Category Hypersonics Before. the Shuttle

Political Considerations

John V. Becker, arguably the father of the X-15, once stated that the project came along at “ … the most propitious of all possible times for its promotion and approval.” At the time it was not considered necessary to have a defined operational program in order to conduct basic research. There were no “glamorous and expensive” manned space projects to compete for funding, and the gen­eral feeling within the nation was one of try­ing to go faster, higher, or further. In today’s environment, as in 1968 when Becker was commenting, it is highly unlikely that a pro­gram such as the X-15 could gain approval.6

This situation should give pause to those who fund aerospace projects solely on the basis of their presumably predictable outcomes and their expected cost effectiveness. Without the X-15’s pioneering work, it is quite possible that the manned space program would have been slowed, conceivably with disastrous consequences for national prestige.7

According to Becker, proceeding with a gen­eral research configuration rather than with a prototype of a vehicle designed to achieve a specific mission as envisioned in 1954 was critical to the ultimate success the X-15 enjoyed. Had the prototype route been taken, Becker believed that “… we would have picked the wrong mission, the wrong struc­ture, the wrong aerodynamic shapes, and the wrong propulsion.” He also believed that a second vital aspect to the success of the X-15 was its ability to conduct research, albeit for very short periods of time, outside the sensi­ble atmosphere.®

The latter proved to be the most important aspect of X-15 research, given the contribu­tions it made to the space program. But in 1954 this could not have been foreseen. Few people then believed that flight into space was imminent, and most thought that flying humans into space was improbable before the next century. Fortunately, the hypersonic aspects of the proposed X-15 enjoyed “virtu­ally unanimous approval,” although ironical­ly the space-oriented results of the X-15 have been of greater value than its contributions to aeronautics.9

A final lesson from the X-15 program is that success comes at a cost. It is highly likely that researchers can never accurately predict the costs of exploring the unknown. If you under-

stand the problems well enough to accurately predict the cost, the research is not necessary. The original cost estimate for the X-15 pro­gram was $10.7 million. Actual costs were still a bargain in comparison with those for Apollo, Space Shuttle, and the International Space Station, but at $300 million, they were over almost 30 times the original estimate.10 Because the X-15’s costs were not subjected to the same scrutiny from the Administration and Congress that today’s aerospace projects undergo, the program continued. One of the risks when exploring the unknown is that you do not understand all the risks. Perhaps politi­cians and administrators should learn this par­ticular lesson from this early and highly suc­cessful program.

Rocket Engines

The XLR99 was the first large man-rated rocket engine that was capable of being throttled and restarted in flight. This com­plexity resulted in many aborted missions (approximately one-tenth of all mission aborts) and significantly added to the devel­opment cost of the engine. When the X-15 program ended, many felt that the throt­tleable feature might have been a needless luxury that complicated and delayed the development of the XLR99.

But in the mid-1960s these attributes were considered vital to the development of a rocket engine to power the Space Shuttle. At the time, Shuttle was to consist of two total­ly reusable stages—essentially a large hyper­sonic aircraft that carried a smaller winged spacecraft much like the NB-52s carried the X-15s. The same basic engine was going to power both stages; the pilots therefore need­ed to be able to control its thrust output. At some points in the early Shuttle concept development phases, the same engines would also be used on-orbit to effect changes in the orbital plane. So the original concept for the Space Shuttle Main Engines (SSME) included the ability to operate at 10 percent of their rated thrust, and to be restarted mul­tiple times during flight.11

In the end, the production SSMEs are throt­tleable within much the same range as the XLR99—65 to 109 percent, in one percent increments. In actuality about the only rou­tine use of this ability is to throttle down as the vehicle reaches the point of maximum dynamic pressure during ascent, easing stresses on the vehicle for a few seconds on each flight. Even this would not have been necessary with a different design for the solid rocket boosters.12 So the complexities required to enable the engine to throttle may, again, have been a needless luxury. Nevertheless, the development pains experi­enced by Reaction Motors provided insight for Pratt & Whitney and Rocketdyne (the two main SSME competitors) during the design and development of the SSMEs.

Human Factors

Coming at a time when serious doubts were being raised concerning man’s ability to han­dle complex tasks in the high-speed, weight­less environment of space, the X-15 became the first program for repetitive, dynamic mon­itoring of pilot heart rate, respiration, and EKG under extreme stress over a wide range of speeds and forces. The Bioastronautics Branch of the AFFTC measured unusually high heart and breathing rates on the parts of the X-15 pilots at points such as launch of the X-15 from the NB-52, engine shutdown, pull­out from reentry, and landing. Heart rates averaged 145 to 160 beats per minute with peaks on some flights of up to 185 beats per minute. Despite the high levels, which caused initial concern, these heart rates were not associated with any physical problems or loss of ability to perform piloting tasks requiring considerable precision. Consequently, theo­retical limits had to be re-evaluated, and Project Mercury as well as later space pro­grams did not have to be concerned about such high heart rates in the absence of other symptoms. In fact, the X-15’s data provided some of the confidence to go ahead with early manned Mercury flights—the downrange bal­listic shots being not entirely dissimilar to the X-15’s mission profile.’3

The bio-instrumentation developed for the X-15 program has allowed similar monitor­ing of many subsequent flight test programs. Incorporated into the pressure suit, pickups are unencumbering and compatible with air­craft electronics. The flexible, spray-on wire leads have since found use in monitoring car­diac patients in ambulances.

Another contribution of the X-15 program was the development of what John Becker calls the “first practical full-pressure suit for pilot protection in space.”14 The David Clark Company had worked with the Navy and the HSFS on an early full-pressure suit for use in high-altitude flights of the Douglas D-558- II; the suit worn by Marion Carl on his high – altitude flights was the first step. This suit was made of a waffle-weave material and had only a cloth enclosure rather than a hel­met. It should be noted that Scott Crossfield was heavily involved in the creation of this suit, the success of which Crossfield attrib­utes to “… David Clark’s genius.”15

The David Clark Company later developed the A/P-22S-2 pressure suit that permitted a higher degree of mobility.16 It consisted of a link-net material covering a rubberized pres­sure garment. Developed specifically for the X-15, the basic pressure suit provided part of the technological basis for the suits used in the Mercury and Gemini programs. It was later refined as the A/P-22S-6 suit that became the standard Air Force operational suit for high altitude flight in aircraft such as the U-2 and SR-71. However, it should be added that the space suit for Project Mercury underwent further development and was pro­duced by the B. F. Goodrich Company rather than the David Clark Company, so the line of development from X-15 to Mercury was not entirely a linear one, and security surround­ing the U-2 and Blackbird programs have obscured some of this history.17

X-15 pilots practiced in a ground-based sim­ulator that included the X-15 cockpit with all of its switches, controls, gauges, and instru­ments. An analog computer converted the pilot’s movements with the controls into instrument readings and indicated what the aircraft would do in flight to respond to con­trol actions. After a flight planner had used the simulator to lay out a flight plan, the pilot and flight planner worked “for days and weeks practicing for a particular flight.” The X-15 simulator was continually updated with data from previous flights to make it more accurate, and eventually a digital computer allowed it to perform at higher fidelity.18

Much has been made of the side-stick con­troller used on the X-15. Although the con­cept has found its way onto other aircraft, it has usually been for reasons other than those that initially drove its use on the X-15. The X-15 designers feared that the high g-loads encountered during acceleration would make it impossible for the pilot to use the conven­tional center stick; such worries are not the reason Airbus Industries has used the con­troller on the A318-series airliners. And although the side-stick controller has proven very popular in the F-16 fighter, it has not been widely adopted. Nevertheless, the X-15 experience provided a wealth of data over a wide range of flight regimes.

Some phases of X-15 flight, such as reentry, were marginally stable, and the aircraft required artificial augmentation (damping) systems to achieve satisfactory stability. The X-15 necessi­tated the development of an early stability aug­mentation system (SAS). The first two X-15s were equipped with a simple fail-safe, fixed – gain system. The X-15-3 was equipped with a triple-redundant adaptive flight control system; the pilot flew via inputs to the augmentation system. Although a point of continuing debate, the X-15 did not incorporate a “fly-by-wire” system if meant to denote a nonmechanically linked control system. Nevertheless, the SAS system did “fly” the X-15-3 based on pilot input rather than the pilot flying it direcdy. This basic concept would find use on an entire generation of aircraft, including such high performance fighters as the F-15. The advent of true fly-by­wire aircraft, such as the F/A-18, would advance the concept even further.


In 1954, the few existing hypersonic wind tunnels were small and presumably unable to simulate the conditions of actual flight at speeds above Mach 5, The realistic fear at the time was that testing in them would fail to produce valid data. The X-15 provided the earliest, and so far most significant, valida­tion of hypersonic wind tunnel data. This was of particular significance since it would be extremely difficult and very expensive to build a large-scale hypersonic wind tunnel.

This general validation, although broadly con­firmed by other missiles and spacecraft, came primarily from the X-15; it made the conven­tional, low-temperature, hypersonic wind tun­nel an accepted source of data for configura­tion development of hypersonic vehicles.

The X-15 program offered an excellent opportunity to compare actual flight data with theory and wind tunnel predictions. The X-15 verified existing wind tunnel tech­niques for approximating interference effects for high-Mach, high angle-of-attack hyper­sonic flight, thus giving increased confi­dence in small-scale techniques for hyper­sonic design studies. Wind tunnel drag meas­urements were also validated, except for a 15 percent discrepancy found in base drag— caused by the sting support used in the wind tunnel. All of this greatly increased the con­fidence of engineers as they set about design­ing the Space Shuttle.

One of the widely held beliefs in the mid – 1950s was the theoretical presumption that the boundary layer (the thin layer of air close to the surface of an aircraft) would be highly stable at hypersonic speeds because of heat flow away from it. This presumption fostered the belief that hypersonic aircraft would enjoy laminar (smooth) airflow over their surfaces. At Mach 6, even wind tunnel extrapolations indicated extensive laminar flow. However, flight data from the X-15 showed that only the leading edges exhibited laminar flow and that turbulent flow occurred over most surfaces. Small surface irregularities, which produced turbulent flow at transonic and supersonic speeds, also did so at Mach 6.20 Thus, engineers had to aban­don their hopeful expectations. Importantly, X-15 flight test data indicated that hyperson­ic flow phenomena were linear above Mach 5, allowing increased confidence during design of the Space Shuttle, which must rou­tinely transition through Mach 25 on its way to and from space. The basic X-15 data were also very useful to the NASP designers while that program was viable.

In a major discovery, the Sommer-Short and Eckert T-prime aerodynamic heating predic­tion theories in use during the late 1950s were found to be 30 to 40 percent in excess of flight test results. Most specialists in fluid mechanics refused to believe the data, but repeated in-flight measurements completely substantiated the initial findings. This led the aerodynamicists to undertake renewed ground-based research to complete their understanding of the phenomena involved— highlighting the value of flight research in doing what Hugh Dryden had predicted for the X-15 in 1956: that it would “separate the real from the imagined.”21

Subsequent wind tunnel testing led to Langley’s adopting the empirical Spaulding – Chi model for hypersonic heating. This eventually allowed the design of lighter vehi­cles with less thermal protection that could more easily be launched into space. The Spaulding-Chi model found its first major use during the design of the Apollo com­mand and service modules and proved to be quite accurate. In 1999 the Spaulding-Chi model was still the primary tool in use.

Based on their X-15 experience, North American devised a computerized mathe­matical model for aerodynamic heating called HASTE (Hypersonic and Supersonic Thermal Evaluation) which gave a workable “first cut” approximation for design studies. HASTE was, for example, used directly in the initial Apollo design study. Subsequent

versions of this basic model were also used early in the Space Shuttle design evolution.

