Category AERONAUTICS

The Air Force Boeing YC-14 Short Take-Off and Landing (STOL) jet transport technology demonstrator flew for the first time on August 9, 1976, from Boeing Field in Seattle, WA, during the period between Phase I and Phase II of the NASA DFBW F-8 program. Two proto­types were built with the second aircraft flying in October 1976. The YC-14 is noteworthy in that it was the first aircraft to fly with a fault – tolerant multichannel redundant digital fly-by-wire flight control system.

A mechanical backup flight control capability was retained. The full authority triply redundant digital fly-by-wire flight control system, designed by the British Marconi Company, performed computational commands for pitch, roll, and yaw that were used to control the eleva­tor, aileron, and rudder actuation systems. The reconfigurable computer architecture divided the basic control path into three subfunctional ele­ments with these elements replicated to provide fault tolerance. The inter­nal element redundancy management function was intended to detect and isolate faulty elements and perform the necessary reconfiguration. The input signal selection methodology was intended to guarantee that all three computers used the same numbers and thus produced identi­cal output values. During normal operation, the overall system output value was selected as the midvalue of the three individual values. The system would continue to operate in the event of a failure of one com­puter by taking the average of the output of the two remaining comput­ers. If they disagreed, both were disabled and the aircraft reverted to the backup manual control system.[1179]

The YC-14 was also noteworthy in that it used optical data links to exchange data between the triply redundant computers. The optical com­munications medium was chosen to eliminate electromagnetic inter­ference effects, electrical grounding loop problems, and the potential propagation of electrical malfunctions between channels. Optical cou­pling was used to maintain interchannel integrity. Each sensor’s out­put was coupled to the other channels so that each computer had data from each of the other sensors. Identical algorithms in each computer were used. They consolidated the data, enabling equalization and fault detection/isolation of the inputs. The computers were synchronized to avoid sampling time differences and to assure that all computers were receiving identical data inputs.[1180]

An important observation involving redundant computer-controlled fly-by-wire flight control systems was derived from the YC-14 flight-test experience. As noted above, the system was designed to ensure that all computers used the same sensor input values and should therefore produce identical outputs. However, a significant fault in the digital
flight control software was encountered during flight-testing that had not been detected during ground laboratory testing. The software fault resulted in incorrect tracking of control law computations in each of the three flight control channels, with each channel performing signal selections on a different set of values. This resulted in different input data for the three channels. Although the discrepancies between each channel’s inputs were small, the cumulative effect led to large tracking errors between flight control channels when airborne.[1181]

Following cancellation of the Air Force YC-14A program in 1979, the two prototypes were placed in storage at Air Force’s Aerospace Maintenance and Regeneration Group (AMARC) at Davis-Monthan AFB, AZ, in April 1980. The first prototype is now displayed at the Pima Air and Space Museum in Tucson, AZ.

The X-31 was the first international experimental aircraft development program in which the U. S. participated. Two X-31 Enhanced Fighter Maneuverability (EFM) demonstrator aircraft were designed and con­structed by Rockwell International Corporation’s North American Aircraft Division and Deutsche Aerospace. Assigned U. S. Navy bureau Nos. 164584 and 164585, the aircraft would be used to obtain data that could be applied to the design of highly maneuverable next-generation fighters. During the conceptual phase of the program, the personnel examined the application of EFM technologies and defined the require­ments for the demonstrator aircraft. Next, the preliminary design of the demonstrator and the manufacturing approach were defined. Technical experts from the U. S. Navy, German Federal Ministry of Defense, and

NASA evaluated all aspects of the design. Detail design and fabrication followed, with the two aircraft being assembled at the Rockwell International (now Boeing) facility at Palmdale, CA. Both aircraft were required to fly a limited flight-test program at Rockwell. The first aircraft flew its first flight on October 11, 1990, piloted by Rockwell chief test pilot Ken Dyson. The second aircraft made its first flight on January 19, 1991, with Deutsche Aerospace chief test pilot Dietrich Seeck at the controls.[1235]

