Category X-15 EXTENDING THE FRONTIERS OF FLIGHT

At the beginning of the X-15 program, researchers used the methods developed by Edward Van Driest and Ernst Eckert to determine the heat-transfer coefficients for temperature calculations. However, the measured heat-transfer coefficients during the early flight program were considerably lower than the predicted values. Based on these preliminary results, derived primarily from the initial low-angle-of-attack flights, engineers modified Eckert’s turbulent-flow method to produce the adiabatic-wall reference-temperature method.[48]

fuselage. The boundary-layer transition was completely unpredictable, but since researchers expected turbulent flow during the major portion of most flights, they normally used turbulent – flow calculations for the entire flight. Next came determining the heat-transfer coefficients, and finally calculating the skin temperature. Due to the tedious work involved in this process, which was done mostly by hand since general-purpose computers were not yet in widespread use, the researchers made many assumptions that simplified the procedure. For instance, it was assumed that temperature did not vary through the thickness of the skin, no heat was transferred along the skin, the specific heat of the skin was constant, solar radiation to the skin was negligible, the emissivity of the skin was constant, and no net heat transfer occurred between surfaces by radiation.[49]

Temperatures calculated using the adiabatic-wall reference-temperature method tended to agree closely with measured data from the flight program. In several instances the calculated temperature was somewhat higher because the analytical method assumed turbulent flow all of the time. This was considered reasonable and sufficient for flight-safety purposes since it erred on the side of caution.-50

In 1957, Lockheed Aircraft Company developed a thermal analyzer program that ran on an IBM 704 digital computer, the largest of its type then available. This program was capable of running the heating prediction equations, including the effects of transient conduction, convection, radiation, and heat storage, that researchers had previously omitted for the sake of expediency. With Lockheed’s assistance, researchers modified the program to reflect the X-15 configuration. The program estimated the heat input to the skin elements using the attached-shock Prandtl- Meyer expansion method for flow conditions, and the adiabatic-wall reference-temperature method for heat transfer. Researchers used the laminar-flow theory of Fay and Riddell to compute the heat input to the stagnation points, with curves developed by Lester Lees used to weight the periphery.-50

One of the primary goals of the X-15 program was to validate the various heat-transfer methods with actual flight results. Many of the early X-15 flights were dedicated to gathering data that the researchers would spend years comparing against wind tunnel and theoretical results. The results were vastly improved heat-transfer models that were used during the Apollo and space shuttle programs. (NASA)

To accompany the Lockheed-developed software, North American developed two other programs to predict structural heating values and their distribution along the airframe. The first program computed local-flow conditions on the aircraft, and the second program used the local-flow conditions to calculate the aerodynamic heat transfer to the skin. The program developed by Lockheed calculated the transient heating of internal structure based on the results of the other two programs.-1521

To evaluate the acceptability of the thermal analyzer program, researchers compared calculated results with actual flight results on several occasions. The values always compared favorably, and were usually slightly better than the hand-calculated values for the same conditions. North American and NASA quickly adopted the automated process based largely on the tremendous labor savings it offered.

After the flight planners established a flight profile on the fixed-base simulator, they digitized the results of a clean flight and input them into the IBM 704 to predict the skin and structural temperatures and thermal gradients for the flight. This was a time-consuming process. Researchers then compared the resulting data with the design conditions to ensure that the X-15 did not violate any structural margins. If any exceptions were uncovered during the comparison, researchers modified the flight profile and the entire process was repeated. Emergency and contingency flight profiles went through the same rigorous process. After the flight, researchers compared the heating predictions with actual flight data and then refined the simulations.-1531

At the October 1953 meeting of the Air Force Scientific Advisory Board (SAB) Aircraft Panel, Chairman Clark B. Millikan asked panel members for their ideas on future aircraft research and development programs. The panel decided that "the time was ripe" for another cooperative (USAF – NACA) research airplane project to further extend the frontiers of flight. Millikan released a statement declaring that the feasibility of an advanced manned research aircraft "should be looked into." The panel member from NACA Langley, Robert R. Gilruth, would later play an important role in coordinating a consensus between the SAB and the NACA.-1771

Contrary to Sanger’s wartime conclusions, by 1954 most experts within the NACA and industry agreed that hypersonic flight would not be possible without major advances in technology. In particular, the unprecedented problems of aerodynamic heating and high-temperature structures appeared to be a potential "barrier" to sustained hypersonic flight. Fortunately, the perceived successes enjoyed by the X-planes led to increased political and philosophical support for a more advanced research aircraft program. The most likely powerplant for the hypersonic research airplane was one of the large rocket engines from the missile programs. Most researchers now believed that manned hypersonic flight was feasible, but it would entail a great deal of research and development. Fortunately, at the time there was less emphasis than now on establishing operational requirements prior to conducting basic research, and, perhaps even more fortunately, there were no large manned space programs that would compete for funding. The time was finally right.1781

The hypersonic research program most likely originated during a meeting of the NACA Interlaboratory Research Airplane Projects Panel held in Washington, D. C., on 4-5 February 1954. The panel chair, Hartley A. Soule, had directed the NACA portion of the cooperative USAF-NACA research airplane program since 1946. In addition to Soule, the panel consisted of Lawrence A. Clousing from Ames, Charles J. Donlan from Langley, William A. Fleming from Lewis, Walter C. Williams from the HSFS, and Clotaire Wood from NACA Headquarters. Two items on the agenda led almost directly to the call for a new research airplane. The first was a discussion concerning Stone’s proposal to use a modified X-2, with the panel deciding that the aircraft was too small to provide meaningful hypersonic research. The second was a proposal to develop a new thin wing for the Douglas D-558-2. This precipitated a discussion on the "advisability of seeking a completely new research airplane and possible effects on such a proposal on requests for major changes to existing research airplanes." The panel concluded that the research utility of the D – 558-2 and X-2 was largely at an end, and instead recommended that NACA Headquarters request detailed goals and requirements for an entirely new vehicle from each of the research laboratories. This action was, in effect, the initial impetus for what became the X-15.1791

On 15 March 1954, Bob Gilruth sent Clark Millikan a letter emphasizing that the major part of the research and development effort over the next decade would be "to realize the speeds of the existing research airplanes with useful, reliable, and efficient aircraft under operational conditions" (i. e., developing Mach 2-3 combat aircraft). Gilruth further noted that a "well directed and sizeable effort will be required to solve a number of critical problems, by developing new materials, methods of structural cooling and insulation, new types of structures, and by obtaining a thorough understanding of the aerodynamics involved." Because many of the problems were not then well defined, "design studies should be started now for manned research aircraft which can explore many of these factors during high-speed flight" and which would be capable of "short excursions into the upper atmosphere to permit research on the problems of space flight and reentry." It was a surprising statement.1801

During the late 1940s and early 1950s, the overwhelming majority of researchers thought very little about manned space flight. Creating a supersonic airplane had proven difficult, and many researchers believed that hypersonic flight, if feasible at all, would probably be restricted to missiles. Manned space flight, with its "multiplicity of enormous technical problems" and "unanswered questions of safe return" would be "a 21st Century enterprise."1811

Within a few years, however, the thinking had changed. By 1954 a growing number of American researchers believed that hypersonic flight extending into space could be achieved much sooner, although very few of them had the foresight to see it coming by 1960. Around this time, the military became involved in supporting hypersonic research and development with a goal of creating new weapons systems. During 1952, for example, the Air Force began sponsoring Dornberger’s manned hypersonic boost-glide concept at Bell as part of Project BoMi.-82

BoMi (and subsequently RoBo) advanced the Sanger-Bredt boost-glide concept by developing, for the first time, a detailed thermal-protection concept. Non-load-bearing, flexible, metallic radiative heat shields ("shingles") and water-cooled, leading-edge structures protected the wings, while passive and active cooling systems controlled the cockpit temperature. NACA researchers, including the Brown study group, read the periodic progress reports of the Bell study-classified Secret by the Air Force-with great interest. Although most were skeptical, a few thought that the project just might work. The Air Force would also fund similar studies by other contractors, particularly Convair and, later, Boeing.-1831

In response to the recommendation of the Research Airplane Projects Panel, NACA Headquarters asked its field installations to explore the requirements for a possible hypersonic research aircraft. Based on the concerns of the 1952 Langley study group, as well as data from Bell regarding BoMi research, it was obvious that a primary goal of any new research airplane would be to provide information about high-temperature aerodynamics and structures. The missile manufacturers concurred.-1841

In response to NACA Headquarters’ request, all of the NACA laboratories set up small ad hoc study groups during March 1954. A comparison of the work of these different NACA groups is interesting because of their different approaches and findings. The Ames group concerned itself solely with suborbital long-range flight and ended up favoring a military-type air-breathing (rather than rocket-powered) aircraft in the Mach 4-5 range. The HSFS suggested a larger, higher – powered conventional configuration generally similar to the Bell X-1 or Douglas D-558-1 research airplanes. The staff at Lewis questioned the need for a piloted airplane at all, arguing that ground studies and the PARD rocket-model operation could provide all of the necessary hypersonic information at much less cost and risk. Lewis researchers believed that possible military applications had unduly burdened previous research airplane programs, and there was no reason to think anything different would happen in this case.-1851

On the other hand, Langley chose to investigate the problem based largely on the hypersonic research it had been conducting since the end of World War II. After the 11-inch hypersonic tunnel became operational in 1947, a group headed by Charles McLellan began conducting limited hypersonic research. This group, which reported to John Becker, who was now the chief of the Aero-Physics Division, provided verification of several newly developed hypersonic theories while it investigated phenomena such as the shock-boundary-layer interaction. Langley also organized a parallel exploratory program into materials and structures optimized for hypersonic flight. Perhaps not surprisingly, Langley decided to determine the feasibility of a hypersonic aircraft capable of a 2- to 3-minute excursion out of the atmosphere to create a brief period of weightlessness in order to explore the effects of space flight. Hugh Dryden would later liken this excursion to the leap of a fish out of water, and coined a new term: space leap.-861

Three men that played important parts in the X-15 program. On the right is Walter C. Williams, the head of the High-Speed Flight Station and a member of the Research Airplane Projects panel that guided the X-15 through its formative stages. In the middle, Hugh L. Dryden, the Director of the NACA. At left is Paul F. Bikle, who came late to the X-15, but guided it through most of its flight program as the director of the Flight Research Center. (NASA)

Langley’s ad hoc hypersonic aircraft study group consisted of John Becker (chairman); Maxime A. Faget,[87] a specialist in rocket propulsion from the Performance Aerodynamics Branch of PARD; Thomas A. Toll, a control specialist from the Stability Research Division; Norris F. Dow, a hot – structures expert from the Structures Research Division; and test pilot James B. Whitten. Unlike the earlier Brown study group, this group intentionally included researchers with previous experience in hypersonics.[88]

The group reached a consensus on the objectives of a hypersonic research aircraft by the end of its first month of study. Although one of the original goals was to investigate the effects of weightlessness, the members soon realized "that the problems of attitude control in space and the transition from airless flight to atmospheric flight during reentry were at least equally significant." The group also began to consider the dynamics of the reentry maneuvers and the associated problems of stability, control, and heating as the most pressing research need. However, another objective would come to dominate virtually every other aspect of the aircraft’s design: research into the 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 problem, as well as the primary research objective.-1891

10 vehicle "would require a much greater expenditure of time and effort" yet "would add little in the fields of stability, control, piloting problems, and structural heating." Considering that no human had yet approached Mach 3, even Mach 7 seemed a stretch.[90]

By the end of April 1954, Becker’s group had completed a tentative design for a winged aircraft and an outline of proposed experiments. The group kept the configuration as conventional as possible to minimize the need for special low-speed and transonic developments without compromising its adequacy as a hypersonic, aerodynamic, and structural research vehicle. However, acknowledging what would become a continuing issue; the group did not consider any of the large rocket engines then under development entirely satisfactory for the airplane. In the absence of the rapid development of a new engine, the group hoped a combination of three or four smaller rocket motors could provide hypersonic velocities.-1911

At this point Floyd Thompson, by now the associate director at Langley, influenced the direction of the Becker study. He made a suggestion that echoed John Stack’s 1945 recommendation that the Bell XS-1 transonic research airplane use a 12% thick wing that would force it to encounter the compressibility efforts that aerodynamicists were most interested in studying. Since the hypersonic airplane would be the first in which aero-thermal-structural considerations constituted the primary research problem, Thompson argued that the aim of the aircraft "should be to penetrate as deeply as possible into the region of [high aerodynamic] heating and to seek fresh design approaches rather than makeshift modifications to conventional designs." His suggestion became policy.-192

Wind-tunnel testing began in mid-1954 and continued through the end of 1955 using the basic Becker design. David E. Fetterman, Jr., Jim A. Penland, and Herbert W. Ridyard led the tests, mainly using the 11-inch tunnel at Langley. The researchers noted that previous hypersonic designs had "been restricted mainly to missile types which were not required to be able to land and which, therefore, had relatively small wings or wings of very low aspect ratio." The researchers concentrated on extrapolating existing data to the Becker design while making sure the concept would be acceptable for a manned aircraft, including the ability to land.-93