At the time of the first Mach 5 X-15 flight, perhaps its greatest contribution to aeronau­tics was to disprove the existence of a “sta­bility barrier" to hypersonic flight that was suspected after earlier research aircraft encountered extreme instability at high supersonic speeds. Although of little conse­quence today, the development of the “wedge” tail allowed the X-15 to successful­ly fly above Mach 5 without the instability that had plagued the X-l series and X-2 air­craft at much lower speeds. The advent of modem fly-by-wire controls and stability augmentation systems based around high speed digital computers have allowed designers to compensate for gross instabili­ties in basic aerodynamic design, and even to tailor an aircraft’s behavior differently for different flight regimes. The era of building a vehicle that is dynamically stable has passed, and with it much of this lesson.

The art of simulation grew with the X-15 pro­gram, not only for pilot training and mission rehearsal, but for research into controllability problems. The same fixed-based simulator used by the pilots could also be used to explore those areas of the flight envelope deemed too risky for actual flight. The X-15 program showed the value of combining wind tunnel testing and simulation in maximizing the knowledge gained from each of the 199 test flights. It also provided a means of com­paring “real” flight data with wind tunnel data. It is interesting to note that the man-in – the-loop simulation first used on X-15 found wide application on the X-30 and the X-33. In fact, DFRC research pilot Stephen D. Ishmael has flown hundreds of hours “in” the X-33, which ironically is an unpiloted vehicle.

Flight Research and Space Flight

Before the X-15, high-speed research air­craft flown at Edwards could be monitored and tracked from Edwards. The trajectory of the X-15 extended much farther from

Edwards than those of the previous research aircraft, requiring two up-range stations where tracking, communications, and telemetry equipment were installed and inte­grated with the control room back at the FRC. Along the X-15 flight route, program personnel also surveyed a series of dry lakebeds for emergency landings and tested them before each flight to ensure they were hard enough to permit the X-15 to land.22 In many ways this parallels the tracking and communications network and the transat­lantic abort sites used by the Space Shuttle.

The opportunity to observe pilot perform­ance under high stress and high g-forces indicated that an extensive ground training program was needed to prepare pilots to han­dle the complex tasks and mission profiles of space flight. The result was a simulation pro­gram that became the foundation for crew training for all human space flight. The pro­gram depended on four types of simulation.

Prior to the first X-15 mission, the abil­ity of the pilot to function under the high g-forces expected during launch and reentry was tested in a closed-loop, six – degree-of-freedom centrifuge at the Naval Air Development Center, Johnsville, Pennsylvania. This project became the prototype for programs set up at the Ames Research Center and the Manned Spacecraft Center at Houston (now the Johnson Space Center).23

A static cockpit mockup provided the means for extensive mission rehearsal— averaging 20 hours per 10 minute flight. Such preparation was directly responsi­ble for the high degree of mission success achieved as pilots rehearsed their pri­mary, alternate, and emergency mission profiles. Similar, but much more elabo­rate, rehearsals are still used by astro­nauts preparing for Space Shuttle flights.

X-15 pilots maintained proficiency by flying an NT-33 or JF-100C variable – stability aircraft whose handling charac-

Подпись: A great deal of what was learned on X-15 went on to build Space Shuttle. (NASA)
Flight Research and Space Flight

teristics could be varied in flight, simu­lating the varied response of the X-15 traversing a wide range of velocities and atmospheric densities. Much of this training is now conducted in advanced motion-based simulators, although the Air Force still operates a variable-stabil­ity aircraft (the VISTA F-16).

Pilots practiced the approach and land­ing maneuver in F-104 aircraft. With landing gear and speed brakes extended, the F-104’s power-off glide ratio approximated that of the unpowered X-15. Shuttle crews continue this same practice using modified Gulfstream Shuttle Training Aircraft (STA).

Astronaut “capsule communicators,” (cap – comms) were a direct outgrowth of the X-15’s practice of using an experienced pilot as the ground communicator for most X-15 mis­sions.24 This practice existed through Mercury, Gemini, and Apollo, and continues today on Space Shuttle missions. It is still believed that a pilot on the ground makes the best person to communicate with the crew, especially in stressful or emergency situations.25

Subsequent flight test work at Edwards relied heavily on the methodology developed
for the X-15. There are no fewer than three high-tech control facilities located at Edwards today; the facility at Dry den, the Riddley Control Center complex at the AFFTC, and the B-2 control complex locat­ed on South Base. Each of these control cen­ters has multiple control rooms for use dur­ing flight test. The X-33 program has built yet another control room, this one located near the launch site at Haystack Butte.26

The X-15 program required a tracking net­work known as “High Range.” Operational techniques were established for real-time flight monitoring which were carried over to the space program. The experience of setting up this control network became something of a legacy to Mercury and later space projects through the personnel involved. Gerald M. Truszynski, as Chief of the Instrumentation Division at the FRC, had participated in set­ting up the High Range, as had Edmond C. Buckley, who headed the Instrument Research Division at Langley, The Tracking and Ground Instrumentation Group at Langley had the responsibility for tracking the Mercury capsules, and it was headed, briefly, by Buckley.27

Buckley soon transferred to NASA Headquarters as assistant director for space
flight operations, with Truszynski joining him in 1960 as an operations engineer. Both con­tinued to be involved in instrumentation and communication until a reorganization under NASA Administrator James Webb created an Office of Tracking and Data Acquisition with Buckley as director. Buckley named Truszynski as his deputy, and in 1962 appointed him to lead the Apollo Task Group that shaped the Apollo tracking and data net – work. JK Much of this same infrastructure was used early in the Space Shuttle program.

Meanwhile, Walter Williams, who had headed the NACA operations at the HSFS/FRC since 1946, was reassigned as Associate Director of the newly formed Space Task Group at Langley in September 1959. He eventually served as the Director of Operations for Mercury, and then as Associate Director of the Manned Spacecraft Center. He also served as operations director in the Mercury Control Center at Cape Canaveral during the Mercury flights of Alan Shepard, Gus Grissom, and John Glenn in 1961 and 1962.2’

Experience from the NASA 1 control room undoubtedly influenced the development of the Mercury Control Center at Cape Canaveral, and perhaps more distantly, even the Mission Control Center (MCC)30 at Houston.31 However, the spacecraft control rooms and their tracking and data acquisition systems drew on many other sources (includ­ing the missile ranges which they shared),32 although the experience setting up the High Range and operating the NASA 1 control room undoubtedly provided some opera­tional perspectives.

An often overlooked area where the X-15 influenced Space Shuttle operations is in the energy management maneuvers immediately prior to landing. By demonstrating that it was possible to make precision unpowered landings with vehicles having a low lift – over-drag ratio, the X-15 program smoothed the path for the slightly later lifting-body program and then for the space shuttle pro­cedures for energy management and landing.

The techniques used by X-15 pilots consist­ed of arriving at a “high key" above the intended landing point. Once he reached the high key, the pilot did not usually need or receive additional information from the con­trol room; he could complete the landing using visual information and his own experi­ence with practice landings in an F-104 con­figured to simulate an X-15 landing. With considerable variation on different missions, the pilot would arrive at the high key on an altitude mission at about 35,000 feet, turn 180 degrees and proceed to a “low key" at about 18,000 feet, where he would turn another 180 degrees and proceed to a landing on Rogers Dry Lake. Depending upon the amount of energy remaining, the pilot could use shallow or tight bank angles and speed brakes as necessary.

Because of their much higher energy, the standard approach for the Space Shuttle con­sists of a variation on this 360-degree approach. As a Shuttle approaches the run­way for landing, if it has excess energy for a normal approach and landing, it dissipates this energy in S-turns (banking turns) until it can slow to a subsonic velocity at about

49,0 feet of altitude some 25 miles from the runway. It then begins the approach and landing phase at about 10,000 feet and an equivalent airspeed of about 320 mph some 8 miles from the runway.33 Early in the Space Shuttle program, a specially-configured T-3834 would accompany the orbiter on the final approach, much as the X-15 chase air­craft did at Edwards. Shuttle pilots practice in a specially-configured Gulfstream Shuttle Training Aircraft, much as the X-15 pilots did in the modified F-104.

Introduction and Author’s Comments

Подпись: Dennis R. Jenkins is an aerospace engineer who spent almost 20 years on the Space Shuttle program for various contractors, and has also spent time on other projects such as the X-33 technology demonstrator. He is also an author who has written over 20 books on aero-space history.

It is a beginning. Over forty-five years have elapsed since the X-15 was conceived; 40 since it first flew. And 31 since the program ended. Although it is usually heralded as the most productive flight research program ever undertaken, no serious history has been assembled to capture its design, develop­ment, operations, and lessons. This mono­graph is the first step towards that history.

Not that a great deal has not previously been written about the X-15, because it has. But most of it has been limited to specific aspects of the program; pilot’s stories, experiments, lessons-leamed, etc. But with the exception of Robert S. Houston’s history published by the Wright Air Development Center in 1958, and later included in the Air Force History Office’s Hypersonic Revolution, no one has attempted to tell the entire story. And the WADC history is taken entirely from the Air Force perspective, with small mention of the other contributors.

In 1954 the X-l series had just broken Mach 2.5. The aircraft that would become the X-15 was being designed to attain Mach 6, and to fly at the edges of space. It would be accom­plished without the use of digital computers, video teleconferencing, the internet, or email. It would, however, come at a terrible financial cost—over 30 times the original estimate.

The X-15 would ultimately exceed all of its original performance goals. Instead of Mach 6 and 250,000 feet, the program would record Mach 6.7 and 354,200 feet. And com­pared against other research (and even oper­ational) aircraft of the era, the X-15 was remarkably safe. Several pilots would get banged up; Jack McKay seriously so, although he would return from his injuries to
fly 22 more X-15 flights. Tragically, Major Michael J. Adams would be killed on Flight 191, the only fatality of the program.

Unfortunately due to the absence of a subse­quent hypersonic mission, aeronautical applications of X-15 technology have been few. Given the major advances in materials and computer technology in the 30 years since the end of the flight research program, it is unlikely that many of the actual hard­ware lessons are still applicable. That being said, the lessons learned from hypersonic modeling, simulation, and the insight gained by being able to evaluate actual X-15 flight research against wind tunnel and predicted results, greatly expanded the confidence of researchers. This allowed the development of Space Shuttle to proceed much smoother than would otherwise have been possible.

In space, however, the X-15 contributed to both Apollo and Space Shuttle. It is interest­ing to note that when the X-15 was con­ceived, there were many that believed its space-oriented aspects should be removed from the program since human space travel was postulated to be many decades in the future. Perhaps the major contribution was the final elimination of a spray-on ablator as a possible thermal protection system for Space Shuttle. This would likely have hap­pened in any case as the ceramic tiles and metal shingles were further developed, but the operational problems encountered with the (admittedly brief) experience on X-15A-2 hastened the departure of the ablators.

Many people assisted in the preparation of this monograph. First and foremost are Betty Love, Dill Hunley, and Pete Merlin at the DFRC History Office. Part of this project

was assembling a detailed flight log (not part of this monograph), and Betty spent many long hours checking my data and researching to fill holes. I am terribly indebted to her. Correspondence continues with several of the program principals—John V. Becker, Scott Crossfield, Pete Knight, and William Dana. Dr. Roger Launius and Steve Garber at the NASA History Office, and Dr. Richard Hallion, Fred Johnsen, Diana Comelisse,
and Jack Weber all provided excellent sup­port for the project. A. J. Lutz and Ray Wagner at the San Diego Aerospace Museum archives, Tony Landis, Brian Lockett, Jay Miller, and Terry Panopalis also provided tremendous assistance to the project.

