Category X-15 EXTENDING THE FRONTIERS OF FLIGHT

In order to get flight-testing under way, North American completed the first two aircraft with interim Reaction Motors XLR11-RM-5 engines. Two XLR11s were installed in each aircraft, producing a total of 11,800 lbf at sea level. These engines were quite familiar to personnel working in the experimental rocket aircraft programs at Edwards, since the Bell X-1, Douglas D – 558-2, and Republic XF-91 all used the same powerplant (or its XLR8 Navy equivalent).-1132!

The basic XLR11 configuration was called G6000C4 by Reaction Motors and consisted of four thrust chambers producing 1,475 lbf each with a turbopump unit, valves, regulators, and controls mounted forward of the chambers. Other variants of the XLR8/XLR11 family used pressure-fed propellants instead of a turbopump. The four chambers were mounted on a support beam assembly that was the main structural member of the engine. A single turbopump provided the pressure to inject the liquid-oxygen and ethyl-alcohol-water propellants, while valves in the oxidizer and fuel lines controlled the flow of the propellants to the chambers. Each thrust chamber contained an igniter, and the pilot could ignite or shut down individual chambers in any sequence, allowing a measure of "thrust stepping." However, once the pilot shut down a chamber, he could not restart that chamber. Fuel circulated through passages in each exhaust nozzle and around each combustion chamber individually for cooling, and then into the firing chambers to be burned. Each engine weighed 345 pounds dry (including pumps) and was approximately 60 inches long, 36 inches high, and 24 inches wide. On paper each engine (including the turbopumps) cost about $80,000, although technicians at Edwards assembled all of the engines used in the X-15 program on site from components left over from earlier programs.-133!

It was surprisingly easy to install the XLR11 in the X-15, considering that the designers had not intended the aircraft to use the engine. Part of this was due to the mounting technique used for the XLR99: the engine was bolted onto a frame structure, which was then bolted into the engine compartment of the aircraft. A new frame was required to mount the two XLR11 engines, but the structural interface to the aircraft remained constant. However, the XLR11 used ethyl alcohol – water for fuel instead of the anhydrous ammonia used in the XLR99. This necessitated some modifications to the system, but none of them were major-fortunately, the two liquids had a

similar consistency and temperature. Surprisingly, no documentation describing the changes seems to have survived; however, as Scott Crossfield remembers:[134]

[S]ince the XLR11 engines were installed as two units including their own fuel pumps, the X – 15 needed only to supply the tank pressures to meet the pumps inlet pressure requirement and the engines didn’t know what airplane they were in. There were, of course, structural changes, i. e., engine mounting and I believe some ballast but nothing very complex. That is a relative statement. The difference in mixture would make the ideal fuel/lox load different but I don’t remember that was a significant problem.

forms, in the Bell X-1 series, Douglas D558-2, and Republic XF-91 programs at Edwards AFB. All of the engines used for the X-15 were made from leftover components from earlier programs. (NASA)

Charlie Feltz remembers that there were no modifications to the fuel tanks. North American had already built and sealed them by the time NASA decided to use the XLR11s. It was determined that both the metal and the sealant were compatible with alcohol, so there was no need to reopen the tanks. There were some minor changes to the plumbing and electrical systems to accommodate the new engines, along with cockpit modifications to provide the appropriate instrumentation and controls.-11351 Nevertheless, considering that North American had designed the airplane with no intention of installing anything but the XLR99, the changes were of little consequence and did not materially delay the program.-11361

In the final installation, the two engines were mounted on a single tubular-steel mounting frame attached to the airplane at three points. The mount canted the upper engine slightly nose-down and the lower engine in a slightly nose-up attitude so that their thrust vectors intersected at the airplane’s center of gravity.-11371

After the last XLR11 flight, NASA placed the remaining engines, spare parts, and special tools into long-term storage. Despite being almost 20 years old, the engines later found their way into the heavyweight lifting bodies.11381

The airframe proposals from Bell, Douglas, North American, and Republic arrived on 9 May 1955. Convair and Northrop evidently decided they had little to offer the competition. Two days later the various evaluation groups (the WADC, NACA, and Navy) received the technical data, and the results were due to the X-15 Project Office by 22 June.*451

In mid-May, Soule, as chair of the NACA evaluation group, sent the evaluation criteria to the NACA laboratories. The criteria included the technical and manufacturing competency of each contractor, the schedule and cost estimates, the design approach, and the research utility of each airplane. Each NACA laboratory had specific technical areas to evaluate. For instance, Ames and Langley were assigned to aerodynamics; Ames, the HSFS, and Langley to flight control; HSFS to crew provisions and carrier aircraft; and the HSFS and Lewis to the engine and propulsion system. Soule expected all the responses no later than 13 June, giving him time to reconcile the results before submitting a consolidated NACA position to the Air Force on 22 June. Later arrangements ensured that engine evaluations, also coordinated among the WADC, NACA, and Navy, would be available to the Research Airplane Committee on 12 July. The final evaluation would take place during a meeting at Wright Field on 25 July.*46*

Given the amount of effort that John Becker and the Langley team had put into their preliminary configuration, one might have thought that all of the contractors would use it as a starting point for their proposals. This was not necessarily the case. The Air Materiel Command had made it clear from the beginning that the Becker concept was "representative of possible solutions."

Becker agreed with this; he in no way thought that his was an optimal design, and the bidders were encouraged to look into other configurations they believed could meet the requirements.*47*

As it turned out, each of the four proposals represented a different approach to the problem, although to the casual observer they all appeared outwardly similar. This is exactly what the government had wanted—the industry’s best responses on building the new airplane. Two of the bidders selected the Bell XLR81 engine, and the other two chose the Reaction Motors XLR30. Despite this, all of the airplanes were of approximately the same size and general configuration. In the end, the government would have to evaluate these varied designs and determine which would most likely allow the desired flight research.

One of the unique items included in the X-15 design was a side-stick controller. Actually, the airplane included two side sticks: one on the right console for the aerodynamic controls, and one on the left console for the ballistic controls. The right and center controllers were linked mechanically and hydraulically to provide simultaneous movement of both sticks; however, the side stick required only one-third as much movement to obtain a given stabilizer motion.[182]

NASA had installed a similar side stick in one of the North American YF-107A aircraft to gain experience with the new controller. A review of early X-15 landing data (using the side-stick) revealed a "striking similarity" with landings made in the YF-107. Despite large differences in speed and L/D ratios, the variations in angle of attack, normal acceleration, pitching velocity, and horizontal stabilizer position exhibited the same tendencies for the pilot to over-control the airplane using the side stick. During the YF-107 program, several flights were generally required before a pilot became proficient at using the controller and could perform relatively smooth landings; the same was true of the X-15.[183]

Regarding the side-stick controller, Bob White commented that "the side aerodynamic control stick designed for the X-15 has received the usual critical analysis associated with a departure from the conventional." As pilots reported their experiences using the side stick, North American began making minor modifications to correct undesirable characteristics. In the end, the company found that most of the initial design features were satisfactory. The most frequent complaint was the location of the stick in relation to the pilot’s arm, since the stick had been located based on Scott Crossfield’s input, and other pilots differed in size and proportions. However, Crossfield was a strong proponent of the side stick and North American soon devised a way to adjust the stick into one of five different fore-aft locations prior to flight based on individual pilot preference. After this, the side stick gained favor rather quickly.-11841

The all-moving horizontal stabilizers deflected symmetrically for longitudinal control (elevators) and differentially for lateral control (ailerons). The rolling tail that had caused so much controversy within the government early in the program proved to be quite satisfactory in operation.