One particular feature, however, differed from later concepts. The initial wind-tunnel tests used a design that incorporated relatively large leading-edge radii for both the wing and vertical stabilizer. The large radii were believed necessary to keep the heat transfer rates within feasible limits. Eventually the researchers discovered the beneficial effects of a leading-edge sweep and found materials capable of withstanding higher temperatures. These allowed smaller radii, resulting in less drag and generally better aerodynamic characteristics. Although the baseline design changed as a result, by this time the researchers were concentrating on evaluating various empennage configurations and elected not to change the wing design on the wind-tunnel models to avoid invalidating previous results.94

While performing the original heating analysis of the proposed reentry from the "space leap," Becker and Peter F. Korycinski from the Compressibility Research Division ran head-on into a major technical problem. At Mach 7, reentry at low angles of attack appeared impossible because of disastrous heating loads. In addition, the dynamic pressures quickly exceeded, by large margins, the limit of 1,000 pounds per square foot (psf) set by structural demands. New tests of the force relationships in the 11-inch tunnel provided Becker and Korycinski with a surprising solution to this problem: if the angle of attack and the associated drag were increased, deceleration would begin at a higher altitude. Slowing down in the thinner (lower-density) atmosphere made the heat-transfer problem much less severe. In other words, Becker and Korycinski surmised, by forcing deceleration to occur sooner, the increased drag associated with

the high angle of attack would significantly reduce the aircraft’s exposure to peak dynamic pressure and high heating rates. Thus, by using "sufficient lift," the Langley researchers found a way to limit the heat loads and heating rates of reentry. Interestingly, this is the same rationale used 15 years later by Max Faget when he designed his MSC-002 (DC-3) space shuttle concept at the Manned Spacecraft Center.1951

On reflection, it became clear to the Becker group that the sufficient-lift concept was a "new manifestation" of Allen’s blunt-body theory and was as applicable to high-lift winged reentry as to the non-lifting missile warheads studied at Ames during 1952. As the group increased the angle of attack to dissipate more of the kinetic energy through heating of the atmosphere (and less in the form of frictional heating of the vehicle itself), the configuration became increasingly "blunt." Some form of speed brakes, again in accord with Allen’s concept, could increase drag and further ease the heating problem.1961

Throughout 1954 the heating problems of high-lift, high-drag reentry came under increasing scrutiny from key Langley researchers. However, another problem soon outweighed the heating consideration: making the configuration stable and controllable at the proposed high-angle-of – attack reentry attitude. Because they were venturing into a new flight regime, the researchers could not determine the exact hypersonic control properties of such a configuration. Nor were they certain they could devise a structure that would survive the anticipated 2,000°F equilibrium temperatures.-1971

The HSFS had forewarned Langley about potential hypersonic stability problems. In December 1953, Air Force Major Chuck Yeager had pushed the Bell X-1A far beyond its expected speed range. As the aircraft approached Mach 2.5, it developed uncontrollable lateral oscillations that nearly proved disastrous.1981 While Yeager frantically tried to regain control, the airplane tumbled for over a minute, losing nearly 10 miles of altitude. At subsonic speed, the aircraft finally entered a conventional spin from which Yeager managed to recover. This incident led to a systematic reinvestigation of the stability characteristics of the X-1A. By mid-1954, findings indicated that the problem that had almost killed Yeager was the loss of effectiveness of the X- 1A’s thin-section horizontal and vertical stabilizers at high speed. The HSFS was not equipped to conduct basic research into solutions, but it coordinated with Langley in an attempt to overcome this problem.

At the same time, Langley and the HSFS began investigating the inertial-coupling phenomenon encountered by the North American F-100A Super Sabre.1991

The Becker group faced a potential stability problem that was several times more severe than that of the X-1A. Preliminary calculations based on data from X-1A wind-tunnel tests indicated that the hypersonic configuration would require a vertical stabilizer the size of one of the X – 1’s wings to maintain directional stability-something that was obviously impractical. Stumped by this problem, Becker sought the advice of his 11-inch hypersonic tunnel researchers. The consensus, reached by wind-tunnel testing and evaluating high-speed data from earlier X-planes, was that an extremely large vertical stabilizer was required if the thin-section stabilizers then in vogue for supersonic aircraft were used. This was largely because of a rapid loss in the lift-curve slope of thin airfoil sections as the Mach number increased. In a radical departure, however, Charles McLellan suggested using a thicker wedge-shaped section with a blunt trailing edge. Some time before, McLellan had conducted a study of the influence of airfoil shape on normal-force characteristics, and his findings had been lying dormant in the NACA literature. Calculations based on these findings indicated that at Mach 7 the wedge shape "should prove many times more effective than the conventional thin shapes optimum for the lower speed." By modifying the proposed configuration to include the wedge-shaped vertical stabilizer, McLellan believed that a reasonably sized vertical stabilizer could correct most directional instability.11001

Charles H. McLellan at NACA Langley, one of the researchers that defined much of the X-15 configuration, proposed the use of a split training edge on the vertical stabilizer to form speed brakes. Perhaps even more importantly, these could also be opened to form a variable-wedge vertical stabilizer as a means of restoring the lift-curve slope at high speeds, thus permitting much smaller surfaces that were easier to design and imposed a smaller drag penalty at lower speeds. The ultimate X-15 configuration did not incorporate the split trailing edge, but the much – later space shuttles did. (NASA)

A new series of experiments in the 11-inch tunnel verified that a vertical stabilizer with a 10- degree wedge angle would allow the proposed aircraft to achieve the range of attitudes required by heating considerations for a safe high-drag, high-lift reentry. Further, it might be possible to use a variable-wedge vertical stabilizer as a means of restoring the lift-curve slope at high speeds, thus permitting much smaller surfaces that would be easier to design and would impose a smaller drag penalty at lower speeds. McLellan calculated that this wedge shape should eliminate the disastrous directional stability decay encountered by the X-1A.-1101

Becker’s group also included speed brakes as part of the vertical stabilizers to reduce the Mach number and heating during reentry. Interestingly, the speed brakes originally proposed by Langley consisted of a split trailing edge; very similar to the one eventually used on the space shuttles. As the speed brakes opened, they effectively increased the included angle of the wedge-shaped vertical stabilizer, and variable deflection of the wedge surfaces made it possible to change the braking effect and stability derivatives through a wide range. The flexibility this made possible could be of great value because a primary use of the airplane would be to study stability, control, and handling characteristics through a wide range of speeds and altitudes. Furthermore, the ability to reenter in a high-drag condition with a large wedge angle greatly extended the range of attitudes for reentry that were permissible in view of heating considerations.-102

Up until this time, the designers of supersonic aircraft had purposely located the horizontal stabilizer well outside potential flow interference from the wings. This usually resulted in the horizontal stabilizer being located partway up the vertical stabilizer, or in some cases (the F-104, for example) on top of the vertical stabilizer. However, researchers at the HSFS suspected that this location was making it difficult, or at times impossible, for aircraft to recover from divergent maneuvers. The same investigations at Langley that verified the effectiveness of the wedge-shape also suggested that an X-shaped empennage would help the aircraft to recover from divergent maneuvers.-110^

The Becker group recognized that the change from a conventional "+" empennage to the "X" configuration would present at least one major new problem: the X-shape empennage projected into the high downwash regions above and below the wing plane, causing a potentially serious loss of longitudinal effectiveness. Researchers at Langley looked for solutions to this new problem. By late 1954 they had an unexpected answer: locate a conventional "+" horizontal stabilizer in the plane of the wing, between the regions of highest downwash. This eliminated the need to use an X-shaped empennage, allowing a far more conventional tail section and control surfaces.^

Although it would come and go from the various preliminary designs, the use of a ventral stabilizer was beginning to gain support. Charles McLellan observed, "At high angles of attack, the effectiveness of the upper and lower vertical stabilizers were markedly different. Effectiveness of the upper tail decreases to zero at about 20 degrees angle of attack. The lower tail exhibits a marked increase in effectiveness because of its penetration into the region of high dynamic pressure produced by the compression side of the wing. Assuming the wing is a flat plate and the flow is two-dimensional, the dynamic pressure below the wing increases 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."[105]

On the structural front, the Becker study evaluated two basic design approaches. In the first, a layer of assumed insulation protected a conventional low-temperature aluminum or stainless steel structure. The alternative was an exposed "hot structure." This design approach and the materials used permitted high structural temperatures without insulation.[106]

Surprisingly, the temperatures expected on the high-altitude "space leap" were significantly higher than for the basic hypersonic research flights. Establishing a design that could withstand the 2,000°F equilibrium temperature was a challenge, and ultimately resulted in the hot-structure concept shown on the lower line of this chart. (NASA)

Analysis of the heating projections for various trajectories showed that the airplane would need to accommodate equilibrium temperatures of over 2,000°F on its lower surface. Unfortunately, no known insulating technique could meet this requirement. Bell was toying with a "double-wall" concept in which a high-temperature outer shell and a layer of insulator would protect the underlying low-temperature structure. This concept would later undergo extensive development, and several contractors proposed it during the X-15 competition, but in 1954 it was in an embryonic state and not applicable to the critical nose and leading-edge regions. However, the Becker group believed that the possibility of local failure of any insulation scheme constituted a serious hazard, as was later tragically demonstrated on the Space Shuttle Columbia. Finally, the problem of accurately measuring heat-transfer rates—one of the primary objectives of the new research aircraft program—would be substantially more difficult to accomplish with an insulated structure.[107]

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 was about 1,200°F, which was far below the estimated equilibrium temperature of 2,000°F. It was clear that some form of heat dissipation—either direct internal cooling or absorption into the structure itself —would be necessary. It was thought 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 structural material preferred by Langley for a hot-structure design.-11081 In mid-1954, Norris Dow began an analysis of an Inconel X structure while other researchers conducted a thermal analysis. In a happy coincidence, the results showed that the skin thickness needed to withstand the expected aerodynamic stresses was about the same as that needed to
absorb the thermal load. This meant that it was possible to solve the structural problem for this transient condition of the Mach 7 research aircraft with no serious weight penalty for heat absorption. This was an unexpected plus for the hot structure. Together with the fact that none of the perceived difficulties of an insulated-type structure (particularly the difficulty of studying structural temperatures) were present, this led the study group to decide in favor of an uninsulated hot-structure design.

Unfortunately, it later proved that the hot structure had problems of its own, especially in the area of non-uniform temperature distribution. Detailed thermal analyses revealed that large temperature differences would develop between the upper and lower wing skins during the pull – up portions of certain trajectories, resulting in intolerable thermal stresses in a conventional structural design. To solve this new problem, researchers devised wing shear members that did not resist unequal expansion of the wing skins. The wing thus was essentially free to deform both span-wise and chord-wise with asymmetrical heating. Although this solved the problem for gross thermal stresses, localized thermal-stress problems still existed near the stringer attachments. The study indicated, however, that proper selection of stringer proportions and spacing would produce an acceptable design that would be free of thermal buckling.-1109

The analyses produced other concerns as well. Differential heating of the wing leading edge resulted in changes to the natural torsional frequency of the wing unless the design used some sort of flexible expansion joint. The hot leading edge expanded faster than the remaining structure, introducing a compression that destabilized the section as a whole and reduced its torsional stiffness. To negate these phenomena, researchers segmented and flexibly mounted the leading edge to reduce thermally induced buckling and bending. Similar techniques found use on the horizontal and vertical stabilizers.

COMFARISQN OF INCONEL X WITH OTHER ALLOYS

TENSILE YIELO STRESS*

К 51

DESIGN TEMR

BTU/SQ FT/SEC

Langley evaluated many materials for the proposed hypersonic research airplane, but the availability of Inconel X and its exceptional strength at extremely high temperatures, made it,

almost by default, the preferred material for a hot-structure design. Coincidently, the researchers at NACA Langley discovered that the skin thickness needed to withstand aerodynamic stress was about the same as the amount of structure needed to absorb the thermal load from the high – altitude mission. (NASA)

Perhaps more worrisome was the question of potential propulsion systems. The most promising configuration was found to be four General Electric A1 or A3 rocket engines, due primarily to the "thrust stepping" this configuration provided.-1110! At the time, rocket engines could not be throttled (even today, most rocket engines cannot be). Several different techniques can be used to throttle a rocket engine, and each takes its toll in mechanical complexity and reliability. However, a crude method of throttling did not actually involve changing the output of the engine, but rather igniting or extinguishing various numbers of small engines. For instance, in a cluster of three 5,000-lbf engines, the available thrust levels (or "steps") would be 5,000, 10,000, and 15,000 lbf. Since most rocket engines were not restartable (again, the concept adds considerable mechanical complexity to the engine), once an engine was extinguished it could not be restarted. Thrust stepping or throttling allowed a much more refined flight profile, and largely defined the propulsion concept for the eventual X-15.-1111-

At this stage of the study, the vehicle concept itself was "little more than an object of about the right general proportions and the correct propulsive characteristics" to achieve hypersonic flight. However, in developing the general requirements, the Langley group envisioned a conceptual research aircraft that would serve as a model for the eventual X-15. The vehicle they conceived was "not proposed as a prototype of any of the particular concepts in vogue in 1954…[but] rather as a general tool for manned hypersonic flight research, able to penetrate the new regime briefly, safely, and without the burdens, restrictions, and delays imposed by operational requirements other than research." 112

Although the Becker group was making excellent progress, their continued investigation of the "space leap" caused considerable controversy. The study called for two distinct research profiles. The first-the basic hypersonic research flights—consisted of a variety of constant angle-of-attack, constant-altitude flights to investigate aero-thermodynamic characteristics. However, the second flight profile explored the problems of future space flight, including investigations into "high-lift and low-L/D [lift over drag] during the reentry pull-up maneuver." Researchers recognized that this was one of the principal problems for manned space flight from both a heating and piloting perspective.-113!