The X-31 had a digital fly-by-wire flight control system that included four digital flight control computers with no analog or mechanical backup. Three synchronous main computers drove the flight control surfaces. The fourth computer served as a tiebreaker in case the three main computers produced conflicting commands. Three thrust vectoring paddles were mounted on the X-31’s aft fuselage adjacent to the engine nozzle. Directed by the DFBW flight control system, the paddles were moved in and out of the exhaust flow with the resultant thrust vectoring augmenting the aerodynamic control surfaces in pitch and yaw control to improve maneuverability. Made of an advanced carbon-fiber-rein­forced composite material, the paddles could sustain temperatures of up to 1,500 degrees Celsius. The X-31 also had movable forward canards for pitch control. As a result of controllability issues identified during the X-31 flight-test program, fixed strakes between the trailing edge of the wing and the engine exhaust were incorporated. They provided additional nose-down pitch control at very high angles of attack. Another fix that was found neces­sary was the addition of small fixed-nose strakes to help control sideslip.[1236]

During flight-test operations at the Rockwell Aerospace facility, the two X-31s flew 108 test missions, validating the use of thrust vectoring to compensate for loss of aerodynamic control at high angles of attack and expanding the poststall envelope up to 40 degrees angle of attack. The poststall envelope refers to the region in which the aircraft dem­onstrated an ability to maintain controlled flight beyond the normal X-31 stall angle of attack of 30 degrees. X-31 flight operations moved to NASA Dryden in February 1992, with the first flight under International Test Organization (ITO) management occurring in April 1992. The ITO initially included about 110 people from NASA, the U. S. Navy, the U. S. Air Force, Rockwell Aerospace, the Federal Republic of Germany, and

Daimler-Benz. The ITO staff was eventually reduced to approximately 60 people. Overall management of the X-31 program came under by the Defense Advanced Research Projects Agency, with NASA responsi­ble for flight-test operations, aircraft maintenance, and research engi­neering after the project moved to Dryden. The ITO director and NASA’s X-31 project manager at Dryden was Gary Trippensee. Pilots included NASA pilot Rogers Smith, U. S. Navy Cdr. Al Groves, German pilots Karl Lang and Dietrich Seeck, Rockwell International pilot Fred Knox, and Air Force Flight Test Center pilot Lt. Col. Jim Wisneski. By July 1992 the X-31 flight envelope was being expanded in preparation for mili­tary utility evaluations that would fly the aircraft against nonthrust vec­tored fighters to evaluate effectiveness in simulated air combat. Thrust vectoring effectiveness at supersonic speed was evaluated out to Mach

1. 28 at an altitude of 35,000 feet.

In early flight-testing, the X-31 flight control system went into a reversionary mode four times in the first nine flights because of dis­agreement between the two air data sources.[1237] The X-31 was very sen­sitive to sideslip. This caused difficulties for the flight control system at higher angles of attack. Below 30 degrees, the nose boom updated the inertial navigation unit with air data. Above angles of attack of 30 degrees, the inertial navigation unit began calculating erroneous side­slip angles as a result of changes in the relative wind vector. To resolve this problem, a so-called Kiel probe replaced the standard NASA Pitot tube to calculate airflow. The Kiel probe was bent 10 degrees downward from the standard pitot configuration. In addition, the sideslip vane was rotated downward 20 degrees relative to the nose boom to compensate for a yawing oscillation that occurred at an angle of attack of 62 degrees. These changes resulted in accurate air data being provided to the iner­tial navigation unit throughout the X-31 flight envelope with false side­slip readings at high angles of attack eliminated.[1238]

Throughout the process of envelope expansion, many modifications to the flight control laws were required because actual aerodynamics of the aircraft were somewhat different from wind tunnel predictions. When the pilots started flying at angles of attack above 50 degrees, they
encountered erratic lateral lurching movements. In an attempt to coun­ter this phenomenon, narrow, 1/4-inch-wide strips of grit were attached to the sides of the nose boom and the radome. These effectively changed the vortex flow across the forward fuselage of the aircraft, reducing the randomness of the lurches and enabling expansion of the flight envelope to the design angle of attack limit of 70 degrees at 1 g. However, pilots encountered unintentional departures from controlled flight as the air­craft approached poststall angles of attack of 60 degrees during Split-S maneuvers.[1239] The asymmetric yawing moment encountered during this maneuver was beyond the capability of the thrust vectoring system to maintain adequate control.[1240] Testing in the Langley full-scale wind tunnel resulted in nose strakes and a modified slightly blunter nose tip design that were fitted to the two aircraft, allowing resumption of the flight-test program. The nose strakes were 6/10 of an inch wide and 20 inches long and forced more symmetric transition of forebody vortexes. The blunted nose tip reduced yaw asymmetries.[1241]