Dennis R. Jenkins Cape Canaveral, Florida February 2000

With the XLR99 engine lagging behind in its development schedule, the X-15 program decided to press ahead with ini­tial flights using two XLR11 engines—the same basic engine that had powered the Bell X-1 on its first supersonic flight. (San Diego Aerospace Museum Collection)

Introduction and Author’s CommentsIntroduction and Author’s CommentsWhen the Reaction Motors XLR99 engine finally became avail­able, the X-15 began setting records that would stand until the advent of the Space Shuttle. Unlike the XLR11, which was “throttleable” by ignit­ing different numbers of thrust chambers, the XLR99 was a truly throttleable engine that could tailor its output for each specif­ic mission. (San Diego Aerospace Museum Collection)

Introduction and Author’s CommentsIntroduction and Author’s CommentsHydraulic lifts were installed in the ramp at the Flight Research Center {now the Dryden Flight Research Center) to lift the X-15 up to the wing pylon on the NB-52 mothership. (Jay Miller Collection)

The early test flights were conducted with a long air data probe protruding from the nose of the X-15. Notice the technician manually retracting the nose landing gear on the X-15, some­thing accomplished after the research air­plane was firmly con­nected to the wing of the NB-52 mothership. (San Diego Aerospace Museum Collection)

Components and Construction

The X-15 was designed with a hot-structure that could absorb the heat generated by its short-duration flight. Remember, the X-15 seldom flew for over ten minutes at a time, and a much shorter time was spent at the maximum speed or dynamic pressure. Development showed the validity of ground

“partial simulation” testing of primary mem­bers stressed under high temperature. A facility was later built at DFRC for heat – stress testing of the entire structure, and sim­ilar testing was accomplished on the YF-12A

after the X-15. Lockheed also used all-mov­ing surfaces on the Blackbird series of Mach 3 aircraft, although it is difficult to ascertain if the X-15 influenced this design choice.

Blackbird and the Space Shuttle structural test article (STA-099).35

The X-15 pioneered the use of corrugations and beading to relieve thermal expansion stresses. Metals with dissimilar expansion coefficients were also used to alleviate stress­es, and the leading edges were segmented to allow for expansion. Around the same time, similar techniques were apparently developed independently by Lockheed for use on Blackbird series of Mach 3+ aircraft.

The X-15 represented the first large-scale use of Inconel X, in addition to extensive use of titanium alloys. This required the develop­ment of new techniques for forming, milling, drilling, and welding that came to be widely

The X-15 designers also had to solve prob­lems relating to high aerodynamic heating in proximity to cryogenic liquids. This led to cryogenic tubing that was used on parts of the Apollo spacecraft, and thermal insulation design features that were later used on the Space Shuttle. An early experience of run­ning a liquid nitrogen cooling line too close to a hydraulic line taught designers about the need to fully understand the nature of the flu­ids they were dealing with. In-flight failures on high altitude flights with the X-15 also taught aerospace engineers about such things as the need to pressurize gear boxes on aux­iliary power units to prevent foaming of the lubricant in the low pressure of space, since that led to material failures.36

used in the aerospace industry. North American pioneered chemical milling, a construction technique that has since been used on other projects.

The differentially deflected horizontal stabi­lizers on the X-15 provided roll and pitch control and allowed designers to eliminate the ailerons that would have provided a severe structural and theromodynamic prob­lem within the thin wing section used on the X-15. This configuration was already being flight tested by less exotic aircraft (YF-107A) at the same time it was used on the X-15, but nevertheless proved extremely valuable. It is common practice today to use differential stabilators on modem aircraft, particularly fighters, although in most cases conventional ailerons are also retained; the flight control system deciding when to use which control surfaces based on conditions.

The all-moving vertical surfaces in lieu of

Although the primary structure of the X-15 proved sound, several detailed design prob­lems were uncovered during early flight tests. A surprise lesson came with the discovery of heretofore unconsidered local heating phe­nomena. Small slots in the wing leading edge, the abrupt contour change along the canopy, and the wing root caused flow disruptions that produced excessive heating and adjacent material failure. The X-15, tested in “typical” panels or sections, demonstrated the problems encountered when those sections are joined and thus precipitated an analytical program designed to predict such local heating stress­es. From this experience, Rockwell engineers closely scrutinized the segmented carbon-car­bon composite leading edge of the Space Shuttle wing. The bimetallic “floating retain­er” concept designed to dissipate stresses across the X-15’s windshield carried over to the Apollo and Space Shuttle windshield designs as well.

conventional rudders has proven somewhat less attractive to aircraft designers. North American used an all-moving vertical sur­face on the A-5 Vigilante, designed not long

On three occasions, excessive aerodynamic heating of the nose-wheel door scoop caused structural deformation, permitting hot bound­ary-layer air to flow into the wheel well,

Monographs in Aerospace History Number 18 — Hypersonics Before the Shuttle


damaging the landing gear, and in one case causing the gear to extend at Mach 4.2 (flight 2-33-56). Although the landing gear remained intact, the disintegration of the tires made the landing very rough. The need for very careful examination of all seals became apparent, and closer scrutiny of surface irreg­ularities, small cracks, and areas of flow interaction became routine. The lessons learned from this influenced the final detailed design of the Space Shuttle to ensure that gaps and panel lines were adequately protect­ed against inadvertent airflow entry.

Other problems from aerodynamic heating included windshield crazing, panel flutter, and skin buckling. Arguably, designers could have prevented these problems through more extensive ground testing and analysis, but a key purpose of flight research is to discover the unexpected. The truly significant lesson from these problems is that defect in subson­ic or supersonic aircraft that are compara­tively minor at slower speeds become much more critical at hypersonic speeds.37

One of the primary concerns during the X-15 development was panel flutter, evidenced by the closing paper presented at the 1956 industry conference. Panel flutter has proven difficult to predict at each speed increment throughout history, and the hypersonic regime was no different. Although the X-15 was conservatively designed, and incorporat­ed all the lessons from first generation super­sonic aircraft, the fuselage side tunnels and the vertical surfaces were prone to develop panel flutter during flight. This led to an industry-wide reevaluation of panel flutter design criteria in 1961-62. Stiffeners and reduced panel sizes alleviated the problems on the X-15, and similar techniques later found general application in the high speed aircraft of the 1960s.’■ The lessons learned at Mach 6 defined criteria later used in the development of the Space Shuttle.

The X-15 provided the first opportunity to study the effects of acoustical fatigue over a wide range of Mach numbers and dynamic pressures. In these first in-flight measure­ments, “boundary layer noise”-related stress­es were found to be a function of g-force, not Mach number. Such fatigue was determined to be no great problem for a structure stressed to normal in-flight loading. This knowledge has allowed for more optimum structural design of missiles and space cap­sules that experience high velocities.

On the X-15, the measurement of velocity was handled by early inertial systems. All three X-15s were initially equipped with analog-type systems which proved to be highly unreliable. Later, two aircraft, includ­ing the X-15-3 with the adaptive control sys­tem, were modified with digital systems. In the subsequent parallel evaluation of analog versus digital inertial systems, the latter was found to be far superior. It was far more flex­ible and could make direct inputs to the adaptive flight control system; it was also subject to less error. Thanks to advances in technology such as laser-ring gyros and dig­ital computers, inertial systems have become inexpensive, highly accurate, and very reli­able.3* In recent years they have been inte­grated with the Global Positioning System (GPS), providing three-dimensional attitude and position information.

During the early test flights, the X-15 relied on simple pilot-static pressure instruments mounted on a typical flight test nose boom. These were not capable of functioning as speeds and altitudes increased. To provide attitude information, the NACA developed the null-sensing “ball-nose” which could survive the thermal environment of the X-15. An extendable pitot tube was added when the velocity envelope was expanded beyond Mach 6. Thus far the ball-nose has not found subsequent application, and probably never will since inertial and GPS systems have evolved so quickly. Interestingly, the Space Shuttle still uses an extendable pitot probe during the landing phase.

The X-15 was the first vehicle to routinely use reaction controls. The HSFS had begun

research on reaction controls in the mid – 1950s using a fixed-base analog control stick with a pilot presentation to determine the effects of control inputs. This was followed by a mechanical simulator to enable the pilot to experience the motions created by reac­tion controls. This device emulated the iner­tial ratios of the X-1B, which incorporated a reaction control system using hydrogen-per­oxide as a monopropellant, decomposed by passing it through a silver-screen catalyst. Because of fatigue cracks later found in the fuel tank of the X-1B, it completed only three flights using the reaction control sys­tem before it was retired in 1958.44

As a result, a JF-104A with a somewhat more refined reaction control system was tested beginning in late 1959 and extending into 1961. The JF-104A flew a zoom-climb maneuver to achieve low dynamic pressures at about 80,000 feet that simulated those at higher altitudes. The techniques for using reaction controls on the X-15, and more importantly, for transferring from aerody­namic controls to reaction controls and back to aerodynamic controls provided a legacy to the space program.41

The X-15-3 was equipped with a Minneapolis Honeywell MH-96 self-adaptive control sys­tem designed for the cancelled Dyna-Soar. The other two X-15s had one controller on the right-hand side of the cockpit for aerodynamic controls and another on the left-hand side for the reaction controls. Thus, the pilot had to use both hands for control during the transition from flying in the atmosphere to flying outside the atmosphere and then back in the opposite direction. Since there was no static stability outside the atmosphere, the pilot had to count­er any induced aircraft motion manually using the reaction controls. The MH-96 had an atti­tude hold feature that maintained the desired attitude except during control inputs. The MH-96 also integrated the aerodynamic and reaction controls in a single controller, greatly improving handling qualities during the transi­tion from aerodynamic to space flight, as well as reducing pilot workload.42

But the basic feature of the MH-96 was auto­matic adjustment of gain (sensitivity) to maintain a desirable dynamic response of the airplane. The MH-96 compared the actual response of the airplane with a preconceived ideal response in terms of yaw, pitch, and roll rates. Initially, Milt Thompson stated that the system was “somewhat unnerving to the pilot” because he was not in “direct control of the aircraft” but was only “commanding a computer that then responded with its own idea of what is necessary in terms of a con­trol output.” However, pilots became “enthu­siastic in their acceptance of it” when they realized that the MH-96 resulted in “more precise command than was possible” with the reaction controls by themselves. Consequently, the X-15-3 with the MH-96 was used for all altitude flights planned above 270,000 feet.4’

There were some problems with the experi­mental system, including one that con­tributed to the death of Mike Adams in X-15- 3 on 15 November 1967. Nevertheless, the MH-96 constituted a significant advance in technology that helped pave the way toward fly-by-wire in the early 1970s. Today, most every aircraft, and several automobiles, fea­ture some variation of a fly-by-wire system with automatic rate-gain adjustment and sta­bility augmentation.44

The Genesis of a Research Airplane

It was not until the mid-1940s that it became apparent to aerodynamic researchers in the United States that it might be possible to build a flight vehicle capable of hypersonic speeds. Until that time, propulsion systems capable of generating the thrust required for such vehi­cles had simply not been considered techni­cally feasible. The large rocket engines that had been developed in Germany during World War II allowed concept studies to be initiated with some hope of success.

Nevertheless, in the immediate post-war peri­od, most researchers believed that hypersonic flight was a domain for unmanned missiles. When an English translation of a technical paper by German scientists Eugen Sanger and Irene Bredt was provided by the U. S. Navy’s Bureau of Aeronautics (BuAer) in 1946, this preconception began to change. Expanding upon ideas conceived as early as 1928, Singer and Bredt had concluded during 1944 that a rocket-powered hypersonic aircraft could be built with only minor advances in technology. The concept of manned aircraft flying at hypersonic speeds was highly stimulating to researchers at the National Advisory Committee for Aeronautics (NACA).’ But although there were numerous paper studies exploring variations of the Sanger and Bredt proposal in the late 1940s, none bore fruit and no hardware construction was undertaken at that time. It was from this background, how­ever, that the concept for a hypersonic research airplane would emerge.2

At the time, there was no established need for a hypersonic aircraft, and it was assumed by many that no operational military3 or civil requirement for hypersonic vehicles would be forthcoming in the foreseeable future. The need for hypersonic research was not over­whelming, but there was a growing body of opinion that it should be undertaken.

The first substantial official support for hyper­sonic research came on 24 June 1952 when the NACA Committee on Aerodynamics passed a resolution to “… increase its program dealing with the problems of unmanned and manned flight in the upper stratosphere at altitudes between 12 and 50 miles,4 and at Mach num­bers between 4 and 10.” This resolution was ratified by the NACA Executive Committee when it met the following month. A study group consisting of Clinton E. Brown (chair­man), William J. O’Sullivan, Jr., and Charles H. Zimmerman was formed on 8 September 1952 at the Langley5 Aeronautical Laboratory. This group endorsed the feasibility of hyper­sonic flight and identified structural heating as the single most important technological prob­lem remaining to be solved.