This brought yet more concerns: "As the speed increases, an increasingly large portion of the aircraft’s weight is borne by centrifugal force until, at satellite velocity, no aerodynamic lift is needed and the aircraft may be operated completely out of the atmosphere. At these speeds the pilot must be able to function for long periods in a weightless condition, which is of considerable concern from the aeromedical standpoint." By employing a high-altitude ballistic trajectory to roughly 250,000 feet, the Becker group expected that the pilot would operate in an essentially weightless condition for approximately 2 minutes. Attitude control was another problem since traditional aerodynamic control surfaces would be useless at very high altitudes. To solve this problem, the group proposed using small hydrogen-peroxide thrusters for attitude control outside the sensible atmosphere.

While the hypersonic research aspect of the Langley proposal enjoyed virtually unanimous support, it is interesting to note that in 1954 most researchers viewed the space-flight aspect with, at best, cautious tolerance. There were few who believed that any space flight was imminent, and most believed that manned space flight in particular would not be achieved until many decades in the future, probably not until the 21st century. For instance, John Becker remembers that even the usually far-sighted John Stack was "not really interested in the reentry problem or in space flight in general." Several researchers opined that the space-flight research was premature and recommended it be eliminated. Fortunately, it remained.114-

Langley’s work throughout 1954 demonstrated one thing: the need for flexibility. Since their inceptions, the Brown and Becker groups had run into one technical problem after another in the pursuit of a conceptual hypersonic aircraft capable of making a space leap. Conventional wisdom had provided experimental and theoretical guidance for the preliminary design of the configuration, but had fallen far short of giving final answers. Contemporary transonic and supersonic aircraft designs dictated that the horizontal stabilizer should be located far above or well below the wing plane, for example, but that was wrong. Ballistics experts committed to pointy-nosed missiles had continued to doubt the worth of Allen’s blunt-body concept, but they too were wrong. Conversely, the instincts of Floyd Thompson, who knew very little about hypersonics but was a 30-year veteran of the vicissitudes of aeronautical research, had been sound. The design and research requirements of a hypersonic vehicle that could possibly fly into space were so radically new and different, Thompson suggested, that only "fresh approaches" could meet them. He was correct.

The X-15 was breaking new ground when it came to structural materials, since it was obvious from the start that most of the wetted surface would be subjected to temperatures up to 1,200°F. Exotic materials made from the rare elements had not advanced sufficiently to permit quantity production of these expensive alloys, so the list of candidate materials was narrowed to corrosion resistant steels, titanium, and nickel-base alloys ("stainless steels"). The following table shows the strength properties of the candidate materials at room temperature; various aluminum alloys are included as a comparison. All properties are for bare sheet stock, except for the AM-355 bar stock. Materials marked with an asterisk were heat-treated.-1591

Although 6A1-4V titanium and AM-350 CRES had good strength efficiencies over a wide temperature range, both of the alloys tended to fall off rapidly above 800°F. Inconel X, on the other hand, had only a gradual drop in strength up to 1,200°F. Because of this stability, North American chose Inconel X for the outer skin for the entire airplane. Regular Inconel (as opposed to Inconel X) was not heat-treatable, but it could be welded and was used in locations where high strength was not of paramount importance or where final closeout welds were necessary following heat treatment of the surrounding structures. To accomplish this, Inconel lands were incorporated into Inconel X structures prior to final heat treatment, and access-hole cover plates made from

Inconel were welded to these lands.-601

North American used high-strength aluminum (2024-T4) to form the inner pressure shell of the cockpit and part of the instrumentation bay. As a relief from high thermal stresses, the company used titanium for the structure of the fuselage and wings. Originally, the company used two titanium alloys: 8-Mn, which was the highest strength alloy then available but was not recommended for welding, and 5A1-2.5Sn, which had acceptable strength and was weldable. Later, North American began using a high-strength and weldable alloy, 6A1-4V, in some areas.

To combat the high concentrated loads from the engine, most of the aft fuselage structure used titanium framing. The majority of the structure used fusion welding, although the company also used a limited amount of resistance welding. North American radiographically inspected all critical welds to ensure quality.-611

The material that presented the most problems was probably the 5Al-2.5Sn titanium, which proved to have inconsistent tensile properties that made it difficult to work with. It also exhibited low ductility and notch sensitivity, and had a poor surface condition. These problems existed in both rolled and extruded forms of the metal. The surface condition was the most important factor governing the formability of titanium, so North American had to remove all oxygen contamination, inclusions, and grind marks by machining, polishing, or chemically milling the metal prior to the final finishing. As a result, North American procured titanium extrusions for the X-15 with sufficient extra material in all dimensions to allow technicians to machine all surfaces prior to use.-621

The limited amount of stretch and shrink that was possible with a titanium extrusion during stretch wrapping presented a different problem when North American went to form the side fairing frames. Each frame was composed of four titanium 5Al-2.5Sn extrusions. One of the problems was that the inside flanges were located in areas that had small bend radii, and it was necessary to prevent compression failure. The small bend radii were "relieved" (some material was removed prior to bending), and a gusset was later welded in to fill the relieved area. The alternative would have been to reduce compression by increasing the pull on the forming machine, thus shifting the bend axis closer to the inboard edge. This, however, would have resulted in a tension failure on the outboard flange.-631

North American found that one of the more interesting aspects of titanium was that a formed part was prone to crack until the residual stresses resulting from the forming had been removed. This delayed cracking could occur within a few minutes, or it might not become evident until weeks later. In response, North American initiated a process that provided stress relief for all parts except "slightly" formed parts, such as skin panels, since they exhibited few problems.-641

Forming the seven different pressure vessel configurations in the X-15 presented its own problems. When compatibility with the contained fluid permitted, titanium was the first choice of material. North American used a 26-inch Cincinnati Hydroform for the hemispherical ends of the 14-inch cylindrical nitrogen tanks with little difficulty. The company also attempted to form the 16-inch hemispheres for the helium tanks on this machine, but the optimum blank size was greater than the maximum machine capacity of 26 inches. Using a smaller-than-optimum blank required excessive hold-down pressure that resulted in small surface cracks. The alternative was to "spin form" the hemispherical ends. Engineers heated the blanks to approximately 1,600°F and used an internally heated spinning chuck to shape the disc. Unfortunately, this resulted in a surface with significant oxygen contamination, so North American used thicker parts and machined them to the correct thickness to eliminate the contamination. Machining was also required to match the hemispheres for each end of the tank prior to welding.-651

Finding the correct material for the main propellant tanks, especially the liquid-oxygen tank, took some investigation. Most steel and common heavy structural alloys gain strength but lose ductility when operated at low temperatures, although Inconel proved to be relatively insensitive to this. The martensitic alloys, such as heat-treated 4130 low-alloy steel and AM-350 CRES precipitation-hardened corrosion-resistant steel, followed predictable curves that showed severe ductility loss as the temperature decreased below -100°F. A titanium alloy containing 5% aluminum and 2.5% tin handled the low temperatures well, but did not have the requisite strength at 1,200°F. North American finally decided to manufacture the primary barrels of the tanks from Inconel X.1661

Initially, engineers used AM-350 CRES, formed on a 7,000-ton hydraulic press using a deep-draw process, for the 32-inch hemispheres of the main propellant tanks. Excessive thinning occurred until the optimum pressure on the press draw ring was determined. Even then, North American encountered some difficulty due to uneven forces from the pressure pins used to secure the blanks, resulting in non-uniformity around the periphery of the hemisphere. The engineers subsequently decided to discard the CRES hemispheres and to remanufacture them from Inconel X.1671

Inconel X proved to be remarkably easy to work with considering its hardness, although the engineers had to make severely formed parts in multiple stages, with annealing accomplished between each stage. Nevertheless, problems arose. One of the first concerned fabricating the large Inconel propellant tank hemispheres. The propellant tanks comprised a large portion of the fuselage and were composed of an outer cylindrical shell and an inner cylinder. Inconel X semi­torus hemispheres at each end of the tank joined these two parts. The hemispheres were formed in two segments, with the split located midway between the inner and outer cylinders. Technicians welded the inner torus segment to the inner cylinder, and the outer torus segments to the outer tank, before joining the two assemblies.1681

After initial attempts to spin the bulkheads from a single, heated Inconel X blank were unsuccessful, the technicians built up the cones by welding smaller pieces together, and performed a complete X-ray inspection of each weld. After the cones were formed to the approximate size, they went through several stages of spinning, with a full annealing process performed after each stage. The first spin blocks used for the hemispheres were made from hardwood, and cast iron was used for the final sizing. A problem developed when transverse cracks began to appear during the spinning of the hemispheres.1691

Both North American and the International Nickel Company investigated the cracks, but determined that the initial welds were nearly perfect and should not have contributed to the problem. Nevertheless, engineers tried different types of welding wire, and varied the speed, feed, and pressure of the spinning lathe, but the welds continued to crack. It was finally determined that the welds were—ironically—too good; they needed to be softer. North American developed a new process that resulted in slightly softer but still acceptable welds, and the cracking stopped.1701

Fabricating the X-15 gave North American engineers some of the first large-scale experience with the newest high-strength alloys of titanium and stainless steel. The main propellant tanks formed an integral part of the fuselage, and after a great deal of investigation, North American manufactured the barrels from Inconel X. The experience gained from building the X-15 provided lessons used during the construction of the Apollo capsules and space shuttle orbiters. (North American Aviation)

North American gained experience in manufacturing the propellant tanks and fuselage structure long before it manufactured the first flight airplane. The company constructed three partial fuselages as ground-test articles for the rocket engines. Reaction Motors at Lake Denmark received two of these, while the third went to the Rocket Engine Test Facility at Edwards. Although

not intended as "practice," they did allow the workers in Inglewood to gain a certain level of expertise on a less-critical assembly before building the real flight articles.-171!

Forming the ogive section of the forward fuselage also presented some problems for North American. The usual method to construct such a structure was to form four semicircular segments of skin and weld them together. However, due to the size of the structure and the need to maintain a precise outer mold line, the engineers decided that the most expedient production method was to make a cone and bulge-form it into the final shape in one operation. The initial cone was made from four pieces of Inconel X welded together and carefully inspected to ensure the quality of the welds. It was then placed in a bulge-form die and gas pressure was applied that forced the part to conform to the shape of the die. This process worked well, with one exception. For reasons that were never fully understood, one of the four pieces of Inconel X used for one cone had a tensile strength about 28,000 psi greater than the others. During formation this piece resisted stretching, causing the welds to distort and creating wrinkles. North American eventually discarded the piece and made another one using four different sheets on Inconel; that one worked fine.^

Both titanium and Inconel were hard metals, and the tools used to form and cut them tended to wear out faster than equivalent tools used in the production of steel or aluminum parts. In addition, it took considerably longer to cut or polish compared to other metals. For instance, it took approximately 15 times longer to machine Inconel X than aluminum. This did not lead to any particular problems during the manufacture of the X-15 (unlike some of the tool contamination issues faced by Lockheed on the Blackbird), but it did slow progress and force North American to rethink issues such as machining versus polishing.!73!