According to Bob White, "the pilot is not aware of what specific type of lateral control is allowing the roll motion. His only concern is in being able to get the aircraft response he calls for when

deflecting the control stick___ From experience to date [after 45 flights], the rolling tail has

provided a good rolling control for the X-15, and there have been no undesirable aircraft motions coupled in any axis because of lateral-control deflection."11881

Conventional rudder pedals actuated the movable portions of the dorsal and ventral vertical stabilizers. Just prior to the landing flare, the pilot would jettison the lower portion of the dorsal stabilizer to provide sufficient ground clearance; otherwise, the dorsal rudder would contact the ground before the landing skids. Speed brakes were located on each side of the fixed portion of the dorsal and ventral stabilizers. Irreversible hydraulic actuators actuated all of the aerodynamic control surfaces.11861

The aerodynamic controls were effective up to about 150,000 feet. Nevertheless, many X-15 pilots manually used the ballistic control system in addition to the aerodynamic controls above 100,000 feet, and the MH-96 on X-15-3 automatically began blending in the ballistic control system thrusters above 90,000 feet. As Neil Armstrong, who was a principle engineer on the MH- 96, commented, "a rule of thumb is that when dynamic pressure on control surfaces reduces to 50 psf, there should be a switchover from aerodynamic to reaction control." Despite some early concerns about controlling a vehicle above the sensible atmosphere, in practice it quickly became routine.11871

The Westinghouse-manufactured stability augmentation system (SAS) dampened the aerodynamic

controls in all three axes. The system consisted of three rate gyros, two pitch-roll servocylinders, one yaw servocylinder, and various electronics, displays, and controls. Essentially, the system included a channel for each axis that sensed the aircraft rate of change in pitch, roll, and yaw, and automatically provided signals to the respective servocylinders to move the horizontal and vertical stabilizers to oppose the airplane angular inputs. An additional interconnect damper, called "yar," provided a crossfeed of the yaw-rate signal to the roll damper. This interconnection was necessary for stability at high angles of attack, primarily because of the high roll input of the lower rudder. The yar interconnect was disabled when the lower rudder was removed during later flights. The authority of the SAS was equal to the pilot’s authority in pitch and yaw, and to twice the pilot’s authority in roll. The pilot could turn dampening on or off for each individual axis, and select the damping gain for each axis. Originally, the SAS gyro package was located in the instrument compartment behind the pilot. However, a vibration at high gains reported by Scott Crossfield during the first X-15 captive flight resulted in North American moving the gyros to the center of gravity compartment under the wings, thus removing the gyro from a point influenced by fuselage bending.-1188-

North American incorporated two side-stick controllers in the X-15 cockpit. The controller on the right console operated the aerodynamic flight control systems while the controller on the left operated the ballistic control system thrusters. The aerodynamic controller was mechanically linked to the conventional center stick. In X-15-3, the MH-96 adaptive flight control system automatically blended the ballistic thrusters in when needed, eliminating the need for the pilot to use the left side-controller. (NASA)

The SAS caused numerous pilot comments. During early flights below Mach 3.5, the dampers used moderate gains and the pilots quickly expressed a desire for "a stiffer aircraft," particularly in pitch and roll. North American subsequently increased the gain, resulting in generally favorable pilot opinions. It is interesting to note that at angles of attack above 8 degrees with low damper gain or with the roll damper off, pilots had great difficulty in controlling the lateral and directional motions to prevent divergence. This was primarily because of an adverse dihedral effect that was present above Mach 2.3. Although this was of some concern to the pilots, and the subject of a great deal of investigation by the researchers, the airplane exhibited acceptable handling characteristics as long as the dampers were functioning. In general, the airplane exhibited about the same handling qualities expected based on extensive simulations at Ames, and the pilots thought the damper-off handling was slightly better than the simulator predicted, but still considered the natural stability to be marginal.-1189

The SAS was unique for the time because it provided 10 pilot-selectable gain rates for each axis. However, the system experienced some annoying problems during development and early operations. During the first studies using the fixed-base simulator, the dampers sustained unwanted limit cycles (or continuous oscillations) from linkage lags and rate limiting. Pilots later observed the phenomenon in flight. The frequency of the limit cycle was about 3.2 cycles per second, resulting in changes in bank angle of about 1 degree. This limit cycle was not constant,

changed due to control input, and had a tendency to "beat." North American was unable to identify a way to eliminate the limit cycles, but modified the electronic filter to reduce its lag. This greatly lowered the amplitude of the limit cycles, and the pilots found the results acceptable.-190

The X-15 made extensive use of a stability augmentation system to dampen the aerodynamic controls in all three axes. The SAS was unique for the time since it provided ten pilot-selectable gain rates for each axis via rotary switches in the cockpit. Flight simulations showed that it would be nearly impossible for a pilot to control the X-15 in some flight regimes without the SAS. (NASA)

Although the modified filter greatly improved the issue with the limit cycles in roll, a new problem soon arose. It became apparent during ground tests that it was possible to excite and sustain a SAS-airplane vibration at 13 cycles per second with the modified filter. A breadboard of the modified filter was flown (flight 2-12-23) at higher damper gains, but Scott Crossfield failed to excite the vibration. During the rollout after landing, however, Crossfield encountered a severe vibration that required disabling the SAS. This experience led to the mistaken belief that the vibration could only occur on the ground. To prevent a recurrence, North American installed a switch that automatically lowered the gain whenever the pilot extended the landing gear.

However, five flights later (2-14-28), Joe Walker encountered a 13-cps vibration during reentry from 169,600 feet. After the flight, Walker reported that the vibration was the most severe he had ever encountered (or ever wanted to). The shaking was triggered by pilot inputs at 130 psf dynamic pressure and continued until the damper gain was reduced and the dynamic pressure climbed above 1,000 psf. Fortunately, the amplitude of the shaking was constrained by the rate limits of the control surface actuators. North American and NASA began investigating the problem again.-11911

TOTAL HOURS

TOTAL FLIGHTS

Failures of the stability augmentation system contributed to the maintenance woes suffered by the X-15 early in the flight program, but oddly, most of the failures were on the ground; the system seldom failed in flight. Nevertheless, an auxiliary stability augmentation system was added to the first two airplanes as insurance against an SAS failure. The X-15-3 did not carry an SAS or ASAS since the mH-96 adaptive flight control system performed both functions. (NASA)

The problem was that the lightly damped horizontal stabilizers were excited at their natural frequency (13 cps) by pilot inputs to the control system. The gyro picked up this vibration and the dampers were able to sustain the vibration with input to the control surfaces. Engineers also found a second natural frequency for the stabilizers at 30 cps. North American subsequently installed notch filters in the SAS and pressure feedback valves in the control surface actuators, eliminating the vibrations.-192

The SAS proved to be unreliable in the beginning, but fortunately most failures occurred during ground testing. The program recorded only seven in-flight failures during the first 78 flights (defined as NB-52 takeoff to X-15 landing). Of these failures, one was an electronic module, three were malfunctioning cockpit gain switches, and three were broken wires in the X-15. Engineers ultimately traced all except the failed electronics module to human error.192

There was never any doubt that the X-15 flight program would take place at Edwards AFB, California. However, Edwards would play a key role as infrastructure was developed to support the X-15. The program was an involved undertaking, and the operational support required was extensive. Logistically, Edwards would become the linchpin of the entire effort.