Poststall pitch control effectiveness, especially with the X-31 center of gravity at the aft allowable design location, was initially marginal.[1242] In these high-angle-of-attack conditions, test pilots rated aircraft response as unsatisfactory. NASA Langley conducted wind tunnel tests of various approaches intended to provide increased nose-down pitch control at high angles of attack. Sixteen different modifications were rapidly tested in the full-scale wind tunnel, with Langley recommending that a pair of strakes 6 inches wide and 65 inches long be mounted along the sides of the aft fuselage to assist in nose-down recovery. These were incorporated on the X-31, with subsequent flight-testing confirming greatly improved nose-down pitch control.[1243] Positive control at 70 degrees angle of attack with a controlled roll around the aircraft velocity vector was demonstrated November 6, 1992. On April 29, 1993, a minimum radius 180-degree post-

stall "Herbst Maneuver” was accomplished for the first time.[1244] During the final phase of evaluation, the X-31s engaged in simulated air combat sce­narios against F/A-18s. During these scenarios, the X-31s were able to outmaneuver the F/A-18s purely through use of poststall maneuvers and without use of thrust vectoring. X-31 test pilots did not support trading off basic fighter characteristics to acquire poststall maneuvering capa­bilities but concluded that improved pitch pointing and velocity vector maneuvering possible with thrust vector control did provide additional options during close-in combat. Thrust vectoring, combined with fully controllable poststall maneuvering, enabled X-31 pilots to position their aircraft in ways that adversary pilots could not counter, but it had to be used selectively and rapidly to be effective.[1245]

In 1994, software was installed in the X-31 to simulate the feasi­bility of stabilizing a tailless aircraft at both subsonic and supersonic speed using thrust vectoring. The aircraft was modified to enable the pilot to destabilize the aircraft with the rudder to lower stability levels to those that would have been encountered if the aircraft had a reduced – size vertical tail. For this purpose, the rudder control surface was used to cancel the stabilizing effects of the vertical tail, and yaw thrust vec­tor commands were applied by the flight control system to restabilize and control the aircraft. The X-31 was flown in the quasi-tailless mode supersonically at 38,000 feet at Mach 1.2, and maneuvers involving roll and yaw doublets, 30-degree bank-to-bank rolls, and windup turns to 2 g were flown. During subsonic testing, simulated precision carrier land­ing approaches and ground attack profiles were successfully evaluated. The quasi-tailless flight-test experiment demonstrated the feasibility of tailless and reduced-tail highly maneuverable fighter/attack aircraft designs. Such designs could have reduced drag and lower weight as well as reduced radar and visual detectability. It determined that thrust vec­toring is a viable flight control effector that can replace the functions provided by a vertical tail and rudder control surface. Potential disad-

vantages include the added weight, complexity, and reliability issues associated with a thrust vectoring system. Additionally, flight condi­tions that require lower engine thrust settings (such as approach and landing) may necessitate provision of additional aerodynamic high – drag devices to enable high-thrust settings to be maintained, ensuring adequate thrust vectoring control. Early integration of such consider­ations into the overall design process, along with an increased level of interaction between propulsion and flight control systems, is required in order to derive the maximum benefit from reduced or tailless aircraft that rely on thrust vectoring for stability and control.[1246]