The windshield glass originally installed on the X-15 was soda-lime-tempered plate glass with a single outer pane and double inner panes. Engineers had based this choice on a predicted maximum temperature of 740°F. Data obtained on early flights indicated that the outer face would encounter temperatures near 1,000°F, with a differential temperature between panes of nearly 750°F. It was apparent that soda-lime glass would not withstand these temperatures. The engineers subsequently selected a newly developed alumino-silicate glass that had higher strength and better thermal properties as a replacement. The 0.375-inch-thick alumino-silicate outer pane withstood temperatures up to 1,500°F during one test. The next test subjected the glass to a surface temperature of 1,050°F with a temperature gradient from the outer to inner surface of 790°F without failure. In actuality, the thermal environment on the X-15 glass was more complicated, although slightly less severe. The outer surface could reach 800°F, while the inner surface could reach 550°F; however, the inner temperature lagged behind the outer temperature. During rapid heat build-up on high-speed missions, the maximum temperature differential reached 480°F at a time when the outer glass was only 570°F. At this point, both the outer and inner panes began to rise in temperature rapidly.774

Technicians at the Flight Research Center installed the alumino-silicate glass in the outer pane of all three X-15s, although they continued to use soda-lime plate glass for the inner panes until the end of the program. Corning Glass Company supplied all of the glass. The thermal qualification test was interesting. Corning heated an 8.4 by 28-inch panel of the glass to 550°F in a salt bath for 3 minutes, and then plunged it into room-temperature tap water. If it did not shatter, it passed the test.!75!

welding, and otherwise joining this material to make a practical machine." Storms described special techniques for contouring the skins that involved hot machining, cold machining, ovens, freezers, cutters, slicers, and rollers. For instance, one special tool fixture needed to control the contour during a heat-treating cycle of the wing skin weighed 4,300 pounds, while the skin it held weighed only 180 pounds. Despite the publicity normally associated with the use of Inconel X, Charlie Feltz remembered that titanium structures gave North American the most trouble. Fortunately, the use of titanium on the X-15 was relatively small, unlike what Lockheed was experiencing across town on the Blackbird.[76]

In June 1959, the $450,000 Rocket Engine Test Facility at Edwards AFB came on line to provide local testing of the XLR99, although it would be almost a year before an XLR99 was available to use in it. This test facility provided a capability for engine checkout and pilot and maintenance – crew familiarization, as well as limited development firings. There were two test areas with a large blockhouse between them that contained various monitoring equipment and provided safe shelter for the ground crew during engine runs. During the early portion of the program, Reaction Motors used one area to test uninstalled engines, while the Air Force fired engines installed in one of the X-15s in the other area. Several "pillboxes" were also located near each area that provided shelter for other ground crews so that they could observe the operation of the engine.78

In preparation for the X-15 program, the Air Force constructed the Rocket Engine Test Facility at Edwards AFB to provide local testing of the XLR99. There were two test areas, each capable of supporting an X-15 during engine tests. For most of the flight program, the XLR99 had to be fired prior to every flight attempt, leading several engineers to complain they were testing the engines to death. Later in the program an engine could fly a second flight if no anomalies had occurred on the first. (U. S. Air Force)

In December 1959, the Air Force formally approved the XLR99 for flight in the X-15. Reaction Motors delivered a ground-test engine to Edwards at the end of May 1960, and the first flight engine at the end of July. Initially, the Air Force procured 10 flight engines, along with six spare injector-chamber assemblies. Later, the Air Force procured one additional flight engine. However, in January 1961, shortly after the first XLR99 test flight, only four engines were available to the flight program while Reaction Motors was assembling four others for delivery later in 1961. Reaction Motors continued to use four engines for ground tests, including two flight engines. Three of these engines were involved in tests to isolate and eliminate vibrations at low power levels, while the fourth investigated extending the Rokide loss that was affecting the life of the thrust chamber.-1791

In addition to ground simulators and the centrifuge, pilots and researchers used aircraft to simulate various aspects of the X-15. For instance, the Lockheed F-104 Starfighter closely approximately the wing loading of an X-15 during landing, and with the right combination of extended landing gear, flaps, and speed brakes, the F-104 at idle thrust did an excellent job of simulating the X-15. For the first 50 or so flights, the pilots dedicated an entire F-104 mission to practicing landing procedures. As new pilots entered the program, they conducted similar practices. Throughout the program, pilots used the F-104s to establish geographic checkpoints and important altitudes around the landing pattern at all the possible landing lakes.-1541

Scott Crossfield and Al White conducted similar work very early in the program using the North American YF-100A equipped with an eight-foot drag chute. Combined with extended gear and speed brakes, the F-100 at idle thrust did an adequate job of simulating the X-15 during landing, although not quite as well as the F-104. The entire process was a bit trickier since it required the in-flight deployment and release of the drag chute.1551

As Al White later remembered, "With gear down, speed brake extended, at idle power, and that drag chute deployed, the airplane was comparable to the X-15 on approach. I would start at about 25,000 feet, pick a spot on the lakebed, and see how close I could come to touching down on that spot. With all the room on the lakebed, it was not necessary to hit a spot, but it is always nice to have that much margin for error. I flew this trainer as much as I could, in preparation for that day that never came." Not flying the X-15 was one of the few disappointments during White’s significant career.-*56

Much of the X-15 flight planning took place prior to the first manned space flight. Since no one had ever left the atmosphere and returned in a winged vehicle (or anything else), there had been concern that the rapidly changing stability and control characteristics in the X-15 as it reentered the atmosphere might pose an unusually demanding piloting task. To address this question, engineers in the Flight Research Department of the Cornell Aeronautical Laboratory conceived the idea of simulating this brief (about 60 seconds duration) but unfamiliar X-15 piloting task in a NT-33A that was owned by the Air Force but operated by Cornell as a variable-stability trainer.-*57*

The NT-33A already had been equipped with a larger internal volume F-94 nose section that contained a three-axis (pitch, roll, and yaw) variable-stability and control system for in-flight simulation purposes. To support the X-15 program, Cornell modified the front cockpit to superficially resemble the X-15, with a side-stick controller on the right-hand console for atmospheric flight control and another side-stick on the left-hand console simulating the ballistic controls. An "instructor" pilot sat in the back cockpit with a normal set of T-33 controls. Jack Beilman at Cornell designed a programmable, non-linear function generator that changed the gains of 32 sensed aerodynamic and rigid-body-motion feedback variables. It also changed the flight-control sensitivities continuously during the

simulated reentry so that the NT-33A stability and control characteristics would match the predicted X-15 characteristics.-158!

The flight plan had the NT-33A entering a shallow dive at about 17,000 feet altitude and then pulling up to a ballistic trajectory that produced about 60 seconds of 0 g-about the same as the initial part of the X-15 reentry. At the same time, the variable-stability system on the NT-33A changed the flight-control sensitivities to simulate going from the vacuum of space to the rapidly increasing dynamic pressure of the atmosphere. Since the normal aerodynamic controls of the X – 15 would be ineffective outside the atmosphere, the pilot used the ballistic controller to establish the correct reentry pitch attitude.-*56

In the NT-33A simulation the "ballistic controller" produced no physical response whatsoever—it only changed the displayed pitch attitude on the instrument panel. (At this point in the simulation, the NT-33A was at 0 g.) In order to maintain the fidelity of the simulation, the X-15 pilot in the front cockpit wore a hood and had no view of the outside world, since there would be little view of the real world in the X-15 at the simulated altitudes. This deception was necessary for the high – angle-of-attack deceleration at the end of the simulated reentry because although the front cockpit instrumentation indicated the pilot was flying an unbanked steep descent (in the X-15), he was actually flying a steep 5-g turn in the NT-33A. The simulator achieved this deception by gradually biasing the attitude indicator to a bank angle of 75 degrees while the X-15 pilot used the ballistic controller to maintain wings-level flight at the proper airspeed, angle of attack, and descent rate on his cockpit instruments. It was a carefully choreographed ballet between the "student" in the front seat and the safety pilot in the back who was trying to keep the NT-33 from becoming a smoking crater in the high desert.*68!

Accordingly, a Cornell team headed by engineering test pilots Bob Harper and Nello Infanti arrived at Edwards in May 1960 to begin a series of flights in the NT-33A in order to provide reentry training for six X-15 pilots (Neil Armstrong, Jack McKay, Forrest Petersen, Bob Rushworth, Joe Walker, and Bob White). Each pilot was to receive six flights in the NT-33A that included a matrix of simulated Mach numbers, altitudes, and various control malfunctions (principally failed

dampers) both separately and simultaneously.1611 Infanti was the "instructor pilot" for each of the X-15 simulation flights in the NT-33A, and the rest of the Cornell team consisted of crew chief Howard Stevens, electronics technician Bud Stahl, and systems engineer Jack Beilman. As Beilman remembers:

During one of the flights, with Neil Armstrong in the front seat, we were simulating failed dampers at something like Mach 3.2 and 100,000 feet altitude. Neil had great difficulty with this simulated undamped X-15 configuration and lost control of the airplane repeatedly.

Nello had to recover from each one of these "lost-control" events using the controls in the back cockpit. [Infanti later recalled that some of these recoveries were "pretty sporty."] The ground crew was monitoring the test radio frequency as usual and followed these simulated flight control problems with great interest.

After landing, the NT-33A taxied to the ramp and Howard Stevens attached the ladder to the cockpits and climbed up to talk to Infanti about the airplane status. I climbed up the ladder front side to talk to Neil Armstrong. He handed me his helmet and knee-pad, got down from the cockpit and we talked about the flight and walked toward the operations building. As we arrived at the door Armstrong extended his right hand to grasp the door handle-but his hand still held the side-stick that he had broken during his last battle with the X-15 dampers-off simulation. I was unaware of any report of this incident during the flight and had not noticed the stick in Armstrong’s hand when he exited the cockpit. Addressing the matter for the first time, Armstrong said-without additional comment—"Here’s your stick!"

[It developed that Infanti had been aware of the broken side-stick after it happened because Armstrong had held it up over his head in the front cockpit for Nello to see.]

After the debriefing, we took the broken side-stick to the NASA workshop where Neil found the necessary metal tubing and repaired the stick while I mostly watched him work. The side-stick was reinstalled and ready for the first flight the next morning. Really good test pilots fix what they break!

In general, the pilots considered the NT-33 flights worthwhile, but there were some "obvious discrepancies or malfunctions" during the early flights. There were also a fair number of delays in the flights due to various system malfunctions caused by the high temperatures at Edwards. Eventually the Cornell crew corrected the malfunctions, but the X-15 pilots considered the first 10 flights unsatisfactory since they did not adequately simulate the X-15 flight profile. This was largely because the programmed trajectories required the NT-33 to fly close to its maximum capabilities: something that was not as easy as it sounds, especially in the heat over the high desert.-1621

The X-15 pilots considered the final six flights, flown during the first half of September 1960, reasonably satisfactory. In fact, the pilots discovered a novel control technique for the divergent closed-loop lateral-directional oscillation encountered at Mach 3.5 and 10 degrees angle of attack with the SAS off during these flights. By using the rudder in conjunction with the turn and bank indicator (which was, in effect, a yaw-rate meter) the pilot was able to damp the oscillations. With this technique, the ailerons were only a steady-state controller; in fact, any attempt to use the ailerons for control caused an immediate divergence. Researchers further investigated this technique on the North American fixed-base simulator with good results.1631

the X-15 flight profile somewhat more convincingly than the NT-33, making it possible to investigate new piloting techniques and control-law modifications without using an X-15. The most limiting factor was that the JF-100C was a single-seat aircraft, meaning that no safety pilot was available to lend assistance if things went wrong. To establish the X-15 flight characteristics on the JF-100C, technicians connected two portable analog computers to the airplane so that the combination became, essentially, a fixed-base simulator. One analog computer simulated the basic F-100C flight characteristics, and researchers manipulated the variable-stability gains until the motion traces matched those obtained from the North American X-15 simulator. Joe Walker and Bob White flew these pseudo fixed-base simulations until they were satisfied that the JF – 100C adequately represented the X-15.[64]

Much of the X-15 flight planning took place prior to the first manned space flight. There was concern that the rapidly changing stability and control characteristics in the X-15 as it reentered the atmosphere might pose an unusually demanding piloting task. To address this, the Cornell Aeronautical Laboratory developed a method of simulating this environment using an NT-33A operated by Cornell as a variable stability trainer. The simulations were hardly ideal, but provided much needed confidence to the original cadre of X-15 pilots. (U. S. Air Force)

The first actual flight of the JF-100C with the new mechanization was made on 24 March and was considered generally satisfactory. The major discrepancies were that the Dutch-roll and roll – subsidence modes appeared to be less stable than those of the actual X-15. Nevertheless, the JF – 100C was capable of performing some interesting simulations. For instance, six flights in late July 1961 simulated the X-15 at Mach 3.5, 84,000 feet, and 10 degrees angle of attack; later flights extended this to Mach 6 and angles of attack of 20 degrees. The aircraft returned to Ames on 11 March 1964 after making 104 flights for pilot checkout, variable-stability research, and X-15

[65]

support.

One of the tasks assigned to the JF-100C was investigating the effects of damper failure on the controllability of the X-15. Researchers had obtained the early wind-tunnel data on sideslip effects with the horizontal stabilizer at zero deflection, and used this data in the 1958 centrifuge program at Johnsville. Based on these data, reentries using an angle of attack of less than 15 degrees were possible even with the roll damper off. On the other hand, reentries at angles greater than 15 degrees (which were required for altitudes above 250,000 feet) with the roll damper off showed a distinct tendency to become uncontrollable because of a pilot-induced oscillation (PIO).[66]

As with a typical PIO, if the pilot released the control stick, the oscillations damped themselves. Nevertheless, researchers suspected that a large portion of the X-15 flight envelope was uncontrollable with the roll dampers off or failed. Investigations were initiated to find a way to alleviate the problem. The first method tried (perhaps because it would have been the easiest to implement) was pilot-display quickening. Sideslip and bank-angle presentations in the cockpit were quickened (i. e., presented with less delay) by including the yaw rate and roll rate, respectively. Researchers experimented with various quickening gains during investigation on the fixed-base simulator, but found no combination that significantly improved the pilot’s ability to handle the instability.-^

Shortly after the centrifuge program was completed, researchers conducted a wind-tunnel test to gather sideslip data with the horizontal stabilizer closer to the normal trim position (which was a large leading-edge-down deflection of -15 to -20 degrees). When researchers programmed the results of these tests into the fixed-base simulator at North American, it showed that the PIO boundary for reentry with the roll damper off had dropped from 15 degrees to only 8 degrees, adding new urgency to finding a solution.-681

To verify the magnitude of the problem in flight, several X-15 pilots explored the fringes of the expected uncontrollable region by setting the airplane up at the appropriate angle of attack and turning the roll and yaw dampers off. In each case, lateral motions began immediately. The pilots experimented with various combinations of angle of attack and control inputs in both the X-15 and the JF-100C to better define the problem.-691

Lawrence W. Taylor and Richard E. Day from the FRC, and Arthur F. Tweedie from North American independently investigated using the rolling tail to control sideslip angle during certain types of instability. An unconventional control technique, called "beta-dot," evolved from these investigations and showed considerable promise on the fixed-base simulator. This technique consisted of sharp lateral control inputs to the left as the nose swung left through zero sideslip (or vice versa to the right). The pilot kept his hands off the stick except when making the sharp lateral inputs, which eliminated the instability induced by inadvertent inputs associated with merely holding onto the center stick. However, when pilots used this technique in the JF-100C, it did not seem to work as well. Further investigations showed that it worked somewhat better in the X-15 when the pilot used the side-stick controller instead of the center stick.-701

It appeared that the beta-dot technique might allow reentries from high altitudes with the dampers failed, if anybody could figure out how to perform the maneuver successfully. As Bob Hoey, the flight planner who later discovered the ventral-off stability fix for the same problem, recalled, "the beta-dot technique is one of those things that is really difficult to explain. You could watch someone make 20 simulated reentries and still not understand what they were doing. The method was based on making a very sharp aileron pulse, timed exactly right, and totally foreign

to normal, intuitive piloting technique. Properly timed, this pulse would completely stop the rolling motion, although not necessarily at wings level. With a little finesse, you could herd the thing back to wings level flight, but, if at any time you reverted to a normal piloting technique, even for a second, you were in big trouble. Art Tweedie [who discovered this method] and Norm Cooper [a North American flight controls expert] could make successful simulator reentries with the dampers off while drinking a cup of coffee! This obviously became a big challenge for the rest of us." Hoey became pretty good at the technique himself, at least in the simulator.-171!