MUROC TO EDWARDS

The Mojave Desert-called the "high desert" because of its altitude-is approximately 100 miles northeast of Los Angeles, just on the other side of the San Gabriel Mountains. First formed during the Pleistocene epoch, and featuring an extremely flat, smooth, and hard surface, Rogers Dry Lake is a playa, or pluvial lake, that spreads out over 44 square miles of the Mojave, making it the largest such geological formation in the world. Its parched clay and silt surface undergoes a cycle of renewal each year as desert winds sweep water from winter rains to smooth the lakebed out to an almost glass-like flatness.-Ш-

Lieutenant Colonel Henry H. "Hap" Arnold decided that Rogers Dry Lake would make a "natural aerodrome," and in September 1933 the Army Air Corps established the Muroc Bombing and Gunnery Range as a training site for squadrons based at March Field near Riverside, California. It continued to serve in that capacity until 23 July 1942, when it became the Muroc Army Air Field. During World War II the primary mission at Muroc was to provide final combat training for aircrews before their deployment overseas.-12-

Until the beginning of World War II, the Army Air Corps conducted the majority of its flight-testing at Wright Field, Ohio. However, the immense volume of testing created by the war was one of the factors that led to a search for a new location to test the first American jet fighter, the Bell XP- 59A Airacomet. The urgent need to complete the program immediately dictated a location with year-round flying weather. In addition, the risks inherent in the radical new technology used in the aircraft dictated an area with many contingency landing areas, and one that minimized the danger of crashing into a populated area. After examining a number of locations around the country, the Army Air Forces selected a site along the north shore of Rogers Dry Lake about six miles away from the training base at Muroc.-13-

When Bell test pilot Robert Stanley arrived at the base in August 1942, he found just three structures: an unfinished hangar, a wooden barrack, and a water tower. Things would begin to change quickly as more than 100 people arrived at the base to support the project. On 2 October 1942, Stanley made the first "official flight" of the XP-59A (it had actually lifted off for the first time on the previous day during high-speed taxi tests), introducing flight-testing to the high desert. Only five years later, on 14 October 1947, Captain Charles E. "Chuck" Yeager became the first man to exceed (barely) the speed of sound in level flight when he achieved Mach 1.06 (approximately 700 mph) at 42,000 feet in the Bell XS-1 research airplane. Muroc’s place in the history books was firmly established.[4]

However, with the arrival of the X-1, flight-testing at Muroc began to assume two distinct identities. The Air Force typically flew the research airplanes, such as the X-3, X-4, X-5, and XF – 92A, in conjunction with the NACA in a methodical fashion to answer largely theoretical questions. The bulk of the testing, however, focused on highly accelerated Air Force and contractor evaluations of prototype operational aircraft, and was often much less methodical as they tried to get new equipment to combat units as quickly as possible at the height of the Cold War.[5]

Not surprisingly, the rather informal approach to safety that prevailed during the late 1940s, and even into the 1950s, was one of the factors that contributed to a horrendous accident rate. There were, of course, a number of other factors. The corps of test pilots at Muroc remained small and commonly averaged more than 100 flying hours per month. They flew a wide variety of different types and models of aircraft, each with its own cockpit and instrument panel configuration. Chuck Yeager, for example, reportedly once flew 27 different types of airplanes in a single one-month period. The year 1948 was particularly tragic, with at least 13 fatalities recorded at or near the base. One such fatality was that of Captain Glen W. Edwards, who was killed in the crash of a Northrop YB-49 flying wing on 4 June 1948. In December 1949 the Air Force renamed the base in his honor, while other pilots have streets named after them.[6]

Edwards AFB, California, hosted the X-15 flight program. The "new" main base complex is located at the center left in this photo, with the NASA Flight Research Center being slightly above the main base on the edge of the lakebed. Rogers Dry Lake was the planned site for all X-15 landings, and 188 times, it worked out that way. Two would land at Cuddeback, one at Delamar, four at Mud, one at Rosamond, one at Silver, and one at Smith Ranch; the X-15-3 broke up in flight and did not land on its last flight. (U. S. Air Force)

On 25 June 1951, the government established the Air Force Flight Test Center (AFFTC) at Edwards, and a $120 million master plan was unveiled for construction at the base. Part of the appropriation paid to remove the Atcheson, Topeka, and Santa Fe railroad from the northern portion of Rogers Dry Lake and bought out the silt mines that had been located along the route. However, the major undertaking was to relocate the entire base two miles west of the original South Base location and construct a 15,000-foot concrete runway. With the increased number of flight test programs at the base, the natural surfaces of the Rogers and Rosamond dry lakebeds took on even greater importance as routine and emergency landing sites. The first AFFTC commander, Brigadier General Albert Boyd, later commented that the dry lakes were nothing less than "God’s gift to the U. S. Air Force." That same year, the USAF Test Pilot School moved from Wright Field to the high desert.[7]

Bell would have seemed a logical choice to develop the new research airplane since the company had developed the X-1 series and X-2 high-speed research aircraft that had ushered in a new era of flight research. They were also doing studies on much faster vehicles in search of the BoMi boost-glide bomber. The company had direct experience with advanced heat-resistant metals and with the practical issues of powering manned aircraft using liquid-fueled rocket engines. In fact, Bell had an in-house group that built rocket engines, including one under consideration for the X-15. Lawrence Bell, Robert Woods, and Walter Dornberger were already legends. Somehow, all of this was lost in the proposal.[48]

Unsurprisingly, Bell engineers decided the Bell-manufactured XLR81 was the most promising engine, and it became the baseline; however, the XLR30 offered certain advantages and Bell proposed the alternative D-171B variant using this engine. The design had three XLR81s arranged in a triangular pattern with one engine mounted above the others, much like the later Space Shuttle Orbiter. Bell believed that the ability to operate a single XLR81 at its 8,000-lbf "half­thrust" setting was an advantage, based on a reported comment from the NACA that "a high percentage of the flight testing would be conducted in the lower speed and altitude ranges." Bell did not record who made the comment, but given that only 36 of the eventual 199 X-15 flights were below Mach 3, it was obviously incorrect. Unfortunately, it seemed to influence the Bell proposal throughout.-^

A throttle lever controlled engine thrust by actuating a series of switches arranged so that thrust increased as the pilot pushed the lever forward in the conventional manner. The initial switch fired the first engine at its 8,000-lbf half-power setting. The second switch caused this engine to go to 14,500-lbf full power. The next switch fired the second engine at its 14,500-lbf setting, resulting in a 29,000-lbf thrust. The last switch started the third engine, resulting in a full thrust of 43,500 lbf. The engineers did not consider the slightly asymmetrical thrust provided by the triangular engine to be a problem.[50]

The selection of a conventional aerodynamic configuration simplified the arrangement of the fuselage and equipment systems. The fuselage had six major sections. The forward section contained the pilot’s compartment, nose gear, and research instrumentation, followed by the forward oxidizer tank. A center section housed the wing carry-through, main landing skids, and pressurization systems, followed by the aft oxidizer tank and fuel tank. The aft section contained the engine and empennage. A pressurized area just behind the cockpit contained the hydraulic and electrical systems, environmental control equipment, and research instrumentation. The hydrogen peroxide supply, the main landing gear, and the structure for suspending the research airplane from the carrier aircraft were located in the center of the fuselage between the two oxidizer tanks. A flush-mounted canopy minimized drag and avoided discontinuities in the airflow that could result in thermal shocks on the glass.[51]