The No. 1 X-31 aircraft was lost on its 292nd flight on January 19, 1995. German test pilot Karl Lang had just finished a series of test maneu­vers and was in the process of recovering back to a landing at Edwards. At an altitude of 20,000 feet, he observed discrepancies in the air data dis­plays along with a master caution light. The aircraft then began a series of diverging pitch oscillations and became uncontrollable. Lang ejected safely at an altitude of 18,000 feet, and the aircraft crashed in an unpop­ulated desert area just north of Edwards. The crash was determined to have resulted from an unanticipated single-point failure in the nose – mounted Kiel probe that provided critical airspeed and altitude data to the aircraft flight control system computers. These data were critical to safe flight, yet the Kiel probe did not include provision for electrical de­icing, presumably because the aircraft would only be flown in clear desert weather conditions. However, during descent to recovery back to Edwards, ice accumulated in the unheated X-31 pitot tube, resulting in the flight control system automatically configuring the aircraft control surfaces for what it assumed were lower airspeed conditions. Unanticipated move­ments of the flight control surfaces caused the aircraft to begin oscillating about all axes followed by an uncontrolled pitch-up to an angle of attack of over 90 degrees.[1247] The subsequent X-31 accident investigation board recommended that training be conducted on the system safety analy­sis process, that procedures be implemented to ensure that all test team members receive configuration change notices, and that improvements be made in the remaining X-31 to prevent similar single-point failures.[1248]

A panel that included former Dryden Research Center director Ken Szalai met at Dryden in early 2004 to review the X-31 accident. The panel noted that the primary contributing factor was the installation of the unheated Kiel probe in place of the original heated Pitot tube. The lack of electrical de-icing capability on the Kiel probe had not been considered a safety risk because X-31 mission rules prohibited flight in precipitation or clouds. However, there was no stipulation specifically restricting flight during potential icing conditions, despite simulations that showed icing of the Pitot static system could lead to loss of con­trol.[1249] Information had been distributed among the X-31’s test pilots and flight-test engineers explaining the Pitot tube change, but a formal process was not in place to ensure that everyone fully understood the implications of the change. Test pilot Lang had noticed anomalies on his cockpit instrumentation and, assuming the presence of icing, told the control room that he was switching on Pitot heat. Shortly afterward, he advised that he was leaving the Pitot heat on for descent and approach to landing. The ground controller then told Lang that the pitot heat switch in the cockpit was not functional. Discrepancies between the X-31’s air­speed and altitude readouts were being observed in the control room, but that information was not shared with the entire control room staff. There was a redundant source of air data and a pilot-selectable alterna­tive control mode that could have saved the aircraft if better commu­nications had existed. Dryden X-31 project manager Gary Trippensee noted that complacency is the enemy of success in flight research; prior to the accident, 523 successful X-31 research missions had been flown.[1250]

In 2000, the remaining X-31 was brought back from long-term storage at NASA Dryden, where it had been since 1995, and reconfig­ured for another round of flight-testing for the Vectoring, Extremely Short Takeoff and Landing Control and Tailless Operation Research (VECTOR) program. This program would explore the use of thrust vec­toring for extremely short takeoff and landing (ESTOL), with a focus on the aircraft carrier environment. An international Cooperative Test Organization was created for the VECTOR program. U. S. partic- ipants/partners were the Navy, Boeing, General Electric, and NASA.[1251]

The Swedish government was represented by Volvo and SAAB, with the German Ministry of Defense and DASA (Daimler-Benz consor­tium) from Germany. The X-31 aircraft was modified to incorporate a Swedish RM-1 engine, the same powerplant used in the Saab JAS-39 Gripen fighter.[1252] On February 24, 2001, flown by U. S. Navy Cdr. Vivian Ragusa, the upgraded X-31 took to the air for the first time from Naval Air Station (NAS) Patuxent River.[1253] German test pilot Rudiger "Rudy” Knopfel, U. S. Marine Corps Maj. Cody Allee, and Navy Lt. J. R. Hansen would fly most of the subsequent ESTOL test program.[1254] The VECTOR X-31 went on to accomplish over 2 years’ of flight-testing, culminating in the final ESTOL flight by Maj. Allee on April 29, 2003.

From April 22 to 29, 2003, the VECTOR X-31 flew 11 test flights, during which fully automated, high-angle-of-attack approaches to land­ing were conducted. The automated flight control system utilized inputs from a special Global Positioning System (GPS)-based navigation sys­tem to maneuver the aircraft to a precise spot above the runway. Known as the Integrity Beacon Landing System (IBLS), it was supplemented by two virtual satellites, or "pseudolites,” on both sides of the runway. Precise spatial position and flight attitude data were inputs for the auto­matic approach control and landing system used in the VECTOR X-31. An ESTOL approach began with the pilot flying into the area covered by the pseudolites; after entering an engagement box, the automatic approach and landing system was activated. The aircraft then assumed a high-angle-of-attack approach attitude and followed a curvilinear path to the touchdown point. Just before touchdown, with the thrust vectoring paddles less than 2 feet above the runway, the X-31A auto­matically reduced its attitude back down to the normal 12-degree angle of attack for landing. An autothrottle system from an F/A-18 and a special autopilot developed by the VECTOR team were coupled with the flight control system to provide the integrated flight and propul­sion control capability used to automatically derotate the aircraft from its steep final approach attitude to touchdown attitude at 2 feet above the runway.

