T-r--......

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The X-IS could be configured to carry wingtip pods. These somewhat obtrusive units were designed to accommodate a wide variety of sensors, including cameras and micro-meteorite samplers. The pods were attached to the wingtips with riveted titanium ftanges. Though somewhat heavy, their addition did not degrade the X-IS's performance substantially, and they remained attached to 56-6670 through its retirement from active flight test.

BLUE S C O U T - - - - - - - - - - -

Drawing Illustrates the proposed two-stage Blue

SCout rocket for launching from the X-15. It con8l8t8 01 the upper stages of the USAF Blue Scout Junior, with 8 small spherical rocket tor Injecting the payload Into orbit.

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Artist's rendering of a "Blue Scout" rocket following launch from an X·15 at 156,000 feet. The recoverable booster was proposed by Ford and North American.

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Scheduled delays in the delivery of the XLR99 powerplant led to a decision to consummate the X-IS powered flight program with a twin-pack XLRll propulsion system. Two XLRll's were placed in a special mount desianed specificallv for the X-IS.

The installed XLR 11 unit presented eight exhaust nozzles in a relatively neat configuration that required a short empennage exlension for accommodation purposes. The dump tube arrangement for Ihe XLRll remained similar to that required for the XLR99.

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The XLR99 utilized liquid oxygen and anhydrous ammonia for propellants. In the standard X-IS it was capable of operating for 87 seconds at maximum thrust and 156 seconds at minimum thrust. Gas generalor hydrogen peroxide flow governed thrust.

The gas generalor was mounled at the powerplant's forward end and exhausted through a large tube that ran the length of the unit. A large number of dump tubes was assembled underneath to accommodale other propellant exhaust requirements. '~

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The XLR99's exhaust nozzle was a regeneratlvely COOled unit that depended on fuel circulating through an integral tube system to control nozzle temperatures. No less than eighteen dump lines exiled around the perimeter of the exhaust nozzle.

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There were numerous propellant dump lines. The two largest were mounted on either side of the empennage section; one accommodated the dumping requirements of the LOX (left), and the other, the dumping requirements of the anhydrous ammonia.

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The anhydrous ammonia dump (vent) line (left) exited the aft end of the right side chine in a long tube that was firmly strapped to a hydrogen peroxide vent line. The LOX dump (vent) line exited the aft end of the left chine in a long tube that was firmly strapped to two other dump lines. The tubes were mostly of stainless steel as their heat and strength requirements were not as high as that for the aircraft's skin and structural members. ti

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A 20-inch diameter helium tank was permanently mounted on the X-15A's empennage, just to the rear of the vertical tail assembly. The helium from this tank, at 3,600 psi, was utilized to expell the liquid hydrogen in the liquid hydrogen tank, and at the same time, create a non-combustible atmosphere.

X-15's, 55-6670 and 56-6671, on the rocket powerplant test site at Edwards AFB, on May 25, 1960. The fortified concrete blockhouse is visible in the center.

XLR99 SOUND PRESSURE LEVELS - - - - - - - - - r

1--25FT-j fh~-+-r----,r---12 FT 148 DB ~ 163 08 VERTICAL 148 D8 VERTICAL 5 FT ABOVE GROUND

147 DB ENGINE BAY 153 DB HORIZ STAB ( LOWER)

The X·15A-2 was designed for increased performance based on its ability to carry a fuel load that was significantly greater than that of the standard X-15. Jettisonable external propellant tanks permitted acceleration out to Mach 2.

(100$ thrust)

156 DB HORIZ STAB ( UPPER)

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The NB-52AfB pylon assembly was a permanently attached unit of primarily aluminum construction. It was designed to support the payload weights associated with . transportation of the X-15, and at the same time, provide an aerodynamically clean fairing for such items as the liquid oxygen and anhydrous ammonia top-off plumbing systems and various communications lines. Attachment points were located at the forward and aft ends of the pylon.

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.J../~ The launch officer's station inside the NB-52A1B was located in the defensive systems officer's station, to the rear of the pilot. Panels were provided giving the launch officer control over communications, propellants, and various related sub-systems. Of primary importance was the top-off operation which required that the X-15 be given a final oxydizer and fuel load immediately prior to launch to compensate for the fuel and oxydizer that had been lost during the ascent.

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The NB-52A (shown) and the NB-52B underwent major modifications that included elimination and reconfiguration of their rear-facing defensive armament stations.

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The NB-52A (shown) and the NB-52B were equipped with forward fuselage modifications that included "bug-eye" photographic and manual monitoring housings and the addition of several observation windows. Cameras were mounted in both front and rear fuselage positions to document all facets of any given X-15 launch. ~

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In order to provide the proposed delta wing X-15 with a speed increment significantly higher than that provided by the NB-52A1B, a study was conducted calling for the X-15 to be launched from the back of the North American XB-70A.

One of the most radical advanced X-15 proposals was this delta wing configuration that appears to have been powered by a pair of faired ventraf scramjets. Noteworthy are the canards mounted just under the cockpit windscreen/canopy.

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The advanced X-15-pilot-dedicated MC-2 high-altitude suit was the first successful full-pressure suit. Earlier versions were of the parlial pressure variety.

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AEROFAX, INC. is pleased to announce the release of the initial titles in its new DATAGRAPH monographic aviation history series. These books are designed to accommodate aircraft histories and related subject areas that are either too large for the smaller MINIGRAPH series or too small for the larger, definitive AEROGRAPH series. Like the MINIGRAPHS and AEROGRAPHS, the DATAGRAPH titles are designed to provide exceptional subject coverage via numerous well-reproduced photographs, an extremely detailed and comprehensive text, and extraordinary quality. Each of these authoritative references has been created for the serious enthusiast and modeler and is designed to provide unparalleled textual and pictorial detail not usually found in other readily available books of this type. Each DATAGRAPH contains over 150 photos, fold-out-type multi-view drawings, color scheme information, systems drawings, and related reference material. If you find the new DATAGRAPH series to your liking and would like to have your name added to our mailing list to receive, free of charge, our quarterly AEROFAX NEWS, please drop us a line at P.O. Box 120127, Arlington, Texas 76012. We would like to hear from you and would particularly appreciate comments, criticisms, and suggestions for future titles. AEROFAX also is in need of interesting, previously unpublished photos of aircraft for use in forthcoming MINIGRAPH, DATAGRAPH, and AEROGRAPH titles. If you have such items in your files, please consider loaning them to AEROFAX so that others can see them, too. You will, of course, be credited if your photo is used, and a free copy of the publication in which It is used will be sent. AEROFAX looks forward to hearing from you. Thanks for you interest, Jay Miller and the AEROFAX, INC. Staff

Boeing NB-52A, 52-003, and North American X-15, 56-6671, are seen over Edwards AFB. The X-15 was still equipped with a twin-pack Reaction Motors XLR11 propulsion system. Visible on X-15 fuselage undersurface is condensation from liquid oxygen tank. Dark gray-green paneling on the upper forward NB-52A fuselage section near observation "bug-eye" is noteworthy. Also visible are white-painted NB-52A flaps and lower nose radome.