An October 1953 meeting of the Air Force’s Scientific Advisory Board (SAB) Aircraft Panel provided additional support for hyper­sonic research. Chairman Clarke Millikan released a statement declaring that the feasi­bility of an advanced manned research aircraft “should be looked into.” The panel member from Langley, Robert R. Gilruth, played an important role in coordinating a consensus of opinion between the SAB and the NACA.

Contrary to Sanger’s conclusions, by 1954 it was generally agreed within the NACA and industry that the potential of hypersonic flight could not be realized without major advances in technology. In particular, the unprecedent­ed problems of aerodynamic heating and high-temperature structures appeared to be so formidable that they were viewed as “barriers” to sustained hypersonic flight.

Fortunately, the successes enjoyed by the sec­ond generation X-ls and other high-speed research programs had increased political and philosophical support for a more advanced research aircraft program. The large rocket engines being developed by the long-range missile (ICBM) programs were seen as a way to provide power for a hypersonic research vehicle. It was now agreed that manned hypersonic flight was feasible. 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 with which to compete for funding. The time was finally right for launch­ing a hypersonic flight research program. s

The specific origins of the hypersonic research program occurred during a meeting of the NACA inter-laboratory Research Airplane Panel held in Washington, DC, on 4­5 February 1954. The panel chairman, Hartley A. Soule, had directed NACA research air­craft activities in the cooperative USAF – NACA program since 1946 and was well versed in the politics and personalities involved. The panel concluded that a wholly new manned research vehicle was needed, and recommended that NACA Headquarters request detailed goals and requirements for such a vehicle from the research laboratories.

In responding to the NACA Headquarters, all of the NACA laboratories set up small ad hoc study groups during March 1954. Langley had been an island of hypersonic study since the end of the war and chose to deal with the problem in more depth than the other labora­tories. After the new 11-inch hypersonic wind tunnel at Langley became operational in 1947, a research group headed by Charles H. McLellan was formed to conduct limited hypersonic research.7 This group, which reported to the Chief of the Langley Aero – Physics Division, John V. Becker, provided verification of newly developed hypersonic theories while investigating such important phenomena as hypersonic shock-boundary – layer interaction. The 11-inch tunnel later served to test preliminary design configura­tions that led to the final hypersonic aircraft configuration. Langley also organized a paral­lel exploratory program into materials and structures optimized for hypersonic flight.

Given this, it was not surprising that a team at Langley was largely responsible for defining the early requirements for the new research airplane. The members of the Langley team included Maxim A. Faget in propulsion; Thomas A. Toll in configuration, stability, and control; Norris F. Dow in structures and mate­rials; and James B. Whitten in piloting. All four fell under the direction of Becker. Besides the almost mandatory elements of stability, control, and piloting, a fourth objective was outlined that would come to dominate virtual­ly every other aspect of the aircraft’s design— it would be optimized for research into the 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.

The preliminary specifications for the research aircraft were surprisingly brief: only four pages of requirements, plus six addition­al pages of supporting data. 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 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 made even more specific: “It shall be possible to design and construct the airplane within 3 years.”9 As John Becker subsequent­ly observed, "… it was obviously impossible that the proposed aircraft be in any sense an optimum hypersonic configuration.”

In developing the general requirements, the team developed a conceptual research aircraft that served as a model for the eventual X-15. The aircraft they conceived was “… not pro­posed as a prototype of any of the particular

The first Bell X-2 (46-674) made its ini­tial unpowered glide flight on 5 August 1954. This aircraft made a total of 17 flights before it was lost on 27 September 1956. Its pilot, Air Force Captain Milburn Apt had flown to a record speed 2,094 mph, thereby becom­ing the first person to exceed Mach 3.

The Genesis of a Research Airplane(NASA/DFRC)

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.” The merits of this approach had been convincing­ly demonstrated by the successes of the X-l and other dedicated research aircraft of the late 1940s and early 1950s.10

Assuming that the new vehicle would be air launched like the X-l and X-2, Langley estab­lished an aircraft size that could conveniently be carried by a Convair B-36, the largest suit­able aircraft available in the inventory. This translated to a gross weight of approximately

30.0 pounds, including 18,000 pounds of fuel and instrumentation.11 A maximum speed of 4,600 mph and an altitude potential of

400.0 feet were envisioned, with the pilot subjected to approximately 4.5g (an accelera­tion equal to 4.5 times the force of gravity) at engine burnout.12

The proposed maximum speed was more than double that achieved by the X-2, and placed the aircraft in a region where heating was the
primary problem associated with structural design, and where very little background information existed. Hypersonic aerodynam­ics was in its infancy in 1954. The few small hypersonic wind tunnels then in existence had been used almost exclusively for fluid mechanics studies, and they were unable to simulate either the high temperatures or the high Reynolds numbers of actual flight. It was generally believed that these wind tunnels did not produce valid results when applied to a full-scale aircraft. The proposed hypersonic research airplane, it was assumed, would pro­vide a bridge over the huge technological gap that appeared to exist between laboratory experimentation and actual flight.12

One aspect of the Langley proposal caused considerable controversy. The Langley team called for two distinct research flight profiles. The first consisted of a variety of constant angle-of-attack, constant altitude, and maneu­vering flights to investigate the aerodynamic and thermodynamic characteristics and limi­tations of then-available technology. These were the essential hypersonic research flights. But the second flight profile was designed to explore some of the problems of manned

space flight by making "… long leaps out of the sensible atmosphere.” This included inves­tigations into high-lift and low-L/D (lift over drag; commonly called a drag coeffi­cient) during the reentry pull-up maneuver” which was recognized as a prime problem for manned space flight from both a heating and piloting perspective.14

This brought other 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 aeromed – ical standpoint.” By employing a high altitude ballistic trajectory to approximately 250,000 feet, the Langley group expected the pilot would operate in an essentially weightless condition for approximately two minutes. Attitude control was another problem, since traditional aerodynamic control surfaces would be useless at the altitudes proposed for the new aircraft; the dynamic pressure would be less than 1 pound per square foot (psf). The
use of small hydrogen-peroxide thrusters for attitude control was proposed.

While the hypersonic research aspect of the Langley proposal enjoyed virtually unani­mous support, it is interesting to note that the space flight aspect was viewed in 1954 with what can best be described as cautious toler­ance. There were few who believed that any space flight was imminent, and most believed that manned space flight in particular was many decades in the future, probably not until the 21st century. Several researchers recom­mended that the space flight research was pre­mature and should be removed from the рто – gram. Fortunately, it remained.15

Hypersonic stability was the first problem of really major proportion encountered in the study. Serious instability had already been encountered with the X-l and X-2 at Mach numbers substantially lower than those expected with the proposed hypersonic research aircraft, and it was considered a major challenge to create a solution that would permit stable flight at Mach 7.

Researchers at Langley discovered through

Подпись:Подпись:The Genesis of a Research AirplaneThe notional research airplane designed by John V. Becker’s group at Langley shows the basis for the eventual X-15. Note the bullet­shaped fuselage (similar to the X-1) and the configuration of the empennage. This was the shape most of the early wind tunnel and analytical studies were per­formed against. (NASA)

wind tunnel testing and evaluating high speed data from earlier X-planes that an extremely large vertical stabilizer was required if the thin sections then in vogue for supersonic aircraft were used. This was largely because of a rapid loss in the lift-curve slope of thin sections as the Mach number increased. The solution devised by McLellan, based on theoretical considerations of the influence of airfoil shape on normal force characteristics, was to replace the thin supersonic-airfoil section of the vertical stabilizer with a 10 degree wedge shape. Further, a variable-wedge vertical sta­bilizer was proposed as a means of restoring the lift-curve slope at high speeds, thus per­mitting much smaller surfaces, which were easier to design structurally and imposed a smaller drag penalty on the airframe. McLellan’s calculations indicated that this wedge shape should eliminate the disastrous directional stability decay encountered by the X-l and X-2.

edge, very similar to the one eventually used on the Space Shuttle orbiters. Both the brak­ing effect and the stability derivatives could be varied through wide ranges by variable deflection of the wedge surfaces. The flexibil­ity made possible by variable wedge deflec­tion was thought to 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. lfi

Two basic structural design approaches had been debated since the initiation of the study—first, a conventional low-temperature design of aluminum or stainless steel protect­ed from the high-temperature environment by a layer of assumed insulation; and second, an exposed hot-structure in which no attempt would be made to provide protection, but in which the metal used and the design approach would permit high structural temperatures.17

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

It was found from analysis of the heating pro­jections for various trajectories that the air­plane would need to accommodate tempera­tures of over 2,000 degrees Fahrenheit on the lower surface of the wing. At the time, there

This chart was used by Becker to demon­strate the relative dif­ferences between the nominal recovery tem­perature, compared to the temperatures expected to be sus­tained by an insulated structure and an appropriately designed heat-sink skin (hot-structure). Inconel X was the material of choice very early in the study.


was no known insulating technique that could meet this requirement. The Bell “double­wall” concept where a non Toad-bearing metal sandwich acted as the basic insulator, would later undergo extensive development, but in 1954, it was in an embryonic state and not applicable to the critical nose and leading edge regions. Furthermore, it required a heavy and space-consuming supplemental liquid cooling system. However, the study group felt that the possibility of local failure of any insu­lation scheme constituted a serious hazard. Finally, the problem of accurately measuring heat-transfer rates—one of the prime objec­tives of the new research aircraft program— would be substantially more difficult to accomplish with an insulated structure.

At the start of the study it was by no means obvious that the hot-structure approach would prove practical either. The permissible design temperature for the best available material was about 1,200 degrees Fahrenheit, which was far below the estimated equilibrium temperature peak of about 2,000 degrees Fahrenheit. It was clear that some form of heat dissipation would have to be employed—either direct internal cooling or heat absorption into the structure
itself. It was felt that either solution would bring a heavy weight penalty.

The availability of Inconel Xі* and its excep­tional strength at extremely high temperatures, made it, almost by default, the structural mate­rial preferred by Langley for a hot-structure design. During mid-1954, an analysis of an Inconel X structure was begun by Becker’s group; concurrently, a detailed thermal analy­sis was conducted. A subsequent stress study indicated that the wing skin thickness should range from 0.05 to 0.10 inches—about the same values found necessary for heat absorp­tion in the thermal analysis.

Thus it was possible to solve the structural problem for the transient conditions of a Mach 7 aircraft with no serious weight penal­ty for heat absorption. This was an unexpect­ed plus for the hot-structure. Together with the fact that none of the perceived difficulties of an insulated-type structure were present, the study group decided in favor of an uninsulat­ed hot-structure design.

Unfortunately, it later proved that the hot – structure had problems of its own, particularly

The Genesis of a Research AirplaneInconel X was easily the best high-tempera­ture alloy available during the 1950s. It possessed a rare combination of high tensile strength and the ability to withstand high temperatures. Although it proved somewhat difficult to work with, it did not impose some of the problems encountered with titanium on other high-speed aircraft projects. (NASA)

in the area of nonuniform temperature distri­bution. Detailed thermal analyses revealed that large temperature differences would develop between the upper and lower wing skin during the pull-up portions of certain tra­jectories. This unequal heating would result in intolerable thermal stresses in a conventional structural design. To solve this new problem, wing shear members were devised which did not offer any resistance to unequal expansion of the wing skins. The wing thus was essen­tially free to deform both spanwise and chord – wise with asymmetrical heating. Although this technique solved the problem of the gross thermal stresses, localized thermal-stress problems still existed in the vicinity of the stringer attachments. The study indicated, however, that proper selection of stringer pro­portions and spacing would produce an acceptable design free from thermal buckling.

During the Langley studies, it was discovered that differential heating of the wing leading edge produced changes in the natural torsion­al frequency of the wing unless some sort of flexible expansion joint was incorporated in its design. 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 this phenomenon, the leading edge was segmented and flexibly mounted in an attempt to reduce thermally induced buck­ling and bending.