One of the unfortunate consequences of selecting the XLR81 was that the red, fuming nitric acid required a large storage volume, which caused the oxidizer to be stored in two tanks (one on either side of the wing carry-through). This was necessary to maintain the center of gravity within acceptable limits, but complicated the attachment of the wing to the fuselage. Bell investigated bolting the wing directly to the oxidizer tank or passing the structure through the tank. This, however, was not considered ideal "since it would present a hazard in the form of a possible fatigue failure as the result of the combination of localized wing loads and tank pressurization loads." The 61S-T aluminum propellant tanks were generally similar to those used on the Bell MX – 776 (GAM-63) RASCAL missile program.[52]

The wing had a leading-edge sweep of 37 degrees to moderate center-of-pressure shifts at subsonic and transonic speeds. Engineers had discovered that higher sweep angles resulted in pitch-up and damping-in-roll difficulties that Bell wanted to avoid. At the same time, researchers found that the aspect ratio was not particularly important, so it was set to provide decent subsonic and landing attitudes. The total wing area was 220 square feet, allowing a reasonable landing speed of 170 mph.1531

Approximately one-third of the vertical stabilizer area was located under the fuselage to maintain high-speed stability. This ventral stabilizer was added "to provide sufficient directional stability to M=7.0. This lower surface is very effective at high Mach numbers because of the compressive flow field below the wing." Bell attempted to provide as much area as possible while still maintaining sufficient clearance for the D-171 to be loaded into the carrier aircraft without resorting to a folding or retractable design. Before the airplane could land, the pilot would jettison the ventral stabilizer to provide sufficient clearance for the landing gear. A parachute lowered the ventral to a safe landing, although Bell noted that deleting the parachute would save a little weight, with the ventral becoming expendable.-1541

Landing skids were a logical choice to save weight but the exact nature of these skids was the subject of some study. A two-skid arrangement-one forward and one aft—was considered too unstable during landing, although a drag chute could be used to overcome this, as was done on the SM-62 Snark missile. Still, the arrangement was undesirable. A nose wheel with a single aft skid was statically stable, but model tests showed that it was dynamically unstable. A good pilot could land the aircraft with this arrangement, but Bell rejected the configuration because it placed too placed a great burden on the pilot. Two forward skids and a single aft skid offered neutral stability, but experience with the Sud-Est SE5003 Baroudeur showed that it still placed a high burden on the pilot. Bell finally selected a conventional tricycle arrangement with a nose wheel and two main skids located midway aft on the fuselage. Both the nose gear and skids were retractable and covered with doors, unlike the eventual X-15 where the rear skids did not retract inside the fuselage.1551

The fully loaded airplane weighed 34,140 pounds at launch, including 21,600 pounds of propellants. The estimated landing weight was 12,595 pounds. Based on a launch at Mach 0.6 and 40,000 from a B-36 carrier aircraft, Bell estimated that the D-171 could exceed the basic performance requirements. The projected maximum altitude during the "space leap" was 400,000 feet. At altitudes between 85,000 and 165,000 feet, the velocity was in excess of 6,600 fps, with a maximum of 6,850 fps at 118,000 feet.1561

A set of reaction controls used eight hydrogen peroxide thrusters: one pointed up and another down at each wing tip for roll control, one up and one down at the tail for pitch control, and one pointing left and one right at the tail for yaw control. A single control stick in the cockpit controlled the thrusters and aerodynamic control systems. Bell noted that "no criteria are available for the design of such controls," so the company arbitrarily assumed that aerodynamic controls would be ineffective at dynamic pressures below 10 psf. Bell expected the X-15 to operate in flight regimes that required reaction controls for about 115 seconds per high-altitude mission, and provided 550 pounds of hydrogen peroxide. Operating all of the thrusters for the entire 115- second flight (something that obviously would not happen) used only 49% of the available propellant.1571

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The Bell entry in the X-15 competition bore a subtle resemblance to their X-2 research airplane that had such an unhappy career. Bell had considerable theoretical experience with thermal protection systems as part of its ongoing work on the Air Force BoMi and RoBo programs, and much practical experience with high-speed X-planes such as the X-1 and X-2. Ultimately, the Bell proposal finished third in the competition. (Bell Aircraft Company)

The researchers at Bell did not believe the hot-structure data provided by the NACA from the Becker studies. This may have reflected a bias on the part of Bell engineers who had been working on alternate high-speed structures for several years. The Bell proposal contained a detailed discussion on why conventional or semi-conventional structures would not work, and the hot – structure concept fell into the latter category.

A survey of available materials showed that Inconel X was the best available high-temperature alloy for a conventional structure—the same conclusion reached at Langley. Bell estimated that an Inconel X airframe would weigh approximately 180% as much as an equivalent structure made from aluminum 75S-T. Bell noted that the "usual expedient" of adding additional material would not relieve all of the thermal stresses unless sufficient material were added to absorb the entire expected heat load, leading to a structure that would be too heavy to accomplish its assigned mission. The Bell engineers also thought that "the stresses and deformations produced by temperature gradients cannot generally be reduced by the simple addition of more material."[58]

The second approach was to use what Bell called semi-conventional structures. In addition to adding sufficient material to absorb the heat load, the designers attempted to develop structures that would be free to warp and bend as they heated. Bell believed that all of the design approaches they tried would fail in operation. For instance, Bell designers decided it would be impossible to use integral propellant tanks in a hot-structure airframe because "no suitable

structural arrangement has been found for attaching propellant tank ends and baffles to the outer shell without introducing serious thermal stresses." When they investigated the use of separate tanks, they found the weight penalty to be severe.

Bell also briefly investigated actively cooled structures, such as the "water wall" concept developed early in the BoMi studies. The basic structure weighed little more than a conventional aluminum airframe, but including the weight of coolant and pumping equipment resulted in the concept being 200-300% heavier.-1591

In a fuzzy look at things to come for the Space Shuttle, Bell investigated a structure protected by external insulation and concluded that "[c]eramic materials would seem attractive for insulation, except that the present state of development for this application is not well enough advanced…."*69!

As it turned out, Bell had an alternative, developed during the ongoing BoMi studies. This unique double-wall structure used air as an insulator, permitting heat transfer by radiation in addition to conduction. The outer wall consisted of a 0.005-inch-thick Inconel X skin panel, approximately 4 inches long and 8 inches wide, welded to a corrugated sheet of Inconel X. The corrugations were

0. 3125 inch deep with 0.3125-inch spacing. An outside retaining strip of Inconel X (approximately 1.25 inches wide and 0.056 inch thick), running along each edge, held each panel in place. The edges of the corrugations, top and bottom, were joggled 0.056 inch so that the outer surface was flush. In the bottom, joggled portion of each of the corrugations, 0.015-inch – deep protruding dimples provided support for the outer wall panels to the inner structure. The combination of the dimple and joggle raised the outer wall panel to a height slightly over 0.375 inch from the inner structure, providing the necessary air space for insulation. The retaining strip was broken into 4-inch lengths to permit expansion relative to the inner structure, and two screws and two floating inverted-type anchor nuts held each retaining strip to the structure.