On the final flight of the VECTOR program, the angle of attack during landing approach was 24 degrees (twice the angle of attack on a normal landing approach). Approach airspeed was 121 knots, or about 30 per­cent lower than the normal 175 knots, and the resultant landing distance was only 1,700 feet, compared to the normal landing distance of nearly 8,000 feet. Maj. Allee commented on the experience of riding along on a VECTOR X-31 automatic approach and landing: "There are no g forces and you sit leaning somewhat backwards in the ejection seat while the nose is pointing sharply upwards. . . . At angle of attacks greater than 15 degrees the pilot cannot see the runway except on the screen on the right – hand side of the instrument panel. . . . Whereas on a normal landing the landscape flashes by, now everything takes place as if in slow motion.”[1255]

Another technical accomplishment demonstrated during the VECTOR X-31 program was the successful test of an advanced Flush Air Data System (FADS). Based on data collected by a dozen sensors located around the nose of the aircraft, the FADS provided accurate air data, including airspeed, altitude, angle of attack, and yaw angle, to the flight control system at angles of attack up to 70 degrees all the way out to supersonic speed.[1256]

The two X-31 aircraft completed a total of 580 flights, a record for an X-plane program. Of these, 559 were research missions and 21 were flown in Europe in support of the 1995 Paris Air Show. Fourteen pilots from NASA, the U. S. Navy, the U. S. Marine Corps, the U. S. Air Force, the German Air Force, Rockwell International, and Deutsche Aerospace flew the aircraft during the original program at Palmdale and Dryden, with two U. S. pilots (one Navy and one Marine Corps) and a German pilot flying the VECTOR X-31 test program at Patuxent River. The surviv­ing X-31, U. S. Navy Bureau No. 164585, flew 288 times, making its last flight on April 29, 2003. This aircraft is now on display at the Deutsches Museum annex at Oberschleifiheim, near Munich, and it will eventually be returned to the United States. The other X-31, bureau No. 164584, had flown 292 times before it was lost on January 19, 1995.

The Engine Component Improvement project was tasked with enhanc­ing performance and lowering fuel consumption of several existing com­mercial aircraft jet engines, in particular Pratt & Whitney’s JT8D and JT9D engines and GE’s CF6. The specific goals included:

• Improving the current versions of the engines without requiring a brand-new design or engine replacement.

• Reducing the amount of fuel a typical jet engine would use on any given flight by 5 to 6 percent.

• Significantly slowing the pace at which the engine’s components would naturally degrade and cause a loss of performance over time.

To do this, researchers tried and tested several ideas, including reduc­ing the clearance between rotating parts, lowering the amount of cool­ing air that is passed through the engine, and making refinements to the aerodynamic design of certain engine parts to raise their efficiency. All together, engineers identified 16 concepts to incorporate into the engines.[1309]

Ultimately, as a result of the Engine Component Improvement efforts, engine parts were incorporated that could resist erosion and warping, better seals were introduced, an improved compressor design was used, and ceramic coatings were added to the gas turbine blades to increase their performance. Tests of the improvements were so promising that many were put into production before the program ended, benefiting the

workhorse airliners at the time, namely the McDonnell-Douglas DC-9 and DC-10, as well as the Boeing 727, 737, and 747.[1310]

Observing the success of fuel cells in space flight, NASA researchers in the late 1990s began to explore the potential for fuel cells to power air­craft. The cells were attractive to NASA’s aeronautics directorate because they have near-zero emissions, are quiet and efficient, and can work in any environment where the temperature is lower than the cell’s oper­ating temperature. Valerie Lyons, NASA’s Division Chief of Power and In-Space Propulsion, said she began pushing for NASA aeronautics to pursue fuel cell research about 10 years ago. "I would venture to say they hadn’t really looked at it before that,” she said. "When I looked at the fuel cell area, I said, ‘This is pretty interesting, can we use this?’”[1478]