With its research objectives and structure now essentially determined, the Langley team turned its attention to the questions of propulsion by examining various existing rocket propulsion systems. The most promis­ing configuration was found to be a grouping of four General Electric A1 or A3 Hermes rocket engines, due primarily to the “thrust stepping” (a crude method of modulating, or throttling, the thrust output) option this con­figuration provided.

The studies prompted the NACA to adopt the official policy that the construction of a manned hypersonic research airplane was fea­sible. In June 1954, Dr. Hugh L. Dryden sent a letter to Lieutenant General Donald Putt at Air Force Headquarters stating that the NACA was interested in the creation of a new manned research aircraft program that would explore hypersonic speeds and altitudes well in excess of those presently being achieved. The letter also recommended that a meeting between the NACA, Air Force Headquarters, and the Air Force SAB be arranged to discuss the project. Putt responded favorably, and also recom­mended that the Navy be invited to participate.

NACA representatives met with members of the Air Force and Navy research and develop­ment groups on 9 July 1954 to present the proposal for a hypersonic research aircraft as an extension of the existing cooperative research airplane program. It was soon dis­covered that the Air Force SAB had been making similar proposals to Air Force Headquarters, and that the Office of Naval Research had already contracted with the Douglas Aircraft Company to determine the feasibility of constructing a manned aircraft capable of achieving 1,000,000 feet altitude. Douglas had concluded that 700,000 foot alti­tudes would be possible from the reentry deceleration standpoint, but that the thermo­structural problem had not been thoroughly analyzed. It was agreed that a cooperative pro­gram would be more cost effective and likely lead to better research data at an earlier time.19

The Navy and Air Force representatives viewed the NACA proposal with favor, although each had some reservations. At the close of the meeting, however, there was agreement that both services would further study the details of the NACA proposal, and that the NACA would take the initiative to secure project approval from the Department of Defense.2"

Less than a month later, the Air Force identi­fied the principal shortcoming of the original Langley proposal—the apparent lack of a suitable rocket engine. In early August the Power Plant Laboratory at the Wright Air Development Center (WADC) pointed out

that “no current rocket engines” entirely satis­fied the NACA requirements, and emphasized that the Hermes engine was not designed to be operated in close proximity to humans—that it usually was fired only when shielded by concrete walls. Other major objections to the Hermes engine centered around its relatively early state of development, its limited design life (intended for missile use, it was not required to operate successfully more than once), and the apparent difficulty of incorpo­rating the ability to throttle it during flight.21 WADC technical personnel who visited Langley on 9 August drew a firm distinction between engines intended for piloted aircraft and those designed for missiles; the NACA immediately recognized the problem, but con­cluded that although program costs would increase, the initial feasibility estimates would not be affected.22

WADC’s official reaction to the NACA pro­posal was submitted to the Air Research and Development Command (ARDC) on 13 August.2’ Colonel V. R. Haugen reported “unanimous” agreement among WADC par­ticipants that the proposal was technically fea­sible; excepting the engine situation, there was no occasion for adverse comment. The evaluation forwarded by Haugen also con­tained a cost estimate of $12,200,000 “distrib­uted over three to four fiscal years" for two research aircraft and necessary government – furnished equipment. Estimated costs includ­ed: $1,500,000 for design work; $9,500,000 for construction and development, including flight test demonstration; $650,000 for gov­ernment furnished equipment, including engines, $300,000 for design studies and specifications; and $250,000 for modification of a carrier aircraft.24 Somewhat prophetically, one WADC official commented informally: “Remember the X-3, the X-5, [and] the X-2 overran 200 percent. This project won’t get started for $12,000,000.”"

On 13 September, the ARDC issued an endorsement of the NACA proposal, and rec­ommended that the Air Force “… initiate a project to design, construct, and operate a new research aircraft similar to that suggested by NACA without delay.” The aircraft, empha­sized ARDC, should be considered a pure research vehicle and should not be pro­grammed as a weapon system prototype. On 4 October 1954, Brigadier General Benjamin S. Kelsey, Deputy Director of Research and Development at Air Force Headquarters, stat­ed that the project would be a joint Navy – NACA-USAF effort managed by the Air Force and guided by a joint steering commit­tee. Air Force Headquarters further pointed out the necessity for funding a special flight test range as part of the project:4′

The NACA Committee on Aeronautics met on 5 October 1954 to consider the hypersonic research aircraft. During the meeting, historic and technical data were reviewed by various committee members including Walter C. Williams, De E. Beeler, and research pilot A. Scott Crossfield from the High-Speed Flight Station (HSFS). Williams’ support was cru­cial. Crossfield would later describe Williams as.. the man of the 20th Century who made more U. S. advanced aeronautical and space programs succeed than all the others together. … He had no peer. None. He was a very strong influence in getting the X-15 program launched in the right direction.”27

Although one Committee member expressed opposition to the proposed hypersonic research aircraft as an extension to the on­going test programs, the rest of the Committee supported the project. The Committee formal­ly adopted a resolution to build a Mach 7 research airplane (attached as an appendix to this monograph).2*

Because the anticipated cost of the project would require support from Department of Defense contingency funds as well as Air Force and Navy R&D funds, a formal Memorandum of Understanding (MoU) was drafted and sent around for signatures begin­ning in early November 1954. The MoU was originated by Trevor Gardner (Air Force Special Assistant for Research and Development), and was forwarded, respec-

tively, for the signatures of J. H. Smith Jr. H (Assistant Secretary of the Navy [Air]) and Hugh L. Dryden (Director of the NACA). Dryden signed the Moll on 23 December 1954, and returned executed copies to the Air Force and Navy.’0

The MoU (attached as an appendix to this monograph) provided that technical direction of the research project would be the responsi­bility of the NACA, acting.. with the advice and assistance of a Research Airplane Committee” composed of one representative each from the Air Force, Navy, and the NACA. Administration of the design and construction phases of the project was assigned to the Air Force. The NACA would conduct the flight research, with extensive support from the Air Force Flight Test Center. The Navy was essen­tially left paying 25 percent of the bills with little active roll in the project, although it would later supply biomedical expertise and a single pilot. The NACA and the Research Airplane Committee were charged with the responsibility for disseminating the research results to the military services and aircraft industry as appropriate based on various secu­rity aspects. The concluding statement on the MoU was: “Accomplishment of this project is a matter of national urgency.’’11

It should be noted that it was not unusual in the late 1940s and early 1950s for the military services to fund the development and con­struction of aircraft for the NACA to use in its flight test programs. This was how most of the testing on the X-l and others had been accom­plished. The eventual X-l5 would be the fastest, highest-flying, and most expensive of these joint projects.12

After the signed copies of the MoU were returned to all participants, the Department of Defense authorized the Air Force to issue invitations to contractors having experience in the development of fighter-type aircraft to participate in the design competition. After the Christmas holidays, on 30 December, the Air Force sent invitation-to-bid letters to 12 prospective contractors; Bell, Boeing,

Chance-Vought, Consolidated (Convair), Douglas, Grumman, Lockheed, Martin, McDonnell, North American, Northrop, and Republic. The letter asked those interested in bidding to notify Wright Field by 10 January 1955, and to attend a bidder’s conference on 18 January 1955.”

Attached to the letter were a preliminary out­line specification, an abstract of the Langley preliminary study, a discussion of possible engines, a list of data requirements, and a cost outline statement. Each bidder was required to satisfy various requirements set forth, except in the case of the NACA abstract which was presented as "… representative of possible solutions.”14

Grumman, Lockheed, and Martin expressed little interest in the competition and did not attend the bidder’s conference, leaving nine possible competitors. At the bidders’ confer­ence, representatives from the contractors met with NACA and Air Force personnel to discuss the competition and the basic design requirements.

During the bidders’ conference, the airframe manufacturers were informed that one prime proposal and one alternate proposal (that might offer an unconventional but superior solution to the problems involved) would be accepted from each company. It also was noted that an engineering study, only, would be required for a modified aircraft where an observer could be substituted for the research instrumentation (a Navy require­ment); that a weight allowance of 800 pounds, a volume of 40 cubic feet, and a power requirement of 2.25 kilowatts (kW) needed to be provided for research instru­mentation; and that the winning design would have to be built in 30 months and be capable of attaining speeds of Mach 6 and altitudes of 250,000 feet. Following the pre­liminary statements concerning the bidding, NACA personnel briefed the various compa­nies in attendance on new information that had resulted from late 1954 wind tunnel research that had taken place at Langley.

Subsequently, between the bidders’ confer­ence and the 9 May submission deadline, Boeing, Chance-Vought, Convair, Grumman, McDonnell, and Northrop notified the Air Force that they did not intend to submit for­mal proposals. This left Bell, Douglas, North American, and Republic. During this period, representatives from these companies met with NACA personnel on numerous occa­sions and reviewed technical information on various aspects of the forthcoming research airplane. The NACA also provided these con­tractors with further information gained as a result of wind tunnel tests in the Ames 10-by – 14 inch supersonic tunnel and the Langley Mach 4 blowdown tunnel.

On 17 January 1955, NACA representatives met with Air Force personnel at Wright Field and were informed that the research airplane was identified as Air Force Project 1226 and would be officially designated X-15.

The Power Plant Laboratory had originally listed the Aerojet XLR73, Bell XLR81, North American NA-5400 (an engine in early devel­opment, still lacking a military designation), and the Reaction Motors XLR10 (and its vari­ants, including the XLR30) as engines that the airframe competitors could use in their designs. Early in January, the laboratory had become concerned that the builders of engines other than those listed might protest the exclu­sion of their products. Consequently there emerged an explanation and justification of the engine selection process. It appeared that the engineers had confidence in the ability of the XLR81 and XLR73 to meet airplane require­ments, had doubts about the suitability of the XLR25 (a Curtiss-Wright product), and held the thrust potential of the XLR8 and XLR11 (similar engines) in low repute. For practical purposes, this exhausted the available Air Force-developed engines suitable for manned aircraft. The XLR10 and NA-5400 were the only Navy-developed engines viewed as acceptable in terms of the competition.35

Earlier, the engine manufacturers had been contacted for specific information about the engines originally listed as suitable for the X-15 program,36 and this information was dis­tributed to all four prospective airframe con­tractors.37 Due to its early development status, there was little data available for the North American NA-5400, and the Reaction Motors XLR10 was “not recommended” at the sug­gestion of the engine manufacturer itself. On 4 February each of the prospective engine contractors (Aerojet, Bell, North American, and Reaction Motors) was asked to submit an engine development proposal.38 Based on this, the Air Force very slightly relaxed the rigid limitations on engine selection, instructing competitors that.. if… an engine not on the approved list offers sufficient advantage, the airframe company may, together with the engine manufacturer, present justification for approval…” to the Air Force.35

On 9 May 1955, Bell, Douglas, North American, and Republic submitted their pro­posals to the Air Force. Two days later the technical data was distributed to the evaluation groups with a request that results be returned by 22 June.40 The final evaluation meeting was scheduled for 25 July at Wright Field.41

Shortly thereafter, Hartley A. Soule, as Chairman of the NACA evaluation group, sent the evaluation rules and processes to the NACA laboratories. The evaluation would be based on the technical and manufacturing competency of each contractor, schedule and cost estimates, design approach, and the research utility of each design. In order to expedite the evaluation, each of the NACA laboratories was assigned specific items to consider with responses to be returned to Soule no later than 13 June.

The evaluation of the engine would be made at the same time, but would be conducted sep­arate from that of the airframe contractor, with the possibility that the chosen engine might not be the one selected by the winning air­frame contractor.

On 10 June the HSFS results were sent to Soule, based on the design approach and

research utility aspects of the airframe, flight control system, propulsion unit, crew provi­sions, handling and launching, and miscella­neous systems. The proposals were ranked:

(1) Douglas; (2) North American; (3) Bell; and (4) Republic. The proposals from Douglas and North American were consid­ered almost equal on the basis of points.

The Ames final evaluation, on 13 June 1955, ranked the proposals: fl) North American;

(2) Douglas; (3) Bell; and (4) Republic. The North American structure was considered to be more representative of future aircraft and thus superior in terms of research utility. Douglas retained a simple and conventional magnesium structure, but in so doing avoided the very thermodynamic problems the research effort wished to explore.