These provided the required air space between the inner and outer walls to minimize heat conduction into the inner structure. Narrow strips of fibrous insulation located beneath the retaining strips prevented boundary air from leaking between the outer panels and their retaining strips.1611

This arrangement allowed the outer wall panels to expand in the direction parallel to the corrugations simply by sliding further under the retaining strips. Separating the skin into elements only 4 inches wide accommodated the thermal expansion of the outer skin of the outer wall. In order to prevent the parallel, free edges of this very thin skin from lifting due to aerodynamic forces, "Pittsburgh" joints interconnected the edges of adjacent panels. This is a standard sheet-metal joint, but in this application "considerable clearance" was used so that the adjacent panels were free to move relative to one another to permit thermal expansions.*621

Two pins set in the basic structure restrained each of the 4-by-8-inch outer wall panels against lateral movement. One of these pins fit snugly into a hole in a small square plate welded to the bottom of two adjacent corrugations, thus preventing any translations. The other pin fit into a slotted hole, permitting expansion but preventing rotation. Thus the outer wall had complete freedom of expansion relative to the underlying aluminum alloy structure. Its shallow depth (0.3125 inch) and uniformity minimized thermal gradients through the wall. Although they cost considerably more to manufacture, Bell proposed using Haynes 188 or similar alloys in areas where temperatures exceeded the capability of Inconel X. Researchers expected that ceramic panels or various sandwich materials could eventually replace the Inconel outer wall.*631

The primary advantage of the double-wall system was that it weighed some 2,000 to 3,000 pounds less than an Inconel X hot structure. The double-wall construction also minimized development time, according to Bell, since the primary structure of the airframe was conventional in every way, including its use of aluminum alloys. This limited, in theory, any development problems for the outer wall. Interestingly, Bell believed that the double-wall construction provided an advantage when it came to research instrumentation. Since the outer panels were easily removable, it greatly simplified the installation of thermocouples, strain gages, pressure orifices, and other sensors.[64]

The wing and empennage used the same double-wall construction, but the leading edges were of unique construction. Bell noted that "it cannot be assumed that the optimum design has been selected since the evaluation…requires a greater time than afforded in this proposal period." Bell engineers did not believe they could accurately predict the heat transfer coefficients, but noted that the equilibrium temperature of the leading edges could approach 2,500°F. At this temperature, Bell was not sure that any metallic alloy would be sufficient, or whether a ceramic was necessary instead. Nevertheless, Bell proposed a metal heat sink. A 0.040-inch-thick Inconel X shell formed the desired leading-edge shape with a chord-wise dimension of approximately 6.5 inches (normal to the leading edge). Properly spaced, welded ribs provided attachment fittings, and intermediate ribs provided support to ensure that air pressure would not deform the shell. Lithium, beryllium, magnesium, or sodium (listed in descending order of preference) filled the leading edge shell as a heat sink.*651

All of the leading edges were easily removable, facilitating the substitution of various types of leading-edge designs for flight research and evaluation. The wing leading edges were single­piece structures on each side of the airplane. The inboard attachment was fixed, but the other attach points were designed to allow span-wise motion to accommodate differences in linear expansion between the wing structure and the leading edge.[66]

At first, Bell selected a Boeing B-50 Superfortress for its carrier aircraft, mainly because it had experience with this type of airplane from the X-1 and X-2 programs. It soon became apparent, however, that the B-50 did not have the capability to carry the D-171 and its support equipment to the altitudes required. Attention then turned to the Convair B-36. A comparison of the two aircraft showed that the B-36 had a much better rate of climb, and could launch the D-171 at Mach 0.6 and 40,000 feet compared to Mach 0.5 and 30,000 feet for the B-50.*67

The basic installation in the B-36 was straightforward, and Convair already had data on the B-36 carrying large aircraft in its bomb bays from Project Fighter Conveyer (FICON).*68 Loading the D – 171 was the same as loading the X-1 or X-2: a pair of hydraulic platforms under the B-36 main landing gear allowed the ground crew to tow the research airplane underneath the raised bomber. Alternately, the bomber straddled an open pit in the ground and crews raised the research airplane into the bomb bays. The D-171 took up the forward three of the four B-36 bomb bays in order to keep the mated center of gravity at an acceptable position. This also minimized B-36 control problems when the D-171 dropped away from the bomber.-*69*

As had been the case with previous research airplanes, the mated pair would take off with the research airplane pilot in the carrier aircraft—not in the D-171. As the carrier climbed through 15,000 feet, the pilot would climb into the research airplane and the canopy would close. Equipment checks of the research airplane

would begin as the carrier climbed through 35,000 feet. When the checks were completed, the carrier aircraft would drop the research airplane.*701

Along with the baseline D-171 design, Bell proposed two slight variations. The D-171A two-seat version was a required response to the government request for proposal. Bell noted that that since the equipment compartment had a differential pressure of 2.5 psi to support the instrumentation, a small increase in structural weight would allow the higher pressure differential necessary to carry a second crew member. The observer would be seated on an upward-firing ejection seat and have two small side windows in a separate canopy. The gross weight was unchanged at 34,140 pounds since the weight of the observer and the ejection seat exactly matched the research instrumentation load normally carried. Performance was also unaffected because the propellant load was identical.1711

The second variant was the D-171B powered by a Reaction Motors XLR30 "Super Viking" engine. Although Bell preferred to use three XLR81 engines, it realized that the XLR30 offered some advantages. The D-171B had an empty weight about 200 pounds more than the baseline configuration, but a launch weight of some 1,000 pounds less. Bell listed the fact that the XLR30 used liquid oxygen as its oxidizer as its greatest disadvantage since this would require a top-off system in the carrier aircraft, which Bell believed would add "considerable greater weight" to the B-36.172 Bell also thought that the minimum thrust capability of the XLR30 (13,500 lbf) was unsatisfactory compared to the Hustler engine (8,000 lbf). On the positive side, the internal propellant tank arrangement for the XLR30-powered airplane was superior because only a single oxidizer tank would be needed, greatly simplifying propellant management for center-of-gravity control. Bell agreed that the single XLR30 thrust chamber (versus three for the XLR81 installation) was also an advantage. Although no two-seat XLR30 aircraft was described in the proposal, it is easy to imagine a two-seat variant since the forward fuselage was identical to that of the D-

171.IZ3]

Bell expected to have the basic design established six months after the contract was signed, and to finalize the design after 18 months. The first airplane would be available for ground tests 34 months after the start of the contract. Bell indicated that they attempted to compress the schedule into the required 30 months, but were unable to do so. It would take 40 months to get to the first glide flight, and six additional months before the first powered flight. Bell expected the government to provide a complete test engine in the 27th month, and a final propulsion system had to be delivered to Bell simultaneously with the first aircraft entering ground tests.1741

The X-15 landing gear was somewhat unusual, both in its approach and in its simplicity. The system consisted of a dual nose wheel and a pair of aft skids. Initially the cast magnesium nose wheels were fitted with standard aircraft tires pressurized with 240-psi nitrogen. The skids consisted of a 4130-steel skid and an Inconel X strut that was attached to the fuselage by trunnion fittings and through bell crank arms that were attached to shock struts inside the aft fuselage. The skids were free in pitch and roll, but fixed in yaw for parallel alignment. Drag braces attached to the fuselage ahead of the trunnion fittings and to the skids at the strut attachment pin. Bungee springs kept the skid in a nose-up position just before landing. Instead of retracting inside the fuselage, the skids and struts folded forward against the outside of the fuselage when retracted. The pilot lowered the landing gear by pulling a handle in the cockpit that attached via
cable to the uplock hooks and released the gear. North American designed the landing gear for an 11,000-pound airplane with a sink rate of 9 fps, touching down between 190 and 230 mph at an angle of attack of approximately 6 degrees.-1124