Dick Day later wrote that "Robert Hoey, lead Air Force engineer on the X-15 project, introduced the control technique to some of the X-15 pilots. Two pilots in particular, Major Robert White and Captain Joe Engle, became so adept at controlling ground and flight simulators that they considered the method would serve as a backup in case of roll damper failure. Fortunately, the beta-dot technique was not required because removing the ventral solved the dampers-off controllability problem. It is worth noting, however, that the complete beta-dot equation was later used in the yaw channel of the Space Shuttle control system to overcome unstable control coupling." It is another enduring legacy of the X-15 program.-721

All of the X-15 pilots trained using this technique, but the actual usefulness of the beta-dot maneuver was questionable. Furthermore, a lateral input in the wrong direction, which was conceivable considering other potential problems clamoring for the attention of the pilot, could be disastrous. One of the reasons the technique was so foreign to the pilot was that the aileron pulse had to be in the same direction as the roll, which is hardly intuitive for most pilots. Then the pilot had to remove the pulse just as the needle on the sideslip indicator hit the null mark. As Hoey remembers, "about half the pilots were dead-set against [the beta-dot maneuver] and essentially refused to consider it as an option. Others conquered the technique and actually became fairly proficient in its use on the fixed-base and in-flight simulations." Pilots flew the in-flight simulations using the NT-33 and JF-100C variable-stability airplanes, which somehow managed to survive the program.731

Researchers at Ames modified a North American JF-100C (53-1709) Super Sabre into a variable – stability trainer that could simulate the X-15 flight profile somewhat more convincingly than the NT-33, making it possible to investigate new piloting techniques and control-law modifications without using an X-15. The most limiting factor was that the JF-100C was a single-seat aircraft, meaning there was not a safety pilot to assist if things went wrong. (NASA)

There were two other answers to the PIO problem at high angles of attack. The first was to make the stability augmentation system truly redundant, at least in the roll axis, by installing the alternate stability augmentation system (ASAS); however, this took almost a year to accomplish. Another answer-discovered by Dick Day and Bob Hoey using the simulator-proved to be remarkably easy, and unexpected: remove the ventral rudder. With the lower rudder on, a considerable portion of the reentry from an altitude mission would be within the uncontrollable region should a damper fail. However, a similar reentry with the lower rudder removed would not enter the predicted uncontrollable region at all. The downside was that the pilots faced significantly reduced flying qualities at low angles of attack without the rudder. Despite a few gripes from the pilots, everybody eventually agreed to remove the lower rudder for almost all of the high-altitude missions. Only a few missions of the X-15A-2 used the ventral rudder, which in this case provided an adequate stand-in for the eventual dummy ramjet. In all, the program would make 73 flights with the ventral rudder on and 126 with it off.1741

By the time of the 1961 industry conference, researchers had determined that the fixed-base simulator and the F-104 in-flight landing pattern simulator were the two most valuable training tools available to the program. The centrifuge and variable-stability aircraft contributed to the overall pilot experience level, but were not necessary for use on a flight-by-flight basis. This mostly explains why only the first group of pilots got the thrills of "riding the wheel" at Johnsville and flying the NT-33 trainer.-1751

After three months of investigations, the Becker group believed that the development of a Mach 7 research aircraft was feasible. Those at NACA Headquarters who followed the progress of their work, as well as the parallel work on hypersonic aircraft concepts at the other NACA laboratories, agreed. It was time to formally present the results to the NACA upper echelon and the Department of Defense.-11^

The preliminary specifications for the research airplane were surprisingly brief: only four pages of requirements, plus six additional pages of supporting data. As John Becker subsequently observed, "it was obviously impossible that the proposed aircraft be in any sense an optimum hypersonic configuration." Nevertheless, Langley believed the design would work. At the same time, a new sense of urgency was present: "As the need for the exploratory data is acute because of the rapid advance of the performance of service [military] aircraft, the minimum practical and reliable airplane is required in order that the development and construction time be kept to a minimum." In other versions of the requirements, this was even more specific: "It shall be possible to design and construct the airplane within 3 years." The researchers were nothing if not ambitious.11^

On 4 May 1954, Hugh Dryden sent a letter to Lieutenant General Donald L. Putt at Air Force Headquarters stating that the NACA wanted to initiate a new manned hypersonic research aircraft program. The letter suggested a meeting between the NACA, Air Force Headquarters, and the Air Force Scientific Advisory Board to discuss the project. Putt responded favorably and recommended inviting the Navy as well. The general also noted that "the Scientific Advisory Board has done some thinking in this area and has formally recommended that the Air Force initiate action on such a program." On 11 June 1954, Dryden sent letters to the Air Force and Navy inviting them to a meeting on 9 July 1954 at NACA Headquarters.117

Attendees included Clark Millikan, Ezra Kotcher from the WADC, and a variety of Air Force and Navy technical representatives. The Air Research and Development Command (ARDC) and Air Force Headquarters also sent policy representatives. During the meeting, Hartley Soule and Walt Williams reviewed the history of previous research airplanes. Hugh Dryden reported the reasons why the NACA believed a new research aircraft was desirable, and said the time had come to determine whether an agreement existed on the objectives and scope of such a project. Dryden emphasized the need for information on full-scale structural heating and on stability and control issues at high speeds and high altitudes. He also indicated that the NACA thought that actual flight-testing combined with theoretical studies and wind-tunnel experiments produced the best results. The Langley study became the starting point for further discussions since it was the most detailed available, with John Becker and John Duberg, who was substituting for Norris Dow, leading the discussions.-118

Those in attendance were in general agreement that a new project was feasible. However, Hugh Dryden, reflecting what John Becker described as "his natural conservatism," stated that the fact it was feasible to build such a research airplane did not necessarily make it worth building; he wanted further study before deciding. The Navy representative indicated that some "military objective" should be included in the program, but Clark Millikan stressed the need for a dedicated research airplane rather than any sort of tactical prototype. The group agreed the performance parameters discussed by the Langley study represented an adequate increment over existing research airplanes, and that a cooperative program would be more cost-effective and more likely to provide better research data at an earlier time. The meeting closed with an agreement that the military would continue studying the NACA proposal, and that Hugh Dryden would seek Department of Defense approval for the project.119

Unexpectedly, the Office of Naval Research (ONR) announced at the meeting that it had already contracted with the Douglas Aircraft Company to investigate a manned vehicle capable of achieving 1,000,000 feet altitude and very high speeds. The configuration evolved by Douglas "did not constitute a detailed design proposal," but was only a "first approach to the problem of a high-altitude high-speed research airplane." Representatives from the NACA agreed to meet with their ONR counterparts on 16 July to further discuss the Douglas study.

Pressure suits, more often called "space suits" by the public, are essentially taken for granted today. Fifty years ago they were still the stuff of science fiction. These suits serve several necessary purposes, with supplying the correct partial pressure of oxygen being the most obvious (although masks or full-face helmets can also accomplish this). The most important purpose, however, is to protect the pilot against the increasingly low atmospheric pressures encountered as altitude increases—pressures that reach essentially zero above about 250,000 feet. At high altitudes, the blood and water in the human body want to boil—not from heat, but from the pressure differential between the body and the environment.-1771

A distant precursor of the full-pressure suit was, arguably, the dry suits used by turn-of-the – century commercial salvage divers, complete with their ported brass helmets and valve fittings. In 1920, renowned London physiologist Dr. John Scott Haldane apparently was the first to suggest that a suit similar to the diver’s ensemble could protect an aviator at high altitudes. There appeared, however, to be little immediate need for such a suit. The normally aspirated piston – powered airplanes of the era were incapable of achieving altitudes much in excess of 20,000 feet, and the major concern at the time was simply keeping the pilot warm. However, the increasing use of supercharged aircraft engines during the late 1920s led to the first serious studies into pressure suits. Suddenly, aircraft could fly above 30,000 feet and the concern was no longer how to keep the aviator warm, but how to protect him from the reduced pressure.-1781

During the early 1930s Mark E. Ridge determined that a suitably constructed pressurized suit would allow him to make a record-breaking altitude flight in an open balloon. His efforts to interest the United States military in this endeavor failed, and instead he contacted John Haldane in London for help. At the time, Haldane was working with Sir Robert Davis of Siebe, Gorman & Company to develop deep-sea diving suits. Together, Haldane and Davis constructed a hypobaric protection suit for Ridge. For a number of reasons, Ridge was never able to put the suit to actual use, although he tested it in a pressure chamber at simulated altitudes up to 90,000 feet.-1791

In 1934 famed aviator Wiley Post commissioned the B. F. Goodrich Company to manufacture a pressure suit of his own design. Unfortunately, the rubberized fabric suit did not work all that well. The basic design was modified by B. F. Goodrich engineer Russell Colley, and after some trial and error, Post was able to use it successfully on several record-breaking flights to altitudes of 50,000 feet.1801

While work on derivatives of the Ridge-Haldane-Davis suit continued in England, the U. S. Army Air Corps finally recognized, somewhat belatedly, the need for a pressurized protective garment for military aviators and started the classified MX-117 research program in 1939. This drew several companies into pressure-suit development, including B. F. Goodrich (with Russell Colley), Bell Aircraft, the Goodyear Rubber Company, the U. S. Rubber Company, and the National Carbon Company. From 1940 through 1943, engineers produced a number of designs that all featured transparent dome-like plastic helmets and airtight, rubberized fabric garments that greatly restricted mobility and range of motion when fully pressurized. The development of segmented, bellows-like joints at the knees, hips, and elbows improved mobility, but still resulted in an extremely clumsy and uncomfortable ensemble. The striking visual aspect of these suits resulted in their being called "tomato worm suits," after the distinctive tomato hornworm.[81]

By 1943 the Army Air Corps had largely lost interest in the concept of a full-pressure suit. The newest long-range bomber, the Boeing B-29 Superfortress, was pressurized and seemed less likely to require the suits than earlier aircraft. As Scott Crossfield later opined, "During World War II the armed services, absorbed with more vital matters, advanced the pressure suit not a whit."-82

After the war, Dr. James P. Henry of the University of Southern California began experimenting with a new concept in aircrew protection. The capstan-type partial-pressure suit operated by imposing mechanical pressure on the body directly, compressing the abdomen and limbs much like the anti-g suits then entering service. The compression was applied by inflatable bladders in the abdominal area and pneumatic tubes (capstans) running along the limbs. A tightly fitting, rubber-lined fabric hood that was fitted with a neck seal and a transparent visor fully enclosed the head.-83

In Worcester, Massachusetts, a small company named after its founder, David Clark, produced anti-g suits for the Air Force and experimental pressure suits for the Navy. Scott Crossfield described Clark as "one of the most interesting men I have ever met in the aviation world." Although Henry had approached the David Clark Company for assistance in developing his suit concept, contracts for anti-g suits between David Clark and the U. S. government made direct cooperation appear to be a conflict of interest. Instead, Clark sent materials and an experienced seamstress, Julia Greene, to help Henry continue his development in California. Just after the war, the Air Force asked Clark to observe a test of the Henry partial-pressure suit in the altitude chamber at Wright Field. Henry demonstrated the suit to a maximum altitude of 90,000 feet, and remained above 65,000 feet for more than 30 minutes; everybody was suitably impressed. The Air Force asked David Clark to produce the Henry design, and all parties soon reached an agreement that included Julia Greene returning to Worcester. David Clark produced the first suit for Jack Woolams, a Bell test pilot scheduled to fly the XS-1, and made additional suits for Chalmers "Slick" Goodlin and a little-known Air Force captain named Chuck Yeager.-84