The 14 June final evaluation from Langley ranked the proposals; (1) North American; (2) Douglas; (3) Republic; and (4) Bell. Langley felt that while the magnesium wing structure of Douglas was feasible, it was feared that local hot spots caused by irregular aerody­namic heating could weaken the structure and be subject to failure. North American’s use of Inconel X was believed to be an advantage.

The final order representing the overall NACA evaluation was (1) North American; (2) Douglas; (3) Bell; and (4) Republic. All of the laboratories involved in this portion of the evaluation considered both the North American and Douglas proposals to be much superior to those submitted by Bell and Republic.

As with the NACA evaluations, the Air Force found little difference between the Douglas and North American designs, point-wise, with both proposals significantly superior to those of Beil and Republic. The Navy evaluation found much the same thing, ranking the pro­posals: (1) Douglas; (2) North American; (3) Republic; and (4) Bell.

On 26-28 July, the Air Force, Navy, and NACA evaluation teams met to coordinate their separate results. The Air Force and the NACA concluded that the North American proposal best accommodated their require­ments. Accordingly, the Navy decided not to. be put in the position of casting the dissenting vote and after short deliberation, agreed to go along with the decision of the Air Force and the NACA. A combined meeting of the Air Force, Navy, and the NACA was held at NACA Fleadquarters on 12 August for the final briefing on the evaluation. Later, the Research Airplane Committee met, accepted the findings of the evaluation groups, and agreed to present the recommendation to the Department of Defense.

Interestingly, the North American proposal was by far the most expensive. The estimat­ed costs for three aircraft plus one static test article and supporting equipment were: Bell, $36.3 million; Douglas, $36.4 million; Republic, $47 million; and North American, $56.1 million.

Because the estimated costs submitted by North American were far – above the amount allocated for the project, the Research Airplane Committee included a recommenda­tion for a funding increase that would need to be approved before the actual contract was signed. A further recommendation, one that would later take on greater importance, called for relaxing the proposed schedule by up to one-and-one-half years. These recommenda­tions were sent to the Assistant Secretary of Defense for Research and Development.

Events took an unexpected twist on 23 August when the North American represen­tative in Dayton verbally informed the Air Force that the company wished to withdraw its proposal. On 30 August, North American sent a letter to the Air Force formally requesting that the company be allowed to withdraw from consideration.43

The Vice President and Chief Engineer for North American, Raymond H. Rice, wrote to the Air Force on 23 September and explained that the company had decided to withdraw



from the competition because it had recently won new bomber and long range interceptor competitions and also had increased activity relating to its on-going F-107 fighter. Having undertaken these projects, North American said it would be unable to accommodate the fast engineering man-hours build-up that would be required to support the desired schedule. Rice went on that, “… due to the apparent interest that has subsequently been expressed in the North American design, the contractor [North American] wishes to extend two alternate courses which have been previ­ously discussed with Air Force personnel: The engineering man-power work load schedule has been reviewed and the contractor wishes to point out that Project 1226 could be han­dled if it were permissible to extend the schedule… over an additional eight month period: in the event the above time extension is not acceptable and in the best interest of the project, the contractor is willing to release the proposal data to the Air Force at no cost.”43

As it turned out, the possibility of extending the schedule had already been approved on 12 August, allowing North American to with­draw its previous letter of retraction once it had been officially informed that it had won the contract.44 Accordingly, on 30 September 1955, the Air Force formally notified North American that its design had been selected as the winner. The other bidders were conse­quently notified of North American’s selec­tion and thanked for their participation.45

By 11 October, the estimate from North American had been reduced from $56,000,000 to $45,000,000 and the maxi­mum annual funds requirement from $26,000,000 to $15,000,000. Shortly there­after, the Department of Defense released the funds needed to start work. More meetings between the Air Force, the NACA, and North American were held on 27-28 October, large­ly to define changes to the aircraft configura­tion. On 18 November, letter contract AF33(600)-31693 was sent to North American, and an executed copy was returned on 8 December 1955.46 The detailed design
and development of the hypersonic research airplane had been underway for just under a year at this point.47

On 1 December 1955, a series of actions48 began that resulted in letter contract AF33(600)-32248 being sent to Reaction Motors, effective on 14 February 1956. Its ini­tial allocation of funds totaled $3,000,000, with an eventual expenditure of about $6,000,000 foreseen as necessary for the delivery of the first flight engine.49

A definitive contract for North American was completed on 11 June 1956, superseding the letter contract and two intervening amend­ments. At that time, $5,315,000 had been committed to North American. The definitive contract allowed the eventual expenditure of $40,263,709 plus a fee of $2,617,075. For this sum, the government was to receive three X-15 research aircraft, a high speed and a low speed wind tunnel model program, a free-spin model, a full-size mockup, propulsion system tests and stands, flight tests, modification of a B-36 carrier aircraft, a flight handbook, a maintenance handbook, technical data, peri­odic reports of several types, ground handling dollies, spare parts, and ground support equip­ment. Exclusive of contract costs were fuel and oil, special test site facilities, and expens­es to operate the B-36. The delivery date for the X-15s was to be 31 October 1958. The quantity of aircraft had been determined by experience; it had been noted during earlier research aircraft programs that two aircraft were enough to handle the anticipated work­load, but three assured that the test pace could be maintained even with one aircraft down.50 This lesson has been largely forgotten in our current budget-conscious era.

A final contract for the engine, the prime unit of government furnished equipment, was effective on 7 September 1956. Superseding the letter contract of February, it covered the expenditure of $10,160,030 plus a fee of $614,000.51 For this sum, Reaction Motors agreed to deliver one engine, a mockup, reports, drawings, and tools.

Hypersonic Research Engine

With little doubt, the most ambitious47 of the

X-15 experiments was the Hypersonic Research Engine (HRE) from the Langley Research Center. At the time that researchers began to consider supersonic-combustion ramjet engines during 1954, the X-15 was not an approved program and played no major role in the engine’s conceptual development. However, events soon transpired that made flight testing of a supersonic ramjet engine desirable, and the Flight Research Center and Langley proposed a joint project to accom­plish just that. The 1962 crash of the X-15-2 opened the door for extensive modification aimed primarily at providing a platform for development of the Mach 8 air-breathing HRE. Then, as now, no tunnel facility existed wherein such an engine could be realistically tested, and rocket boosters could not give steady-state tests or return the equipment.48

The actual prototype engine was to be carried attached to the lower ventral of the X-15A-2. Twenty-nine inches were added to the fuse­lage between the existing tanks for the liquid hydrogen to power the HRE, two external fuel tanks were added, and the entire aircraft was coated with an ablative-type insulator.

During 1965, Garrett-AirResearch was put under contract to provide six prototype engines by mid-1969. As would happen, the development effort necessary to produce a workable engine had been severely underesti­mated, and Garrett quickly ran into problems that caused serious delays in the project.

In the meantime flight-test evaluations were made of the modified aircraft itself and of a dummy HRE attached to the X-15A-2. On the first and only maximum-speed test of the X-15 A-2 in 1967, shock impingement off the dummy HRE caused severe heating damage to the lower empennage, and very nearly resulted in loss of the aircraft. Though quick­ly repaired, the X-15A-2 never flew again. Hindsight would place the blame for this design oversight on haste and insufficient flow interaction studies. A key lesson learned from this episode was not to hang external stores or pylons on hypersonic air-

craft, at least not without far more extensive study of underside flow patterns. As John Becker later observed, “Flight testing on the X-15A-2 would have been long-delayed, hazardous, very costly, and fortunately never came about.”4V

When the X-15 flight program was terminat­ed, the HRE degenerated into a costly wind tunnel program using partial-simulation test models. The HRE was eventually tunnel test­ed in 1969, and the primary objective of achieving supersonic combustion was met, although the thrust produced was less than the drag created. HRE engineers nonetheless claim a success in that the objective was supersonic combustion, not a workable engine. The program continued until 1975 and never achieved a positive net thrust, although it still contributed to the technology base, albeit at a very high cost. A hindsight study conducted in 1976 concluded that the HRE’s fuel-cooled structure was its main contribution to future scramjets.50

X-15 Design and Development

Harrison A. “Stormy” Storms, Jr. led the North American X-15 design team, along with project engineer Charles H. Feltz. These two had a difficult job ahead of them, for although giving the appearance of having a rather sim­ple configuration, the X-15 was perhaps the most technologically complex single-seat air­craft of its day. Directly assisting Storms and Feltz was test pilot A. Scott Crossfield, who had worked for the NACA prior to joining North American with the intended purpose of working on the X-15 program. Crossfield describes Storms as a man of wonderful imagination, technical depth, and courage… with a love affair with the X-15. He was a tremendous ally and kept the objectivity of the program intact….” According to Crossfield, Feltz was a remarkable ‘can do and did’ engineer who was very much a source of the X-15 success story.”1

Storms himself remembers his first verbal instructions from Hartley Sould: “You have a little airplane and a big engine with a large thrust margin. We want to go to 250,000 feet altitude and Mach 6. We want to study aero­dynamic heating. We do not want to worry about aerodynamic stability and control, or the airplane breaking up. So if you make any errors, make them on the strong side. You should have enough thrust to do the job.” Adds Storms, “and so we did.”2

Crossfield’s X-15 input proved particularly noteworthy during the early days of the development program as his experience per­mitted the generation of logical arguments that led to major improvements to the X-15. He played a key role, for instance, in con­vincing the Air Force that an encapsulated ejection system was both impractical and

X-15 Design and DevelopmentBy the time of the first industry conference in 1956, this was the design baseline for the North American X-15. Note the tall ver­tical stabilizer, and the fact that it does not have the distinctive wedge shape of the final unit. Also notice how far forward the fuselage tunnels extend—well past the canopy. (NASA)

unnecessary. His arguments in favor of an ejection seat, capable of permitting safe emergency egress at speeds between 80 mph and Mach 4 and altitudes from sea level to 120,000 feet, saved significant money, weight, and development time.

There has been considerable interest over whether Crossfield made the right decision in leaving the NACA since it effectively locked him out of the high-speed, high-alti­tude portion of the X-15 flight program. Crossfield has no regrets: “… I made the right decision to go to North American. I am an engineer, aerodynamicist, and designer by training… While I would very much have liked to participate in the flight research pro­gram, I am pretty well convinced that f was needed to supply a lot of the impetus that allowed the program to succeed in timeli­ness, in resources, and in technical return…. I was on the program for nine years from conception to closing the circle in flight test. Every step: concept, criteria, requirements, performance specifications, detailed specifi­cations, manufacturing, quality control, and flight operations had all become an [obses­sion] to fight for, protect, and share—almost with a passion.”3

Although the first, and perhaps the most influential pilot to contribute to the X-15 program, Crossfield was not the only one to do so. In fact, all of the initially assigned X-15 pilots participated in the development phases, being called on to evaluate various operational systems and approaches, as well as such factors as cockpit layout, control sys­tems, and guidance schemes. They worked jointly with engineers in conducting the sim­ulator programs designed to study the aspects of planned flight missions believed to present potential difficulties. A fixed – base simulator was developed at North American’s Los Angeles facility, containing a working X-15 cockpit and control system that included actual hydraulic and control – system hardware. Following use at North American, it was subsequently relocated to the Flight Research Center4 (FRC) at Edwards AFB. Once flight research began, the simulator was constantly refined with the results of the flight test program, and late in its life the original analog computers were replaced by much faster digital units. For the life of the program, every X-15 flight was preceded by 10-20 hours in the simulator.