Three major test series of the landing-gear system were conducted prior to the first glide flights: 1) a dynamic-model test of stability during the landing run, 2) nose-wheel shimmy tests using the actual nose gear, and 3) full-scale skid tests at the lake-bed landing site.-1125

North American used the model tests to investigate the stability of the tricycle arrangement. Engineers constructed a 1/10-scale model that accurately reflected the size, weight, and mass moments of inertia for yaw and roll, but did not simulate the aerodynamic characteristics of the X-15. Scale-size metal skids were manufactured so that they could be installed in either the original mid-fuselage location proposed by North American, or the aft fuselage location eventually built. North American catapulted the model along a concrete runway using a 100-foot length of 0.625-inch-diameter shock cord. High-speed movie cameras on overhead towers recorded each run. The tests revealed some minor nose-wheel instability, which the company subsequently corrected.-1124

Researchers at Langley then tested the revised full-scale nose gear using the landing-loads track facility at speeds up to 125 mph. These tests evaluated the nose gear on smooth concrete, uneven concrete, wet pavement, sandy pavement, uneven tire pressure, one flat tire, and unbalanced wheels. Given that the X-15 was to land only on dry lake beds, some of the tests seemed extreme. Throughout the tests the co-rotating wheel arrangement proved extremely stable, with no tendency to shimmy. Researchers, therefore, concluded the shimmy damper and torque links were unnecessary and North American subsequently removed them, saving 25 pounds.-1127

North American conducted the landing-gear-skid tests on Rogers Dry Lake during April 1958. For these tests, researchers mounted the complete main gear on a two-wheel trailer vehicle and towed it behind a truck at speeds up to 70 mph. After the truck reached full speed, an electric switch actuated a bomb-release solenoid that dropped a 6,000-pound load on the skid landing gear. Instruments on the gear recorded vertical and drag loads, and shock-strut position. High­speed cameras mounted in the truck and trailer recorded the motion of the gear and skids. Test runs included straight-line landing on smooth lake surfaces, "fishtail" runs on rutted and bumpy areas near the edges of the lake, and one landing on the concrete runway just to make sure. The results of all the tests were satisfactory. Skid wear on the lake beds was light, and engineers determined that the skids would last for three or four landings. The tests revealed that the X-15 should leave depressions approximately 0.03 inch deep in the lake bed. As expected, wear on the concrete runway was severe, but the tests showed the X-15 could land on concrete if necessary.-1124

Despite all the tests, the first four actual landings pointed out several deficiencies in the landing gear, mainly because the aircraft was heavier than anticipated and sink rates were slightly higher. North American replaced the shock struts with higher-capacity units, and strengthened some of the structure inside the fuselage. The fourth landing resulted in X-15-1 breaking in half. This was not strictly a design error; Scott Crossfield had been unable to fully jettison the propellants prior to an emergency landing, and the airplane was significantly overweight. However, the landing gear contributed because the gas and oil mixture in the shock strut foamed, keeping the rear skids from absorbing as much of the impact as they should have. This forced a higher than normal load on the nose gear, aggravating the structural problem caused by being overweight.729

The X-15 was unique, even among X-Planes, in using a landing gear consisting of rear skids and a nose wheel. The skids solved several problems for designers since they were relatively small and could be stowed mostly outside the airframe. Interestingly, the X-15 landing gear was lowered by the pilot pulling a mechanical handle that was connected to a cable that released the uplock hooks and allowed a bungee to extend the skids. A similar system would have been used on the X-20 Dyna-Soar if that program had not been cancelled. (NASA)

In addition, during some of the early landings, engineers found that the nose wheel tire marks left on the dry lake bed were not continuous. After initial contact, the tire marks became very faint or disappeared for short distance and then reappeared. This puzzled the engineers since all of the early drop tests of the landing gear had been satisfactory.-1^00

The engineers became concerned that the nose-gear extension mechanism was not working properly. Normally, technicians manually retracted the nose gear after attaching the X-15 to the NB-52, and then they pumped dry nitrogen gas into the shock strut to preload it to 1,404 psi. Charlie Feltz had suggested this method as a way to minimize the size and weight of the nose gear compartment. What the engineers discovered was that upon lowering the landing gear, an orifice in the strut trapped the nitrogen gas below it and most of the shock-absorption oil above it. The design of the metering valve was such that it prevented a rapid change in position of the oil and nitrogen in the 10 seconds between gear extension and wheel touchdown. To better understand the problem, engineers conducted additional dynamic tests using the original test apparatus. Initial tests operated the apparatus with the nose gear serviced in the extended position, as had been done in the original tests. The performance appeared normal. The engineers then modified the test rig to allow the gear to be serviced in the retracted position, as was done on the airplane. A delay of 10 seconds was introduced between the gear being lowered and touching down, and the abnormal behavior was reproduced almost exactly.-1201

At first, engineers modified the orifice in the shock, but this failed to resolve the problem. After additional tests, the engineers determined that they could not pressurize the strut in its retracted (compressed) position. Unfortunately, the nose wheel compartment was not large enough to allow the nose gear to be retracted in its extended position. The final solution was to mount redundant nitrogen bottles on the gear strut itself. When the gear reached its fully extended position, a valve actuated and released the nitrogen to pressurize the strut. This worked and the first modified nose gears were available in July 1960. However, the engineers kept evaluating the problem and, later in the program, changed the design again. This time they installed a floating piston inside the strut that kept the oil and gas separated. Technicians could now pressurize the strut in the compressed position before flight, allowing the removal of the nitrogen bottles.-1202

During 1961, engineers instrumented the skids to gather additional data on skid landing gear in support of the Dyna-Soar program and possible future vehicles, such as the space shuttle. Standard NASA instrumentation was used to provide airplane upper-mass response, shock-strut force and displacement, main – and nose-gear drag forces, nose-gear vertical force, horizontal – and vertical-stabilizer setting, horizontal stabilizer load, airplane angle of attack, and airplane pitch velocity during the impact and slideout portion of a landing. Tests were conducted at the end of normal research flights while the pilots landed normally and performed specific control movements during slideout. Phototheodolite cameras on the ground furnished data for landing coordinates, airplane altitude, flight-path velocity, and vertical velocity at touchdown. The instrumentation remained on all three airplanes for the remainder of the flight program to monitor the severity of each landing.-1203

Landing-gear loads continued to be high, despite the minor modifications made early in the flight program. An analytical study of the landing dynamics showed that several important parameters affecting the landing loads were actually aerodynamic factors. One of the primary culprits was a down-load from the horizontal stabilizer caused by both the pilot and SAS. Immediately prior to touchdown, the stabilizer trim position was set to between 4 and 5 degrees with the leading edge down. If the pilot pulled back on the stick and put the leading edge further down, the landing loads increased. If the pilot pushed the stick forward to get the leading edge up, the loads decreased. Another factor affecting the gear loads was lift from the wing. Unfortunately, the severe nose-down angle of the X-15 after nose-gear touchdown effectively pushed the airplane into the ground, further increasing the stress on the landing gear. Unfortunately, this was an unchangeable consequence of the airplane configuration, and a similar problem occurred on the space shuttle orbiters.[204]