These early partial-pressure suits did, in fact, work. On 25 August 1949, Major Frank K. "Pete" Everest was flying the first X-1 on an altitude flight when the canopy cracked and the cockpit depressurized. The laced partial-pressure suit automatically activated, squeezing Everest along the torso, arms, and legs, supporting his skin and keeping his blood from boiling. He landed, uncomfortable but unhurt. This was the first recorded use of a partial-pressure suit under emergency conditions.-1851

Continued improvements resulted in the T-1 suit, the first standardized partial-pressure suit used by the Air Force. The Air Force used the T-1 suit in a variety of aircraft, including the stripped-down "featherweight" versions of the Convair B-36 intercontinental bomber that frequently flew missions lasting in excess of 24 hours at altitudes above 50,000 feet. Unfortunately, the T-1 suit was not a particularly comfortable garment.-861

The discomfort of the so-called "Henry suit" was an unfortunate aspect of the fundamental design of partial-pressure suits. This was at least partially eliminated in the subsequent MC-1, MC-3, and MC-4 series (the MC-2 suit was an experimental full-pressure suit to be discussed later) by the placement and adjustment of panels during customized fitting. However, the suits did accomplish their main purpose: to protect the wearer from the effects of emergency decompression at altitude.-1871

Taking a different route, after the war the U. S. Navy began investigating the possibility of developing a full-pressure suit in cooperation with B. F. Goodrich and Russell Colley. This led to a progressive series of refinements of the basic design that resulted, in the early 1950s, in the first practical U. S. full-pressure suit. At the same time, the David Clark Company was also experimenting with full-pressure suits under Navy auspices. On 21 August 1953, Marine Corps Lieutenant Colonel Marion E. Carl took one of the D-558-2 aircraft to an unofficial record altitude of 83,235 feet while wearing a David Clark full-pressure suit.-1881

The Navy’s adventures in full-pressure suit development took some intriguing turns, and Scott Crossfield covers them well in his autobiography. The Navy ended up concentrating on the Goodrich designs. One of these was the Model H, an early developmental suit that the Navy considered unacceptable for operational use but showed a great deal of promise. Consequently, in a perfect example of interservice rivalry, the Air Force and Navy began separate development efforts—both based on the Model H—to perfect an operational full-pressure suit. By the early 1960s the Navy had progressed through a series of developmental models to the Mark IV, Model 3, Type 1, a production suit that Navy aircrews wore on high-altitude flights for several years.-1891

Air Force experience at high altitudes in the B-36 confirmed the need for a full-pressure suit to replace the partial-pressure suits used by the bomber crews. In response, the Air Force drafted a requirement for a suit to provide a minimum of 12 hours of protection above 55,000 feet. The goal was to construct a "fully mobile suit" that would weigh less than 30 pounds, operate with an internal pressure of 5 psi, and provide the user with sufficient oxygen partial pressure for breathing, adequate counterpressure over the body, and suitable ventilation.1901

Whatever the political nuances involved, in 1955 the Air Force issued a request for proposals for a full-pressure suit. Several contracts were awarded and the two leading designs were designated the XMC-2-ILC (International Latex Corporation) and the XMC-2-DC (David Clark Company). The ILC approach resulted in an unwieldy garment that used convoluted metal joints and metal bearing rings, and had limited mobility under pressure; it was known, however, to provide the required pressure protection. Unfortunately, the joint bearings produced painful pressure points on the body and were hazardous during bailout or ejection—hardly an ideal solution.1911

On the other hand, the David Clark suit featured a major breakthrough in suit design with the use of a new "distorted-angle fabric," called Link-Net, to control inflation and enhance range of motion. This eliminated the need for the tomato-worm bellows at the limb joints. David Clark had been developing this same basic suit with the Navy before that service opted to go with the Goodrich design. The Air Force selected the David Clark suit for further development.1921

The new Link-Net fabric was the result of an intensive effort by the company to develop a new partial-pressure suit fabric using both Navy and company money. Originally, David Clark had constructed several torso mockups using different unsupported sheet-rubber materials, but quickly discarded these when it became evident that a rupture in the material could cause the entire suit to collapse. The company began looking for a supported-rubber material that would meet the sealing requirements but would not collapse when punctured. Ultimately, David Clark selected a neoprene-coated nylon. A puncture in this material would result in a small leak, but not a sudden expulsion of gas.1931

The enormous advantages offered by the Link-Net fabric were hard to grasp. Coupled with advances in regulators and other mechanical pieces, David Clark could now produce a workable full-pressure suit that weighed about 35 pounds. Previously, during the early X-15 proposal effort, North American had estimated a suit would weigh 110 pounds.[94]

Further tests showed that two layers of nylon marquisette arranged with opposite bias provided the maximum strength in high-stress areas. This improved Link-Net material consisted of a series of parallel cords that looped each other at frequent intervals. The loops were interlocked but not connected so that the cords could slide over each other and feed from one section of the suit to another to allow the suit to deform easily as the pilot moved. The main characteristic required of the Link-Net was the lowest possible resistance to bending and twisting, but the elasticity had to be minimal since the suit could not increase appreciably in volume while under pressure. The use of a relatively non-elastic cord in the construction of Link-Net made it possible to satisfy these seemingly contradictory requirements. Clark chose nylon for the Link-Net because of its high tensile strength, low weight, and low bulk ratio.-195

The X-15 provided the first impetus to develop a workable full-pressure suit, and Scott Crossfield and Dr. David M. Clark were instrumental in the effort. The first X-15 full-pressure suit, the XMC – 2 (S794-3C) was demonstrated by Scott Crossfield in the human centrifuge at the Aero Medical Laboratory on 14 October 1957. Two 15-second runs were made at 7 g, and the following day an additional 23 tests were conducted to demonstrate the anti-g capability of the suit. (U. S. Air Force)

The first prototype David Clark Model S794 suit provided a learning experience for the company. For instance, the initial anti-g bladders were fabricated using neoprene-coated nylon, but failed during testing. New bladders incorporated a nylon-oxford restraint cover, and these passed the pressure tests. Materials evaluated for the gloves included leather/nylon, leather/nylon/Link-Net, and all leather. Eventually, the company found the best combination was leather covering the hand, a stainless-steel palm restrainer stitched inside nylon tape supported by nylon tape around the back, Link-Net from the wrist up to the top zipper, and a black cabretta top seam. However, pilots quickly found that gloves constructed in the straight position made it impossible to hold an object, such as a control stick, for more than 15-20 minutes while the glove was pressurized. When the company used a natural semi-closed position to construct the glove, the pilots could hold an object for up to 2 hours without serious discomfort. Perhaps the most surprising material used in the prototype suit was the kangaroo leather for the boots, which turned out to be soft and comfortable as well as sufficiently durable."

The construction of two "production" full-pressure suits (S794-1 and S794-2) followed. These suits were an improvement in terms of production and mobility but were, in reality, still prototypes. One of the major changes was extending the use of Link-Net material further from the joints to increase the amount of "draw" and provide additional mobility. Eventually David Clark concluded that the entire suit should use Link-Net. David Clark delivered these two suits to the Aero Medical Laboratory at Wright Field for testing and evaluation, and used the lessons learned to construct the first X-15 suit for Scott Crossfield."

Unfortunately, the reliability demonstrated during the PFRT program did not continue at Edwards. Early in the flight program, vibrations, premature chamber failures, pump seal leaks, and corrosion problems plagued operations. Potentially the most serious problem was a 1,600-cycle vibration. Fortunately, the natural frequencies of the engines dampened the vibration below 100 g. However, between 100 and 200 g, the vibration could be dampened or could become divergent, depending on a complex set of circumstances that could not be predicted in advance, and the vibration always diverged above 200 g.[80]

The vibrations caused a great deal of concern at Edwards. On 12 May 1960, as the program was trying to get ready for the first XLR99 flight, the Air Force called a meeting to discuss the problem. Although Reaction Motors had experienced only one vibration shutdown every 50 engine starts at Lake Denmark, personnel at Edwards reported that there had been eight malfunction shutdowns out of 17 attempted starts. The vibration began when the main-propellant valves opened for final chamber start, although the engines had not experienced vibrations during the igniter phase. Since the demonstrated rate of occurrence had jumped from 2% at Lake Denmark to 47% at Edwards, nobody could ignore the problem. Engineers discovered that the 1,600-cycle vibration corresponded to the engine-engine mount resonant frequency, and that Reaction Motors had not seen the vibration using the earlier non-flight-rated engine mounts at Lake Denmark. As a temporary expedient, Reaction Motors installed an accelerometer that shut the engine down when the vibration amplitude reached 120 g, a move the company believed would permit flight­testing to begin.-181

The engine (serial number 105) used at Edwards differed only slightly in configuration from those used at Lake Denmark; for example, it used an oxidizer-to-fuel ratio of 1.15:1 instead of 1.25:1. The desired operating ratio at altitude was 1.25:1, and this is what Reaction Motors had used during their tests. However, to simulate the 1.25:1 ratio on the ground, the engine had to run at 1.15:1 to compensate for atmospheric and propellant density differences at the lower altitude. Reaction Motors had tested this reduced oxidizer-to-fuel ratio only twice at Lake Denmark, and had not encountered vibrations either time. The company recommended a series of actions, including checking for purge gas leaks at the PSTS, changing the propellant ratio back to 1.25:1, and performing more engine test firings.-82

By the beginning of June 1960, the problem did not seem to be getting any better. The Air Force conducted two tests with 17 starts on engine 105 at Edwards, with two vibration shutdowns using the ground orifice (1.15:1 ratio). When engineers reinstalled the flight orifice (1.25:1 ratio), three of five starts resulted in vibration shutdowns. Reaction Motors conducted 18 starts on engine 104, and three of the four initial starts resulted in vibration shutdowns, but all restarts were successful.82

A series of minor changes made to engine 104 by Reaction Motors seemed to ease the problem, and between the middle of July and the middle of August 1960, the engine accumulated 25 starts at Edwards without any vibration-induced shutdowns. In fact, only a single malfunction shutdown of any type was experienced, which was attributed to a severe "throttle chop" that the turbopump governor could not keep up with. Other XLR99s had experienced similar problems, and Reaction Motors warned the pilots to move the throttle slowly to avoid the situation.-1841

The Propulsion System Test Stand was the unlikely name for a non-flight X-15 fuselage that was used to test rocket engines. At least two of the fuselages were manufactured, one for Reaction Motors and one for Edwards AFB. Here technicians install an XLR99 in the PSTS in preparation for a test. (NASA)

Still, as late as the meeting of the Technical Advisory Group on 9-10 November 1960, the vibration problem persisted and the Air Force launched an effort to solve the problem. This program used two engines (006 and 012) at Lake Denmark and completed a series of baseline tests by the end of November that showed a 30% incidence rate of vibration shutdowns with the flight orifices installed. Reaction Motors found that modifying the liquid-oxygen inlet substantially lowered the incident rate of vibration shutdowns. Since this modification did not seem to have any other noticeable effect on the engine, the Air Force adopted it as a temporary fix.[85]

Separately, Reaction Motors determined that o-ring deterioration at the casing joint caused fuel pump seal leaks. Replacing the o-ring was difficult because it took technicians two or three shifts to remove the turbine exhaust duct, stator blades, rotor, and inlet housing; just to remove the exhaust duct necessitated the removal and re-safety-wiring of 60 bolts. Thus, although the o – ring failure itself was not serious, since it simply resulted in a steam leak, the repair required removing the engine from the aircraft, performing a time-consuming engine disassembly, and revalidating the engine installation. This process directly contributed to early flight delays using the XLR99.[86]

Ironically, the corrosion problem appeared to be the result of the unusually long engine life. With a few exceptions, the materials used by Reaction Motors for the turbopump were compatible with the various propellants, but those in contact with the hydrogen peroxide were experiencing more corrosion than desired. There were also some instances of galvanic action between the magnesium pump case and steel parts with decomposed peroxide as an electrolyte. As one

researcher noted, "the only thing really compatible with peroxide is more peroxide." There were no obvious fixes, so the program lived with the problem.[87]

The premature failure of the thrust chambers was of more concern. To insulate the stainless-steel cooling tubes from the 5,000°F flame, Reaction Motors used a 0.005-inch-thick, flame-sprayed Nichrome®-881 undercoat with 0.010 inch of oxygen-acetylene flame-sprayed Rokide Z zirconia as an insulating, erosion-resistant top coating. In service, the Rokide coating began to spall or flake due to thermal cycling from the large number of engine starts, and from vibration effects from an unstable flame. For instance, by January 1961 about 50 square inches of Rokide coating had peeled off engine 108 at Edwards, including 14 inches during a single vibration shutdown. The loss of the coating exposed the cooling tubes to the heat and erosive effects of the flame, overheating the ammonia coolant within the tubes and reducing the amount of cooling available. The superheated ammonia vapors also attacked the stainless steel and formed a very brittle nitrided layer. At the same time, the combustion gases began to melt and erode the tube surface. As this condition continued, the effective thickness of the tube wall gradually decreased until it burst. Raw ammonia then leaked into the chamber, causing more hot spots and eventually the complete failure of the chamber.-1891