A ground simulation of the dynamic envi-



X-15 Design and DevelopmentПодпись:Подпись: VELOCITY 6600 FT PER SEC DESIGN ALTITUDE 260,000 FT LANDING SPEED 164 KN POWER PLANT*RMl MAX THRUST (40.000 FT) 57,000 LB MIN THRUST (40,000 FT) 17,000 LB WING AREA 200 SQ FT SWEEP c/4 25 0E6BEES THICKNESS 5 PERCENT ASPECT RATIO 2.5 WEIGHT LAUNCHING 31.275 LB BURN-OUT 12.971 LB PROPELLANT 13.304 LB X-15 Design and DevelopmentOne of the more con­troversial features of the North American design was the fuse­lage tunnels that car­ried the propellant lines and engine con­trols around the full monocoque propellant tanks, shown in this 1956 sketch. Originally these tunnels extend­ed forward ahead of the cockpit, and the NACA worried they would create unac­ceptable vortices. (NASA)

ronment was provided by use of the Navy centrifuge at the Naval Air Development Center (NADC) Johnsville, Pennsylvania. Over 400 simulated reentries5 were flown during an initial round of tests completed on 12 July 1958; Iven Kincheloe, Joe Walker, Scott Crossfield, A1 White, Robert White, Neil Armstrong, and Jack McKay participat­ed. The primary objective of the program was to assess the pilot’s ability to make emergency reentries under high dynamic conditions following a failure of the stability augmentation system. The results were gen­erally encouraging.6

When the contracts with North American had been signed, the X-15 was some three years away from actual flight test. Although most of the basic research into materials and structural science had been completed, a great deal of work remained to be accom­plished. This included the development of fabrication and assembly techniques for Inconel X and the new hot-structure design. North American met the challenge of each problem with a practical solution, and even­tually some 2,000,000 engineering man­hours and 4,000 wind tunnel hours in 13 dif­ferent wind tunnels were logged.

The original North American proposal gave rise to several questions which prompted a meeting at Wright-Patterson AFB on 24-25 October 1955. Subsequent meetings were held at the North American Inglewood plant on 28-29 October and 14-15 November. Major discussion items included North American’s use of fuselage tunnels and all­moving horizontal stabilizers (the “rolling – tail”). The rolling-tail operated differentially to provide roll control, and symmetrically to provide pitch control; this allowed the elimi­nation of conventional ailerons. North American had gained considerable experi­ence with all-moving control surfaces on the YF-107A fighter. In this instance the use of differentially operated surfaces simplified the construction of the wing, and allowed elimination of protuberances that would have been necessary if aileron actuators had been incorporated in the thin wing. Such pro­tuberances would have disturbed the airflow and created another heating problem.

One other significant difference between the configuration of the NACA design and that of the actual X-15 stemmed from North American’s use of full-monocoque propel­lant tanks in the center fuselage and the use

The interior layout of the fuselage did not change much after the 1956 conference. Note the helium tank locat­ed in the middle of the LOX tank. The hydro­gen-peroxide (H202) was used to power the turbopump on the XLR99 rocket engine.

X-15 Design and Development(NASA)

of tunnels on both sides of the fuselage to accommodate the propellant lines and engine controls that ordinarily would have been contained within the fuselage. The NACA expressed concern that the tunnels might cre­ate undesirable vortices that would interfere with the vertical stabilizer, and suggested that the tunnels be kept as short as possible in the area ahead of the wing. North American agreed to make the investigation of the tunnels’ effects a subject of an early wind tunnel-model testing program.7

During the spring and summer of 1956, sev­eral scale models were exposed to rather intensive wind tunnel tests. A 1/50-scale – model was tested in the 11 – inch hypersonic and 9-inch blowdown tunnels at Langley, and another in a North American wind tun­nel. A 1/15-scale model was also tested at Langley and a rotary-derivative model was tested at Ames. The various wind tunnel pro­grams included investigations of the speed
brakes, horizontal stabilizers without dihe­dral, several possible locations for the hori­zontal stabilizer, modifications of the vertical stabilizer, the fuselage tunnels, and control effectiveness, particularly of the rolling-tail. Another subject in which there was consid­erable interest was determining the cross­section radii for the leading edges of the var­ious surfaces.

On 11 June 1956, North American received a production go-ahead for the three X-15 air­frames (although the first metal was not cut for the first aircraft until September). Four days later, on 15 June 1956, the Air Force assigned three serial numbers (56-6670 through 56-6672) to the X-15 program.8

By July, the NACA felt that sufficient progress had been made on the X-15 devel­opment to make an industry conference on the project worthwhile.8 The first Conference on the Progress of the X-15 Project was held



X-15 Design and DevelopmentSeven different wind tunnels are represent­ed in this chart show­ing how the extreme front of the fuselage tunnels began to be modified. Note the large speed brakes on the vertical stabilizer.

“LAC on the chart is the Langley Aeronautical Laboratory, while “AAL” is the Ames Aeronautical Laboratory.


at Langley on 25-26 October 1956. There were 313 attendees representing the Air Force, the NACA, Navy, various universities and colleges, and most of the major aero­space contractors. It was evident from the papers that a considerable amount of progress had already been made, but that a few significant problems still lay ahead.10

A comparison of the suggested configuration contained in the original NACA proposal and the North American configuration pre­sented to the industry conference revealed that the span of the X-15 had been reduced from 27.4 feet to only 22 feet and that the North American fuselage had grown from the suggested 47.5-foot overall-length to 49 feet. North American followed the NACA suggestion by selecting Inconel X as the major structural material and in the design of a multispar wing with extensive use of cor­rugated webs.11

One of the papers summarized the aerody­namic characteristics that had been obtained by tests in eight different wind tunnels.12 These tests had been made at Mach numbers ranging from less than 0.1 to about 6.9, and investigated such problems as the effects of speed brake deflection on drag, the lift-drag relationship of the entire aircraft, of individ­ual components such as the wings and fuse­lage tunnels, and of combinations of individ­ual components. One of the interesting prod­ucts was a finding that almost half of the total lift at high Mach numbers would be derived from the fuselage tunnels. Another
result was the confirmation of the NACA’s prediction that the original fuselage tunnels would cause longitudinal instability; for sub­sequent testing the tunnels had been short­ened in the area ahead of the wing, greatly reducing the instability. Still other wind tun­nel tests had been conducted in an effort to establish the effect of the vertical and hori­zontal tail surfaces on longitudinal, direc­tional, and lateral stability.

It should be noted that wind tunnel testing in the late 1950s was, and still is, an inexact sci­ence. For example, small (3- to 4-inch) mod­els of the X-15 were “flown” in the hyper­velocity free-flight facility at Ames. The models were made out of cast aluminum, cast bronze, or various plastics, and were actually fairly fragile. Despite this, the goal was to shoot the model out of a gun at tremendous speeds in order to observe shock wave patterns across the shape. As often as not, what researchers saw were pieces of X-15 models flying down the range side­ways, Fortunately, enough of the models remained intact to acquire meaningful data.13

Other papers presented at the industry con­ference dealt with research into the effect of the aircraft’s aerodynamic characteristics on the pilot’s control. Pilot-controlled simula­tion flights for the exit and reentry phases had been conducted; researchers reported that the pilots had found the early configura­tions nearly uncontrollable without damping, and that even with dampers the airplane pos­sessed only minimum stability during parts

These charts show the expected tempera­tures and skin thick­ness for various parts of the Х-15’s fuselage.

X-15 Design and DevelopmentNote the large differ­ence between top-side temperatures and those on the bottom of the fuselage. (NASA)



of the programmed flight plan. A program utilizing a free-flying model had proved low – speed stability and control to be adequate. Since some aerodynamicists had questioned North American’s use of the rolling-tail instead of ailerons, the free-flying model had also been used to investigate that feature. The results indicated that the rolling-tail would provide the necessary lateral control.

Several papers presented at the conference dealt with aerodynamic heating. One of these was a summary of the experience gained with the Bell X-1B and X-2. The information was incomplete and not fully applicable to the X-15, but it did provide a basis for compari­son with the results of the wind tunnel and analytical studies. Another paper dealt with the results of the structural temperature esti­mates that had been arrived at analytically. It was apparent from the contents of the papers that the engineers compiling them were con­fronted by a paradox—in order to attain an adequate and reasonably safe research vehi­cle, they had to foresee and compensate for the very aerodynamic heating problems that were to be explored by the completed aircraft.

In addition to the papers on the theoretical aspects of aerodynamic heating, a report was made on the structural design that had been accomplished at the time of the conference. Critical loads would be encountered during the accelerations at launch weight and during reentry into the atmosphere, but since maxi­mum temperatures would be encountered only during the latter, the paper was largely
confined to the results of the investigations of the load-temperature relationships that were anticipated for the reentry phase. The selection of Inconel X skin for the multispar box-beam wing was justified on the basis of the strength and favorable creep characteris­tics of that material at 1,200 degrees Fahrenheit. A milled bar of Inconel X was to be used for the leading edge since that por­tion of the wing acted as a heat sink. The internal structure of the wing was to be of titanium-alloy sheet and extrusion construc­tion. The front and rear spars were to be flat web-channel sections with the intermediate spars and ribs of corrugated titanium webs.

For purposes of the tests the maximum tem­perature differences between the upper and lower wing surfaces had been estimated to be 400 degrees Fahrenheit and that between the skin and the center of the spar as 960 degrees Fahrenheit. Laboratory tests indicated that such differences could be tolerated without any adverse effects on the structure. Other tests had proven that thermal stresses for the Inconel-titanium structure were less than those encountered in similar structures con­structed entirely on Inconel X, Full-scale tests had been made to determine the effects of temperature on the buckling and ultimate strength of a box beam. Simply heating the test structure produced no surface buckles. Compression buckles had appeared when ultimate loads were applied at normal tem­peratures but the buckles disappeared with the removal of the load. Tests at higher tem­peratures and involving large temperature





The wing of the X-15 was constructed from Inconel X skins over a titanium struc­ture. Unlike many air­craft, there was not a continuous spar across both wings. Instead, each wing was bolted to the fuselage. (NASA)




X-15 Design and Development

differences had finally led to the failure of the test box, but it seemed safe to conclude that . thermal stresses had very little effect on the ultimate strength of the box.”

Tests similar to those conducted on the wing structure had also been performed on the hor­izontal stabilizer. The planned stabilizer struc­ture differed from the wing in that it incorpo­rated a stainless steel spar about halfway between the leading and trailing edges, and an Inconel X spar three and one-half inches from the leading edge. The remainder of the inter­nal structure consisted of titanium compo­nents and the skin was Inconel X sheet. Tests of the stabilizer had indicated that a design which would prevent all skin buckling would be inordinately heavy, so engineers decided to tolerate temporary buckles. The proposed sta­bilizer had flutter characteristics that were within acceptable limits.

The front and rear fuselage were semimono – coque structures of titanium ribs, Inconel X outer skin, and an inner aluminum skin insu­lated with spun glass. The integral propellant tanks in the center fuselage were of full monocoque construction. The full mono – coque design used only slightly thicker skins than the semimonocoque design, possessed adequate heat sink properties, and reduced stresses caused by temperature differences by placing all of the material at the surface. It seemed, therefore, that the resulting structure was ideal for use as a pressure tank. The thickness of the monocoque walls would also make sealing easier and leaks less likely.

The fuselage side tunnels presented yet another problem. As the tunnels would pro­tect the side portions of the propellant tanks from aerodynamic heating, the sides would not expand as rapidly as the areas exposed to the air, and another undesirable compressive stress had to be anticipated. It was thought that beading the skin of the areas protected by the tunnels would provide a satisfactory solution, but beading introduced further complications by reducing the structure’s ability to carry pressure loads. Ultimately, however, the techniques proved successful.

Like most rocket engines of the period, the XLR99 would use liquid oxygen as an oxi­dizer, and a non-cryogenic fuel, in this case anhydrous ammonia.14 Each of the two main propellant tanks was to be divided into three compartments by curved bulkheads; the two compartments furthest from the aircraft cen­ter of gravity were equipped with slosh baf­fles. Plumbing was to be installed in a single compartment, the compartment sealed by a bulkhead, and the process repeated until all the compartments were completed. The tank ends were to be semicurved in shape to keep them as flat as possible, to reduce weight, and to permit thermal expansion of the tank shell. This entire structure was to be of weld­ed Inconel X.