The most severe problem, however, was weight. The design landing weight had been 11,000 pounds. The initial landing weight of the airplane was 13,230 pounds, and by 1965 this had crept up to 15,500 pounds on a routine basis. Emergency landings with a partial propellant load could be as high as 17,000 pounds. The only way to execute safely a landing at 17,000 pounds was for the pilot to perform an active push maneuver to obtain low horizontal stabilizer settings. This would still exceed the design load on the airplane, but would most probably be below the yield (destructive) limit.-1203

The nose gear was more conventional, consisting of a pair of wheels and tires. Note how short the nose gear strut is, resulting in severe loads during landing. The length of the nose strut was largely dictated by the amount of room available to stow it when retracted. Space shuttle orbiters suffer from a similarly short nose gear strut. (NASA)

By 1965 the problem was no longer one of understanding the nature of the loads, but rather one of how best to reduce them. North American introduced a near-constant series of minor modifications to the skids, their struts, and the surrounding structure in an effort to provide additional margin for the landing gear. Of all the factors that affected gear loads, the most

difficult to control—without restricting the research role of the airplane—was weight. Engineers determined they could reduce landing gear loads if they prevented the stabilizer angle from moving in the leading-edge-down direction during landings. Training the pilots to perform a push maneuver during landing accomplished this. In addition, North American installed a switch in the cockpit that disengaged the SAS at main gear touchdown to prevent the dampers from forcing the stabilizer leading edge down. Experience showed that under normal circumstances the pilots were efficient at pushing the stick at the right moment, even though the maneuver had to occur within 0.4 second after main gear touchdown to be effective in reducing gear loads. However, this maneuver was unnatural for the pilots, who tended to revert to habits formed through long hours of previous experience during emergencies and pull back on the stick. For this reason, the FRC began developing an automatic stick-kicker.-1206

In fact, this very condition occurred during the Jack McKay’s accident in X-15-2. The airplane was 1,000 pounds heavy with residual propellants, and as he landed, McKay pulled back on the stick, driving the stabilizer leading edge down to its maximum value. As it happened, the flaps failed on this flight and resulted in a down-load on the main wing, and therefore on the main landing gear. The combined resulted was a severely overstressed gear that, of course, failed.-207

Following the accident with X-15-2, engineers considered designing a new landing gear for the modified X-15A-2. The original location of the nose gear was approximately 23 feet ahead of the center of gravity, and moving the landing gear back could significantly reduce main-gear loads, with the forward bulkhead of the liquid-oxygen tank representing the rear-most location in the existing airframe. One of the ideas engineers investigated was moving the nose landing gear rearward to the instrumentation compartment behind the pilot. The nose gear would occupy the lower half of the compartment, with most of the instrumentation that normally resided there being moved forward to the old nose-gear compartment ahead of the pilot.208

However, fiscal and schedule constraints involved with repairing the aircraft precluded such major modifications, and the existing gear locations were reused on the modified airplane. Nevertheless, engineers made some basic changes, such as increasing the shock strut stroke from 3.66 inches to 5.03 inches, and modifying the relief valve setting from 17,000 pounds to 22,000 pounds.

North American manufactured two sets of strengthened struts—one set that was the same length as the original units, and another set that was lengthened from 53.6 inches to 59.0 inches. The longer units provided sufficient ground clearance to land with the functional ramjet sill attached to the ventral, but it appears that all flights of the X-15A-2 used the shorter units. Engineers also lengthened the skid 6.75 inches. In addition, engineers made some changes to the nose gear, primarily increasing the shock strut stroke to accommodate the increased length of the airplane. North American lowered the trunnion 9 inches to allow an attitude at nose-gear touchdown similar to that of the basic X-15. Despite these changes, the landing dynamics of the new gear were not appreciably changed, and X-15A-2 inherited most of the deficiencies of the basic system.209-

In addition, during the first part of 1965, North American investigated increasing the capability of the X-15A-2 gear. NASA wanted the maximum landing weight with the "short" main landing gear to increase to 16,374 pounds normal and 18,519 pounds emergency. The "long" gear used with the ramjet would increase to 17,855 pounds normal and 20,000 pounds emergency. A preliminary analysis indicated that incorporating the stick-kicker and changing the shock-strut relief valve setting would allow these increases. However, at 20,000 pounds there were concerns about whether the fuselage structure just behind the cockpit would be strong enough.219

Researchers installed a prototype stick-kicker in the FRC fixed-base simulator in May 1965 to determine the optimum stick forces. Subsequently, engineers installed the first stick-kicker in X – 15-3 during the weather down period at the beginning of 1966, and in X-15-1 by the end of that year. Apparently, NASA never installed the stick-kicker in X-15A-2. An emergency landing at 17,700 pounds, the highest landing weight yet encountered by the program, illustrated the effectiveness of the stick-kicker.-211

North American also conducted an investigation during early 1965 to determine the modifications needed to increase the landing weight of X-15-1 and X-15-3 to 16,000 pounds normal and 17,000 pounds emergency. The analysis included the use of the stick-kicker to rotate the horizontal stabilizer at landing to reduce main skid loads, although this would not eliminate the need to modify the skids for the higher weights. Preliminary studies showed that relocating the nose gear trunnion (as done on X-15A-2) would appreciably reduce landing loads on the other two airplanes, even without the addition of a stick-kicker.-212-

At the same time, engineers studied the feasibility of incorporating a third main skid attached to the fixed portion of the ventral stabilizer. This third skid could redistribute the landing loads and relieve the critically stressed gear components, particularly if either the stick-kicker of the landing flaps failed to operate. NASA installed the skid on X-15-3 in time for flight 3-52-78 on 18 June 1966, and by the end of 1966 it had used the third skid for four landings. These landings, however, were not at a sufficient weight to require the skid, and during the slideout the third skid contacted the lake surface with little or no load applied to it. Nevertheless, the third skid seemed like a good idea and NASA modified X-15-1 in time for flight 1-71-121 on 22 March 1967.

NASA did not install the third skid on X-15A-2 since it would have interfered with the ramjet installation.215-

The X-15A-2 experienced some of the more bizarre problems with landing gear. On the second flight (2-33-56) of the modified aircraft, after obtaining a maximum Mach number of 5.23, the nose gear unexpectedly extended as the airplane decelerated below Mach 4.2. William P. Albrecht, the X-15 project engineer for the flight, wrote that "[u]pon arrival in the Edwards area, chase aircraft confirmed that the nose gear was extended fully, and that the tires appeared badly burned, although still inflated. Major Rushworth elected to land the X-l5, and skillfully did so. The tires remained intact on touchdown but disintegrated after approximately 300 feet of rollout, the remainder of the 5,630 foot rollout being taken by the magnesium rims of the nose wheels." Considering the circumstance, it was a good landing.-214!