In January 1961 the X-15 Project Office and the Materials Central Division of the Aeronautical Systems Division at Wright Field initiated a study of methods to improve the chamber life of the XLR99. Two possible approaches were to attempt to improve the Rokide coating system, or to develop an improved coating. The Air Force contract with Reaction Motors already included an effort to improve the Rokide coating, but researchers expressed little faith that this would achieve any measurable results. This resulted in the Air Force initiating a program to develop an alternate coating. In the meantime, engineers at the NASA Flight Research Center (FRC) surveyed other rocket engine manufacturers to find out whether they had developed workable processes. Both Rocketdyne and Aerojet were doing extensive laboratory testing of ceramics applied with plasma – arc devices, but neither had put the process into production. Both companies indicated that their experience with flame-sprayed alumina and zirconia had been unsatisfactory. Instead, Rocketdyne was working on metal-ceramic graduated coatings, and Aerojet was investigating the use of refractory metal (molybdenum and tungsten) overcoats on top of ceramics.-1901

At the time, the Air Force already had a contract with the Plasmakote Corporation to study graduated coatings in general, and this contract was reoriented to solving the XLR99 problem specifically. A second contract, this one with the University of Dayton, was reoriented to provide realistic techniques for laboratory evaluations of the coatings.-1911

A graduated coating consisted of sprayed layers of metal and ceramic; the composition changed from 100% metal at the substrate to 100% ceramic at the top surface. This removed the traditionally weak, sensitive interface between the metal and ceramic layers. Researchers produced the coatings by spraying mixed powders with an arc-plasma jet and gradually changing the ratio of metal and ceramic powders, with most of the coatings using combinations of zirconia with Nichrome, molybdenum, or tungsten. The FRC recommended adopting the new technique immediately as a way to repair damaged chambers at Edwards. They noted that engine 101 had been patched using Rokide coating, but the engine would soon need to be repaired again since the coating was not lasting. The Air Force and NASA decided that the next patch on engine 101 would use the new process, and NASA built a special fixture at the FRC to allow the chamber of a fully assembled engine to be coated.-921

Before the new coating was applied, NASA tested an existing Rokide chamber for 5.5 minutes, and 25 square inches of Rokide coating was lost during the test. Engineers then applied a

graduated coating segmented into areas using several different top coats, including tantalum carbide, titanium carbide, titanium nitride, zirconia with 10% molybdenum, and zirconia with 1% nickel. This chamber ran for 5.75 minutes, and only 3 square inches of the new coatings were lost. However encouraging, the tests were of relatively short duration and researchers did not consider them conclusive. One thing that became apparent during the tests was that it would be extremely difficult or impossible to reclaim failed chambers if the coating wore thin or was lost, since the internal damage to the tube might be sufficient to cause it to fail with no visible damage.[93]

One of the most significant issues experienced by the XLR99 during the flight program was the premature failure of the thrust chambers. Researchers eventually traced this to the spalling or flaking of the Rokide Z zirconia coating that had been applied to the inside of the chamber as an insulator. Although improved coatings were eventually developed, the Flight Research Center also developed an in-house capability to recoat the chambers when necessary, resulting in a significant cost savings compared to sending the chambers back to Reaction Motors or procuring new chambers. (NASA)

The Technical Advisory Group met on 11-12 January 1961 at the Reaction Motors facility at Lake Denmark. All in attendance agreed that chamber durability needed to be increased, and supported the development of a quick-change orifice to simplify ground runs. The group also recommended that the X-15 Project Office initiate the procurement of six spare chambers and sufficient long – lead material to construct six more. It could not be determined whether these chambers were actually procured.-194

Some documentation indicates that the XLR99 was redesignated YLR99 on 29 December 1961, although nothing appears to have changed on the engines themselves. The original source documentation from the period is inconsistent in its use of XLR99 or YLR99; this history will use XLR99 throughout simply to avoid confusion.1951

By March 1962, technicians at the FRC had the necessary equipment and training to recoat the chambers as needed. The cost of the tooling had come to almost $10,000, but the cost to recoat a chamber was only about $2,000-much less than the cost of procuring a new chamber from Reaction Motors. The coating finally approved for use consisted of 30 mils of molybdenum primer in the throat and 10 mils elsewhere, followed by 6 mils of a graduated Nichrome-zirconia coating and then 6 mils of a zirconia topcoat. NASA used this coating process for the duration of the flight program with generally satisfactory results.1961

As is the case with almost any new technology, some things can never be fully understood. One of the harder things to grasp when dealing with complex mechanical devices is component matching (or mismatching), i. e., why some items will work in a particular assembly and other seemingly identical items will not. For example, during the initial checkout of engines 108 and 111 at Edwards, both engines exhibited excessive vibrations. NASA replaced the igniter in engine 108 with a spare that reduced the vibration to acceptable levels. The igniter that had been removed from 108 was then installed in 111 and its vibration was reduced to acceptable levels. Compatibility was not a particular problem, but scenarios such as this did point out some puzzling inconsistencies.-1971

The concept of using a large aircraft to carry a smaller one aloft was not necessarily new, but the X-1 program was the first research effort that made extensive use of the idea. The original series of X-planes used two modified Boeing B-29s and three Boeing B-50s as carrier aircraft. However, despite the fact that thousands of B-29s and B-50s had been built, by the end of 1950 maintenance personnel at Edwards were finding that it was difficult to obtain replacement parts, especially for the B-29s. The performance of the aircraft had proven adequate for the original X-1 aircraft, but as the research airplanes got heavier, the performance of even the more-powerful B – 50s became marginal. In addition, the ability to take off at high gross weights was limited in the heat that was typical of the high desert during the summer months. Obviously, the research programs needed to find a better solution.1761

B-36

Three of the four competitors had sized their X-15 concepts around the premise of using a Convair B-36 as the carrier aircraft (Douglas had chosen a B-50). Easily the largest piston – powered bomber to enter operational service, the B-36 could fly over 400 mph and some versions could climb well above 50,000 feet. Convair manufactured 385 of the giant bombers between June 1948 and August 1954. The B-36 would have carried the X-15 partially enclosed in its bomb bays, much like the X-1 and X-2 had been in earlier projects. This arrangement had several advantages, particularly that the pilot could move freely between the X-15 and B-36 during the cruise to the launch location. This was extremely advantageous if problems developed that required jettisoning the X-15 prior to launch. The B-36 was also a large aircraft with more than adequate room for a propellant top-off system (liquid oxygen and ammonia), power sources, communications equipment, breathing oxygen, and monitoring instruments and controls. Launch would have occurred at approximately Mach 0.6 at altitudes between 30,000 and 50,000 feet. At the first industry conference in 1956, engineers at North American anticipated that a B-36 would be modified beginning in the middle of 1957 and ready for flight tests in October 1958.[77]

During their proposal effort, North American evaluated four different schemes for loading the research airplane into the bomber, which were generally similar to those of the other bidders. Engineers quickly rejected the idea of using a pit (like the X-1 and operationally for the GRB – 36D/RF-84K FICON project) because of the potential "fire hazard and accumulation of fumes." Similarly, they eliminated a plan to jack up the carrier aircraft nose gear, because of "the jockeying necessary to position the research aircraft plus the precarious position of the B-36." The most complicated scheme involved physically removing the vertical stabilizer from the research airplane, sliding the X-15 under the bomber, and then reattaching the vertical once the airplane was in the bomb bay. The potential loss of structural integrity that would result from frequently removing the vertical eventually eliminated this option.-178

North American had originally selected a Convair B-36 very heavy bomber as the carrier aircraft for the X-15. However, just before modifications were to begin, NASA and the Air Force decided to replace the B-36 with a much newer Boeing B-52 Stratofortress. The B-52 was a good deal faster than the B-36, providing a better launch environment for the research airplane and reducing maintenance requirements for the ground crew. (North American Aviation)

Ramp loading, which was similar to another method used in the FICON project, became the chosen solution.-1791 Loading the X-15 into the carrier aircraft began with "running the B-36 main landing gear bogies up on permanent concrete ramps by use of commercially available electric cable hoists attached to the gear struts." The ground crew then towed the research airplane under the bomber and hoisted it into the bomb bays.1801

The X-15 was suspended from three points: one on either side of the aft fuselage attached to the rear wing spar, and a third on the centerline behind the canopy firmly supported by the structure of the forward liquid-oxygen tank bulkhead. The same types of cartridges used by tactical aircraft to jettison external fuel tanks were used to explosively separate the shackles.-1811

The only major structural modification made to the B-36 would be the removal of bulkhead no. 7, which separated bomb bays 2 and 3, along with some compensating structural stiffening.1871 The X-15 would occupy most of the three forward bomb bays. Since the B-36 used a single set of doors to cover the aft two bomb bays, shorter doors were necessary to cover only bay no. 4.1831 Interestingly, the remaining 16-foot doors covering the last bomb bay would still be functional. A small, fixed fairing replaced the doors that normally covered bomb bay nos. 1 and 2. North American proposed installing a 9-foot-diameter, 6.5-foot-long heated compartment in the front of bomb bay no. 1, equipped with its own entrance hatch on the bottom of the fuselage. The compartment could seat three crewmembers, and included oxygen and intercom connections. A 36-inch hatch opened into the bomb bay, and a catwalk on both sides of the bomb bay allowed access to the X-15 in flight. An aerodynamic fairing with a rubber-sealing strip ran the full length of the bomb-bay opening.1841

One of the more interesting suggestions concerning the carrier aircraft was that "a bank of powerful lights be turned on several minutes prior to launching so that the pilot [of the research airplane] will not be blinded by the sudden glare of daylight during launching."1851

The B-36 was equipped with a 1,000-gallon liquid-oxygen tank and a 100-gallon ammonia tank to top off the research airplane’s propellants. This was surprising because Bell and Douglas, as well as Reaction Motors, believed the rate of ammonia boil-off was so slow that no topping-off would be required. Suspended in the bomb bay above the X-15, the tanks allowed the propellants to be gravity-fed into the airplane. A nitrogen bottle pressurized and purged the tanks, and lines running outside the fuselage to the former tail turret allowed the carrier aircraft to jettison and vent the rocket propellants.1861

The "High Altitude and High Speed Study" by the El Segundo Division of the Douglas Aircraft Company had been funded by the ONR as a follow-on to the D-558 research aircraft that loosely competed with the Air Force X-1 series. Duane N. Morris led the study under the direction of the chief of the Aerodynamic Section, Kermit E. Van Every. Although the concept is generally mentioned—briefly—in most histories of the X-15, what is almost always overlooked is how insightful it was regarding many of the challenges that would be experienced by the X-15 a few years later.129

By the spring of 1954, when the X-15 approval process began, Douglas had not accomplished a detailed design for a new airplane, but recognized many of the same problems as John Becker and the researchers at Langley. The Douglas engineers also examined peripheral subjects—carrier aircraft, landing locations, etc.—that the initial Langley studies did not address in any detail.121

One interesting aspect of the Douglas Model 671 was that the contractor and the Navy had agreed that the aircraft was to have two mission profiles: high speed and high altitude (with the emphasis on the latter). This was in distinct contrast to the ongoing Langley studies that eventually led to the X-15. Although the Becker team at Langley was interested in research outside the sensible atmosphere, there was a great deal of skepticism on the part of others in the NACA and the Air Force. Douglas did not have this problem—the ONR strongly supported potential high-altitude research.

Excepting the Langley work, the Douglas study was probably the first serious attempt to define a hypersonic research airplane. Most of the other companies investigating hypersonics were oriented toward producing operational vehicles, such as the ICBMs and BoMi. Because of this, they usually concentrated on a different set of problems, frequently at the expense of a basic understanding of the challenges of hypersonic flight. The introduction from the Douglas study provides a good background:11221

The purpose of the high altitude study…is to establish the feasibility of extending human flight boundaries to extreme altitudes, and to investigate the problems connected with the design of an airplane for such flights.

The project is partially a result of man’s eternal desire to go higher, faster, or further than he did last year. Of far more importance, however, is the experience gained in the design of aircraft for high-speed, high-altitude flight, the collection of basic information on the upper atmosphere, and the evaluation of human tolerance and adaptation to the conditions of flight at extreme altitudes and speeds.

The design of an airplane for such a purpose cannot be based on standard procedures, nor necessarily even on extrapolation of present research airplane designs. Most of the major problems are entirely new, such as carrying a pilot into regions of the atmosphere where the physiological dangers are completely unknown, and providing him with a safe return to Earth. The type of flight resembles those of hypersonic, long-range, guided missiles currently under study, with all of their complications plus the additional problems of carrying a man and landing in a proper manner.

The study consists of a first approach to the design of a high-altitude airplane. It attempts to outline most of the major problems and to indicate some tentative solutions. As with any preliminary investigation into an unknown regime, it is doubtful that adequate solutions have been presented to every problem of high-altitude flight, or even that all of the problems have been considered. It would certainly appear, however, that the major difficulties are not insurmountable.

The Model 671 was 41.25 feet long (47.00 feet with the pitot boom), spanned only 18 feet with 81 square feet of area, and had an all-up weight of 22,200 pounds. In many respects, it showed an obvious family lineage to the previous D-558s. The fuselage consisted of a set of integral propellant tanks, and dive brakes were located on each side aft, as in most contemporary fighters.