The expected acceleration of the X-15 pre­sented several unique human factors concerns early in the program. It was expected that the pilot would be subjected to an acceleration of up to 5g. It seemed advisable to develop a

X-15 Design and Development

One of the innovations proposed by North American was the use of monocoque propel­lant tanks, leading to the use of the contro­versial fuselage tun­nels. The forward-most part of the LOX tank was equipped with slosh baffles. (NASA)

side-stick controller that would allow the pilot’s arm to be supported by an armrest while still allowing him of full control over the aircraft.15 Coupled with the fact that there were two separate attitude-control systems on the X-15, this resulted in a unique control stick arrangement. A conventional center stick, similar to that installed in most fighter – type aircraft of the era, was connected to the aerodynamic control surfaces through a sta­bility-augmentation (damper) system. A side – stick controller on the right console was con­nected to the same aerodynamic control sur­faces and augmentation system. Either stick could be used interchangeably, although the flight manual16 describes using the center stick “during normal periods of longitudinal and vertical acceleration.” The center stick was occasionally omitted from flights later in the flight research program based on pilot prefer­ences. Another side-stick controller on the left console operated the so-called “ballistic con­trol” system17 (thrusters) that provided attitude control at high altitudes. The flight manual warns that “velocity tends to sustain itself after the stick is returned to the neutral posi­tion. A subsequent stick movement opposite to the initial one is required to cancel the orig­inal attitude change.”

At the time of the industry conference in 1956, the design for the X-15 side controller had not been definitely established but a summary of the previous experience with such controllers was available. Experimental controllers had been installed on a Grumman F9F-2, Lockheed TV-2, Convair F-102, and on a simulator. The pilots who had tried side controllers had reported no difficulty in maneuvering, but they generally felt that greater efforts would have to be made to eliminate backlash and to control friction forces; they had also urged that efforts be made to give the side controllers a more “natural” feel.

Another problem which had not been thor­oughly explored at the time of the 1956 con ference concerned the proposed reaction con­trols that would be necessary for the X-15 as dynamic pressures decreased to the point where the aerodynamic controls would no longer be effective. Analog computer and ground simulator studies were then under way in an effort to determine the best relationship between the control thrust and the pilot’s movement of the control stick. Attempts were also being made to determine the amount of propellant that would be required for the reac-





The X-15 contained two side-stick con­trollers; one for the aerodynamic controls (shown), and one on the other console for the reaction controls. Although the side-stick proved very success­ful on the X-15, it would be another 20 years before one was installed on an opera­tional aircraft (the General Dynamics F-16). (NASA)















X-15 Design and Development

tion controls. No significant problems were uncovered during these early investigations, but it was clear that the pilot would have to give almost constant attention to such a con­trol system and that pilots should be given extensive practice on simulators before being allowed to attempt actual flight.

Some of the anticipated difficulties in the field of instrumentation arose because available strain gauges were not considered satisfactory at the expected high temperatures and because of difficulties in recording the output of ther­mocouples. Large structural deformations of wings and empennage were to be recorded by cameras in special camera compartments. Another instrumentation problem arose because the sensing of static pressure, ordi­narily difficult at high Mach numbers, was compounded in the case of the X-15 by heat­ing that would be too great for any conven­tional probe and by the low pressure at the high altitudes to be explored. The answer was to develop a stable-platform-integrating- accelerometer system to provide velocity, alti­tude, pitch, yaw, and roll angle information.

Still another instrumentation difficulty was created by the desirability of presenting the
pilot with angle-of-attack and side slip infor­mation, especially for the critical exit and reentry periods. Any device to furnish this information would have to be located ahead of the aircraft’s own flow disturbances, be structurally sound at elevated temperatures, accurate at low pressures, and cause a mini­ma) flow disturbance so as not to interfere with the heat transfer studies that were to be conducted in the forward area of the fuse­lage. These requirements had resulted in the design of a ball-nose16 capable of withstand­ing 1,200 degrees Fahrenheit. A six-inch diameter Inconel X sphere located in the extreme nose of the X-15 was gimbaled19 and servo-driven in two planes. It had five open­ings: a total-head port opening directly for­ward and two pairs of angle-sensing ports in the pitch and yaw planes, located at an angle of 30 to 40 degrees from the central port. Pitch and yaw could be sensed as pressure differences and these differences were con­verted into signals that would cause the ser­vos to realign the sphere in the relative wind.

Based largely on urgings from Scott Crossfield, the Air Force agreed to allow North American to design an ejection seat and to make a study justifying the selection

X-15 Design and DevelopmentAlthough the ejection seat showed at the 1956 industry confer­ence did not resemble the final unit used in the X-15s, the basic concepts remained the same. Restraining the pilot’s head, arms, and legs during ejec­tion at high dynamic pressures presented one of the major chal-. lenges to seat devel­opment. (NASA)

of a seat in preference to a capsule system.20 Two main criteria had governed the selec­tion of an escape system for the X-15, and these criteria were not necessarily comple­mentary. The first requirement was that the system be the most suitable that could be designed while remaining compatible with the airplane. The second was that no system would be selected that would delay the development of the X-15 or leave the pilot without any method of escape when the time arrived for flight research. The four possible escape systems that were consid­ered included cockpit capsules, nose cap­sules, a canopy shielded seat, and a stable – seat with a pressure-suit. An analysis of the expected flight hazards had indicated that because of the fuel exhaustion and low aerodynamic loads, the accident potential at peak speeds and altitudes was only about two percent of the total.

Capsule-like systems had been tried before, most notably in the X-2 where the entire for­ward fuselage could be detached from the rest of the aircraft. Model tests showed these to be very unstable and prone to tumble at a high rate of rotation. They also added a great deal of weight and complexity to the aircraft,21

The final decision for a stable-seat with a pressure-suit was made because most of the potential accidents could be expected to occur at speeds of Mach 4 or less, because system reliability always decreased with sys­tem complexity, and finally, because it was the system that imposed the smallest weight and size penalties upon the aircraft. The selected system would not function success­fully at altitudes above 120,000 feet or speeds in excess of Mach 4, but designers, particu­larly Scott Crossfield, held that the aircraft itself would offer the best protection in the areas of the performance envelope where the seat-suit combination was inadequate.

Cockpit and instrument cooling, pressuriza­tion, suit ventilation, windshield defogging, and fire protection were all to be provided from a liquid nitrogen supply. Vaporization of the liquid nitrogen would keep the pilot’s environment within comfortable limits at all times. An interesting aspect of the cooling problem was an estimate that only 1.5 per­cent of the system’s capacity would be applied to the pilot; the remaining 98.5 per­cent was required for equipment. Cockpit temperatures were to be limited to no more than 150 degrees Fahrenheit, the maximum

X-15 Design and DevelopmentThis chart shows that 92 percent of the expected X-15 acci­dents would happen below Mach 2 and

90,0 feet. This esti­mate supported Scott Crossfield’s request to use an ejection seat and pressure suit instead of a more complex escape cap­sule. (NASA)

limit for some of the equipment. The pilot would not be subjected to that temperature, however, as the pressure suit ventilation would enable him to select a comfortable temperature level. Cockpit pressure was to be maintained at the 35,000 foot level.

The effects of flight accelerations upon the pilot’s physiological condition and upon his ability to avoid inadvertent control move­ments had not been completely explored, but it was recognized that high accelera­tions could pose medical and restraint diffi­culties. In addition to the accelerations that would be encountered during the exit and reentry phases of the X-15’s flights, a very high acceleration of short duration would be produced during the landings. This was a result of the location of the main skids at the rear of the aircraft. Once the skids touched down, the entire aircraft would act as if it were hinged at the skid attachment points and the nose section would slam downward. Reproduction of this landing acceleration on simulators showed that because of the short duration, no real prob­lem existed. There were, however, numer­ous complaints about the severity of the jolts both in the simulator and once actual landings began.

The final paper presented to the 1956 indus­try conference was an excellent summary of the development effort and a review of the major problems that were known at that time. The author, Lawrence R Greene from North American, considered flutter to be an unsolved problem, primarily because of a lack of basic data on aero-thermal-elastic relationships and because little experimental data was available on flutter at hypersonic Mach numbers. He pointed out that available data on high-speed flutter had been derived from experiments conducted at Mach 3 or less, and that not all of the data obtained at those speeds were applicable to the problems faced by the designers of the X-15. As it turned out, panel flutter was encountered early in the flight test program, leading to a change in the design criteria for high-speed aircraft. Another difficulty was the newness of Inconel X as a structural material and the necessity of experimenting with fabrication techniques that would permit its use as the primary structural material for the X-15. Problems were also expected to arise in con­nection with sealing materials, most of which were known to react unfavorably when subjected to high temperature condi­tions.22 Although North American did encounter initial problems in using Inconel

X-15 Design and DevelopmentDespite its perform­ance potential, the basic cockpit design of the X-15 was quite conventional, with the exception of the side – stick controllers. The engine instrumenta­tion on the lower left of the instrument panel would be differ­ent for the XLR11 flights. The addition of the MH-96 in the X-15-3 would necessi­tate some changes in the instrumentation. See page 63 for a photo. (NASA)

X and titanium during the construction of the X-l5, it was able to work through the diffi­culties with no major delays.

A development engineering inspection was held at the North American Inglewood plant on 12-13 December 1956. This inspection of a full-scale tnockup was intended to reveal unsatisfactory design features before fabrication of the aircraft got under way. Thirty-four of the forty-nine individuals who participated in the inspection were rep­resentatives of the Air Force; twenty-two of them from WADC. The important role of the Air Force was also evident from the composition of the committee that would review the requests for alteration.2’ Major E. C. Freeman, of ARDC, served as committee chairman, Mr. F. Orazio of WADC and Lieutenant Colonel Keith G. Lindell of Air Force Headquarters were committee mem­bers, and Captain Chester E. McCollough, Jr. of the ARDC and Captain Iven C. Kincheloe, Jr. of the Air Force Flight Test Center (AFFTC) served as advisors. The Navy and the NACA each provided a single committee member; three additional advi­sors were drawn from the NACA.

The inspection committee considered 84 requests for alterations, rejected 12, and placed 22 in a category for further study. The majority of the 50 changes that were accept­ed were minor, such as the addition of longi­tudinal trim indications from the stick posi­tion and trim switches, relocation of the bat­tery switch, removal of landing gear warning lights, rearrangement and redesign of warn­ing lights, and improved markings for sever­al instruments and controls.

Some of the most interesting comments were rejected by the committee. For instance, the suggestions that the aerodynamic and reac­tion controller motions be made similar, that the reaction controls be made operable by the same controller used for the aerodynam­ic controls, or that a third controller combin­ing the functions of the aerodynamic and reaction controllers be added to the right console, were all rejected on the grounds that actual flight experience was needed with the controllers already selected before a decision could be made on worthwhile improvements or combinations. As two of the three sugges­tions on the controllers came from potential pilots of the X-l5 (Joseph A. Walker and

X-15 Design and DevelopmentThe vertical stabilizer was one of the most obvious changes between the industry conference configura­tion and the final vehi­cle. The first design did not use the exag­gerated wedge-shape of the final unit. It was also more traditional, using a fixed forward portion and a conven­tional appearing rud­der. The final version used an all-moving design. Note the rud­der splits to become speed brakes, much like the shuttle design 25 years later. (NASA)

Iven C. Kincheloe, Jr.24), it would appear that the planned controllers were not all that might have been desired.

A request that the pilot be provided with continuous information on the nose-wheel door position (loss of the door could produce severe structural damage) was rejected because the committee felt that the previous­ly approved suggestion for gear-up inspec­tion panels would make such information unnecessary. This particular item would come back to haunt the program during the flight research phase.

After the completion of the development engineering inspection, the X-15 airframe design changed only in relatively minor details. North American essentially built the X-15 described at the industry conference in October and inspected in mockup in December 1956. Continued wind tunnel test­ing resulted in some external modifications, particularly of the vertical stabilizer, and some weight changes occurred as plans became more definite. But while work on the airframe progressed smoothly, with few unexpected problems, the project as a whole did encounter difficulties, some of them seri­ous enough to threaten long delays. In fact, North American’s rapid preparation of draw­ings and production planning served to high­light the lack of progress on some of the components and subsystems that were essen­tial to the success of the program.