The subsequent investigation revealed that the nose-gear uplock hook was severely bent, the point of the hook having opened by approximately 0.25 inches. However, engineers determined the hook had not bent far enough to release the gear without the occurrence of some other deflection. The pilot lowered the X-15 landing gear via a simple cable arrangement that connected the landing gear extension handle in the cockpit to the uplock hook. Engineers measured the slack in the landing gear actuating cable (used to compensate for fuselage expansion due to heating effects) at 1.18 inches after the flight, within the specified limits. However, an analysis by North American indicated that the thermal growth of the fuselage was approximately 1.90 inches for this flight. This pointed out that the slack allowance was inadequate. Since the same mechanism operated all three landing gear components, it could not be ascertained in advance which of the three landing gears (left main, right main, or nose) would be first affected by partial actuation of the extension system, since that one with the least cable loading (due to friction, air loads, etc.) would tend to operate first. NASA duplicated the failure in the High-Temperature Loads Calibration Laboratory by simulating the fuselage expansion and applying heat to the nose-gear door. As Albrecht observed afterward, "Needless to say some modification to the landing gear mechanism seems to be in order."-215-

North American modified the cable to provide 2.25 inches of slack to compensate for thermal expansion. Although the engineers did not believe the problem affected the other two airplanes, they also received the modification. The only major drawback to this modification was that the pilot now had to pull the gear handle through almost 14 inches of travel to release the landing gear, which led to several complaints. Subsequently, engineers at the FRC designed a differential pulley that shortened the pull to 11 inches.-1216

These modifications, however, did not totally fix X-15A-2. During the next flight (2-34-57) on 29 September 1964, Bob Rushworth experienced a similar, but less intense, noise and aircraft trim change at Mach 4.5: the small nose-gear scoop door opened. This had already happened several times during the flight program on all three airplanes, fortunately without disastrous results.

There were two initial thoughts on how to fix the problem. The first was to eliminate the scoop door altogether; except for inspection and servicing, the door would be bolted shut prior to flight. Alternately, engineers could design a new uplock for the scoop door that featured a positive retention of the door roller on the uplock hook. In the end, NASA selected the second route and installed a new uplock hook, scoop door hook, and associated bell cranks.-217

NASA conducted two captive-carry flights of X-15A-2 to verify proper deployment of the redesigned nose scoop door and nose landing gear after cold soak. During flight 2-C-58 the nose gear required approximately 5.4 seconds to lock down—an unacceptably long time. Subsequent inspection showed that an incorrect orifice had been installed in the nose-gear snubber (which controlled the deployment rate). NASA installed the correct orifice, and the deployment time on flight 2-C-59 was an acceptable 2.7 seconds. Researchers collected data on both these captive flights data regarding the scoop door hook position and scoop door roller loads. Hook movement was negligible (less than 1/16 inch) and NASA subsequently modified the other two airplanes as well. Jack McKay took X-15A-2 on a perfect flight (2-35-60) on 30 November 1964.

Like the rear skids, the nose wheel was lowered by the pilot pulling a handle that was connected to a cable that released the uplocks. On two separate flights, the nose gear extended while Major Robert A. Rushworth was flying the X-15A-2 above Mach 4, resulting in some interesting flying characteristics and two sets of burned tires. Researchers finally deduced that the fuselage of the airplane was expanding due to heat, and that the landing gear release cable did not have enough slack to compensate. North American increased the slack in the cable, but the pilots now had to pull the release handle more than 14 inches to get the landing gear to deploy. (NASA)

However, it did not end there. Rushworth was in the cockpit again for the next fight (2-36-63) of X-15A-2 on 17 February 1965 when the right main skid extended at Mach 4.3 and 85,000 feet. The chase pilot was able to verify that the gear appeared structurally sound, and Rushworth managed to make a normal landing. Investigation of the right-hand main skid uplock revealed that thermally induced bowing of the main strut caused excessive loading of the main uplock hook. Ground heating tests of the main-gear struts during a "hot-flight" profile caused bending of the hook and release of the gear. Consequently, NASA modified the main-gear uplock to include a stronger hook, a Belleville washer mounting system to accommodate approximately 0.14 inch bowing of the strut, and a stronger support structure. In addition, it was necessary to reinforce the sheet-metal fuselage longeron structure around the main-gear drag-brace anchor fittings. While the repair itself was not complicated, access was extremely difficult since it

To test the hypothesis that the fuselage expanded more than the release cable, researchers at the Flight Research Center heated one of the X-15 forward fuselages using heat lamps. The test confirmed the theory. (NASA)

This ended the significant problem with the landing gear on the X-15A-2 (and the other airplanes), although the ever-increasing landing weight continued to be a concern and a set of small modifications (such as stronger struts) continued to be implemented until the end of the flight program.

On 30 September 1946, Walter C. Williams and four other engineers from NACA Langley arrived at the Muroc Army Air Field to assist in flight-testing the XS-1. It was supposed to be a temporary assignment, and the group did not even have an official name, although they called themselves the NACA Muroc Flight Test Unit. This marked the beginning of the joint USAF-NACA research airplane program that would culminate with the X-15. The NACA Muroc Flight Test Unit received permanent status from Hugh Dryden on 7 September 1947, with Walt Williams named as head of a group that now numbered 27 people.[8]

By the time the NACA redesignated the unit the High-Speed Flight Research Station (HSFRS) on 14 November 1949, roughly 100 people worked for Williams, who was named chief of the station. In February 1953, the Air Force formally leased 175 acres at the north end of the main taxiway to the NACA for a permanent installation.[9]

On 26 June 1954, the now-200-strong NACA contingent moved from its primitive quarters on South Base to a new headquarters, located in Building 4800 on the north side of the new Edwards flight line. This building still serves as the core of the Dryden Flight Research Center (DFRC). The new facility cost $3.8 million. In contrast to the small NACA station at Edwards, the Air Force contingent at the AFFTC numbered over 8,000. The NACA organization was renamed the High­Speed Flight Station (HSFS) on 1 July 1954, with Williams still in charge.-10

On 27 January 1953, a ground-breaking ceremony was held at the site of the future NACA High­Speed Flight Station on the new main base at Edwards AFB. Shown in the photo are (left to right) Gerald Truszynski, Head of Instrumentation Division; Joseph Vensel, Head of the Operations Branch; Walter Williams, Head of the Station, scooping the first shovel full of dirt; Marion Kent,

Head of Personnel; and California state official Arthur Samet. (NASA)

On 1 October 1958, the National Aeronautics and Space Administration (NASA) replaced the NACA, and on 27 September 1959, NASA Headquarters redesignated the HSFS the Flight Research Center (FRC). By the time the station became a center, Williams was gone. At the behest of Hugh Dryden, on 14 September 1959 he had joined Project Mercury (the first American manned space effort) as its operations director. In his place came Paul F. Bikle, a Pennsylvanian with long experience in flight-testing at the nearby AFFTC. Bikle replaced Williams on 15 September 1959, oversaw its transition to the FRC, and remained for the next 12 years. Bikle believed in doing things quietly and with a minimum of fuss and outside attention. "Under Paul Bikle," one engineer recalled, "we were well aware that headquarters was 3,000 miles away." Like Williams before him, Bikle impressed those who encountered him with his bluntness, drive, and engineering sense.-1111

The first challenge faced by Bikle was shifting from planning for the X-15 program to conducting it. He needed people, and asked Ira Abbott at Headquarters for authority to add 80 new positions to the rapidly growing X-15 team. The personnel and facilities at the FRC expanded throughout the 1960s, with the budget going from $3.28 million in 1959 to $20.85 million in 1963, and to $32.97 million in 1968. The staff went from 292 to a peak of 669 in 1965; by the end of the X – 15 program, the staff was down to 566.-112

The employees of the NACA High-Speed Flight Station are gathered for a 1954 photo shoot on the front steps of building 4800, the new NACA facility at Edwards AFB. This new building was considerably larger than the earlier NACA buildings on South Base to support a staff that had increased from 132 in 1950 to 250 in 1954. As the workload increased and more research flights were completed the complement of employees grew to 662 in 1966, largely the result of the X-15 flight program. (NASA)