A conventional configuration was deliberately chosen for the study, and no benefits have yet been discovered for any unconventional arrangement. Actually, for the prime objective of attaining very high altitudes, the general shape of the airplane is relatively unimportant. Stability and control must be provided, and it must be possible to create sufficient lift for the pullout and for landing; but, in contrast to the usual airplane design, the reduction of drag is not a critical problem and high drag is to some extent beneficial. The planform of the wing is unimportant from an aerodynamic standpoint at the higher supersonic Mach numbers. Therefore, it was possible to select the planform based on weight and structure and landing conditions. These considerations led to the choice of an essentially unswept wing of moderate taper and aspect ratio.11231

The empennage of the Model 671 was completely conventional and looked much like that of the Mach 2 D-558-2 that preceded it. However, Douglas realized that the design of the stabilizers was one of the greater unknowns of the design. "The tail surfaces are of proper size for stability at the lower supersonic Mach numbers, but there is some question of their adequacy at very high supersonic speeds. Further experimental data in this speed range are necessary before modifications are attempted. In addition, it may be possible to accept a certain amount of instability with the proper automatic servo controls." Unlike the Becker group, Douglas did not have access to a hypersonic wind tunnel.[124]

Nevertheless, preliminary investigations at Douglas indicated that "extremely large tail surfaces, approaching the wing area in size, are required to provide complete stability at the maximum Mach number of about 7." Engineers investigated several methods to improve stability, with the most obvious being to increase the size of the vertical stabilizer. However, placing additional area above the fuselage might introduce lateral directional dynamic stability problems "due to an unfavorable inclination in the principle axis of inertia and the large aerodynamic rolling moment due to sideslip (the dihedral effect)." The preferred arrangement was to add a ventral stabilizer and keep the ventral and dorsal units as symmetrical as possible. However, Douglas recognized that a large ventral stabilizer would present difficulties in ground handling and during landing. The engineers proposed that the fin should be folded on the ground, unfold after takeoff, and then be jettisoned just before touchdown. Alternately, Douglas believed that some sort of autopilot could be devised that would allow the use of more conventional-sized control surfaces.11251

Douglas conducted an evaluation of available power plants, and reached much the same conclusions the X-15 program would eventually come to. The desired engine should produce about 50,000 lbf with a propellant consumption of about 200 pounds per second. The only powerplant that met the requirements was the Reaction Motors XLR30-RM-2 rocket engine, which used liquid oxygen and anhydrous ammonia propellants. The high (245 lbf-sec/lbm) specific impulse (thrust per fuel consumption) was desirable since it provided "a maximum amount of energy for a given quantity of propellant." The high density of the propellants allowed a smaller tank size for a given propellant weight, allowing a smaller airframe. However, the researchers worried that since the original application was a missile, it would be difficult to make the engine safe enough for a manned aircraft.-11261

Douglas had some interesting observations about drag and power-to-weight ratios:11271

The function of drag in the overall performance must be reconsidered. The effect of drag is practically negligible in the power-on ascending phase of flight (for a high altitude launch), because of the very large thrust to weight ratio. Throughout the vacuum trajectory, the aerodynamic shape of the airplane is completely unimportant. During the descending phase of flight, a large drag is very beneficial in aiding in the pullout, and the highest possible drag is desired within the limits of the pilot and the structure. In fact, during the pullout it has been assumed that drag brakes would be extended in order to decelerate as soon as possible. However, because of excessive decelerative forces acting upon the pilot, it is necessary to gradually retract the brakes as denser air is entered, until they are fully retracted in the later stages of flight.

For a given propulsion unit (i. e., fixed thrust and fuel consumption), the overall performance of the present design [Model 671] is much more dependent upon the ratio of fuel weight to gross weight that it is upon the minimum drag or the optimum lift-drag ratio. Even though the fuel is expended in approximately the first 75 seconds of flight (a relatively small fraction of the total flight time), the ultimate performance as measured by the maximum altitude is affected to a great extent by small changes in the fuel to gross weight ratio. As an example, an increase in fuel weight/gross weight from 0.65 to 0.70 results in an increase in peak altitude of about 35% for a typical vertical flight trajectory, other parameters remaining

constant.

To better understand the nature of the various propellants then available for rocket engines, engineers reviewed numerous reports by the Caltech Jet Propulsion Laboratory, the NACA, and RAND. Only two oxidizers—oxygen and either red fuming or white fuming nitric acid-seemed to offer any increase in performance. Douglas was seeking better propellants than the liquid oxygen and alcohol used in the Reaction Motors LR8, effectively ruling out nitric acid since it was less dense than oxygen. The available fuels were alcohol (CH3OH or C2H5OH), anhydrous ammonia (NH3), hydrazine (N2H4), and gasoline. Alcohol offered no improvement, and hydrazine was too expensive and too difficult to handle safely, narrowing the choice to anhydrous ammonia and gasoline. Interestingly, Douglas ruled out liquid hydrogen because "on the basis of density, hydrogen is seen to be a very poor fuel." It would be 20 years before the Centaur upper stage would prove them wrong.-1128!

The Douglas Model D-671 was a proposed follow-on to the successful D-558 series of research airplanes developed under Navy auspices and flown at the High-Speed Flight Station. Preliminary investigation showed the concept was capable of roughly the same performance as the eventual X-15, but the Navy declined further development of the Douglas concept when it joined the X-15 program in late 1954. (Douglas Aircraft Company)

An auxiliary power unit (APU) rated at about 8 horsepower was necessary to support the electrical requirements of the instruments, controls, and radio. Investigation showed that the lightest alternative would be a small turbine generator using hydrogen peroxide or ethylene oxide monopropellant. The Walter Kidde Company and American Machine and Foundry Company were developing units that could satisfy the requirements. Both companies claimed they could develop a 10-horsepower hydrogen peroxide unit that weighed about 56 pounds, including propellants for 30-horsepower-minutes. Given the trouble of the future X-15 APUs, perhaps North American should have better reviewed this part of the Douglas report.-129!

to obtain reasonable estimates." They continued that "it is unfortunate that the largest contributing factor to the high temperatures of reentry, the convective heating from the boundary layer, is the one about which there is the least knowledge." Nevertheless, they took some educated guesses.[130]

The expected average heat level approached 1,400°F, with peak temperatures above 3,300°F on the wing leading edges and nose. Douglas believed "it would be impossible to design a structure for this temperature [1,400°F] which satisfies both the stress and weight requirements…." To overcome this, engineers recommended the use of some as-yet-undeveloped "good insulating material" with a density of 20 pounds per cubic foot and an insulating value of 0.20 British Thermal Units (Btu) per pound. For the purposes of the study, Douglas used a C – 110M titanium – alloy structure and skin protected by an unspecified ablative coating. Water sprayed into stainless-steel sections of the wing leading edges and nose area allowed superheated steam to remove unwanted heat, keeping these areas below their melting points. Alternately, Douglas investigated injecting cool gas (bottled oxygen) into the boundary layer to provide cooling. The study noted, however, that "none of these systems have yet been proven by practical application." The designers protected only a few areas, such as the cockpit, with batt insulation since the study assumed no heat transfer to the interior of the aircraft.-1131

Not surprisingly, Douglas chose an air-launch configuration. What is interesting is that the launch parameters were Mach 0.75 at 40,000 feet—well beyond the capabilities of anything except the Boeing B-52, which was still in the early stages of testing. Douglas summarized the need for an air launch by noting that "[t]he performance is increased, but the prime reason for the high altitude launch is the added safety which 40,000 feet of altitude gives the pilot when he takes over under his own rocket power." Trade studies conducted by Douglas indicated that an increase in launch altitude from sea level to 40,000 feet would result in a 200,000-foot increment in maximum altitude on a typical high-altitude mission. Additional benefits of a higher launch altitude diminished rapidly above 40,000 feet since most of the initial improvement was due to decreasing air density.-132

interested in the high-altitude research and at one point estimated the D-671 could reach 1,000,000 feet altitude. Although Douglas only conducted a minimal amount of research into the concept before it was cancelled, they foresaw many of the issues that would ultimately confront the X-15 development effort. (Douglas Aircraft Company)

Engineers spent a great deal of time studying possible flight paths, but "no attempt has been made in the present study to determine an absolute optimum flight path, because of the large number of variables involved." The designers noted that the airframe and propulsion systems could theoretically support a maximum altitude in excess of 1,130,000 feet; however, based on a conservative pullout altitude of 30,000 feet, the vehicle was more realistically limited to 770,000 feet. The pullout altitude (and the limiting decelerations, which were really the issue) was "directly traceable to the single limiting factor of the presence of a human pilot." The 770,000-foot, 84- degree profile resulted in a 10-g pullout maneuver, about the then-known limit of human tolerance.-1133!

Some thought was given to using a "braking thrust," which would allow a small amount of propellant to be saved and used during reentry. Either a mechanical thrust reverser would be installed on the rocket engine, or the airplane would reenter tail-first. This technique would have allowed slightly higher flights by reducing the stresses imposed by the pullout maneuver, although less propellant would be available for the ascent. The designers did not pursue this concept since entering tail-first involved undesirable risks, and the mechanical complexity of a thrust reverser seemed unnecessary, at least initially.!134!

The theoretical maximum performance was 6,150 mph and 190,000 feet for the speed profile, and 5,200 mph and 1,130,000 feet for the altitude profile (but limited, as discussed above). Landings would be made at Edwards AFB because of its "long runways and considerable latitude in the choice of direction and position of touchdown." The study noted that there would be little opportunity to control either the range or the heading by any appreciable amount after engine burnout. "Since the airplane must land without power at a specified landing site, it is obvious that it must be aimed toward the landing site at launch." Douglas estimated that a misalignment of 5 degrees in azimuth at burnout would result in a lateral miss of over 45 miles.!133!

One of the concerns expressed by Douglas was that "rocket thrust will not be sufficiently reproducible from flight to flight, either in magnitude or in alignment." Engineers estimated a thrust misalignment of less than one-half of a degree could impart 500 pounds of side force on the aircraft, causing it to go significantly off course. Researchers investigated several possible solutions to thrust misalignment, including using a larger rudder, using the auxiliary reaction control system, installing movable vanes in the exhaust,!136! performing gas separation in the nozzle,-!137! and mounting the rocket engine on a gimbal. All of these methods contained various problems or unknowns that caused the engineers to reject them. Further consideration showed that thrust misalignment was largely a non-issue since early low-speed flights would uncover any deficiencies, allowing engineers to correct them prior to beginning high-speed flights.!138!

The estimated landing speed was 213 mph, with a stall speed of 177 mph. Engineers accepted this relatively high speed "given the experimental nature of the aircraft and the high skill level of the pilots that will be flying it." The study noted that the slower speeds were possible if high-lift leading-edge devices were used or the area of the wing was increased. However, the increased weight and/or the resulting complications in the leading-edge cooling system appeared to make these changes undesirable.!139!

The high-altitude profile would use "flywheels, gyroscopes, or small auxiliary jets" for directional control outside the atmosphere, with Douglas favoring hydrogen peroxide jets in the wing tips and at the rear of the fuselage. Flywheels were rejected because they were too complex (for a three- axis system), and gyroscopes were too heavy. Each of the hydrogen peroxide thrusters would generate about 100 lbf and use 1 pound of propellant per second of operation. The engineers arbitrarily assumed that a 25-pound supply of propellant was required since no data existed on potential usage during flight. A catalyst turned the liquid hydrogen peroxide to steam at 400-psi pressure.-114^

The projected performance of the airplane caused Douglas engineers to investigate escape capsules for the pilot: "Because of the high altitude and high speed performance of the aircraft, it is believed that all ordinary bailout procedures, such as escape chutes and ejection seats, are of no value to the pilot." At the time, Douglas believed that ejection seats were only "suitable up to a Mach number of approximately one at sea level, with somewhat higher speeds being safe at higher altitudes." Instead, the engineers decided to jettison the entire forward section of the fuselage, including the pilot’s compartment, much like the Bell X-2. The total weight penalty for the capsule was about 150 pounds. The study dismissed pressure suits, stating that "it is very doubtful that sufficient pressurization equipment could be carried by the pilot during…ejection… to sustain suit pressurization from the maximum altitude to a safety zone within the earth’s lower atmosphere." Douglas stated flatly that "an ejection seat or other ordinary bailout techniques will be inadequate in view of the problem of high speeds and high altitudes." Scott Crossfield would later disagree.-1141!

In order to withstand the reentry temperatures, the cockpit windscreen used two 0.5-inch layers of quartz with a 0.25-inch vented air gap between them. This would keep the inner windscreen below 200°F. A thin sheet of treated glass placed inside the inner quartz layer reduced ultraviolet and other harmful radiation. Although the potential dangers of radiation above the atmosphere were largely unknown, Douglas predicted that little harm would come from the short flights (a few minutes) envisioned for the D-558-3. However, "proper precautions to prevent any one pilot from making too many successive flights in a weeks or months time interval should be taken….’,[142]

One of the technical innovations of the eventual X-15 program was the "ball nose" that sensed the angle of attack and angle of sideslip during high-speed and high-altitude flight. The Douglas study foresaw the need for a new pitch and yaw sensor "capable of sensing exceedingly low forces or pressures, but capable of withstanding the maximum dynamic pressures encountered during the complete pullout." However, Douglas thought that "the instrument need not be precise, for it is only to serve as a guide for pointing the nose into the wind at heights where a pilot might otherwise lose all sense of orientation." Four possible solutions emerged:[143]