One of NASA’s main fuel cell initiatives was the Next Generation Clean Aircraft Power (NEXCAP) program, which sought to use the cells as auxiliary power units (APUs) for aircraft.[1479] The APUs could be used in onboard electrical systems to power a grounded aircraft, providing an alternative to wasting fuel and producing emissions by drawing on power from an idling engine. NASA researchers hoped that the APU research would eventually lead to the design and test flight in 2030 of an electric airplane, which would rely on fuel cells for propulsion.[1480] While NASA’s electric airplane never came to fruition, Boeing flew the first piloted airplane powered by hydrogen fuel cells in 2008.[1481] NASA researchers maintain a strong interest in the potentiality of fuel cells to meet the future energy needs of the Nation.

NASA and AeroVironment upgraded the Pathfinder UAV to an improved configuration known as the Pathfinder Plus, which had a longer wing­span, two additional electric motors for a total of eight motors, and improved solar cells. The two additional motors were more efficient than the six carried forward from Pathfinder. The Pathfinder Plus UAV, which was intended to serve as a "transitional” aircraft between Pathfinder and the next-generation Centurion UAV, had a wingspan of 121 feet that

included use of four of the five wing sections from the original Pathfinder. The new center wing section contained more efficient solar cells that converted approximately 19 percent of the solar energy into electrical power for the vehicle’s motors, avionics, and communications system. This compared with an efficiency rating of only 14 percent for the solar cells on the remaining four wing sections from the original Pathfinder. The addition of the fifth wing section enabled Pathfinder Plus to gen­erate 12,500 watts, as compared with the original Pathfinder’s 7,500 watts. Pathfinder Plus had a gross weight of 700 pounds, up to a 150- pound payload capacity, an aspect ratio of 15 to 1, a wing chord of 8 feet, and a power-off glide ratio of 21 to 1. Pathfinder Plus enabled higher – altitude flights and was used to qualify the next-generation Centurion wing panel structural design, airfoil, and SunPower Corporation’s solar array. Several flight tests were conducted in Hawaii. On its last flight, on August 6, 1998, Pathfinder Plus set a world altitude record of 80,201 feet for solar-power – and propeller-driven aircraft. These flight tests demon­strated the power, aerodynamic, and systems technologies needed for the Centurion and confirmed the practical utility of using high-altitude remotely controlled solar powered aircraft for commercial purposes.[1542]

During the early 1970s, the Ames Flight Simulator for Advanced Aircraft (FSAA) became operational and the first tilt rotor simulations were suc­cessfully accomplished. By 1975, the Army decided to augment the rotary wing flight dynamics research at Ames as NASA initiated the fabrica­tion of the Vertical Motion Simulator (VMS). This simulator, with very large vertical and horizontal motion capability, was a national asset well suited for rotary wing research.

At Langley, a major instrument flight rules (IFR) investigation was conducted under the VTOL Approach and Landing Technology (VALT) program. The VALT Boeing-Vertol CH-47 Chinook helicopter was the pri­mary research vehicle for exploring the control/display/task relationships. In addition, the Sikorsky SH-3 Sea King helicopter was used as a testbed for exploring the merits and defining the electro-optical parameter require­ments associated with advanced "real-world” display concepts. The objec­tive was to identify systems that might be capable of providing a pilot an "out-the-window display” during IFR flight conditions through the use of fog-cutting sensors or advanced computer-generated visual situation dis­plays. The VALT CH-47 flights were conducted at the Wallops Flight Center, where the NASA Aeronautical Research Radar Complex provided omni­directional tracking coverage. This facility permitted the investigation of a wide variety of approach trajectories and selection of any desired wind direction relative to the final approach heading. Computer-graphic dis­plays were generated on the ground and transmitted via video link to the aircraft for presentation in the pilots’ instrument panel. The integrated flight-test system permitted manual, augmented, or completely automatic control for executing flight trajectories that could be optimized from the standpoints of fuel, time, airspace utilization, ride qualities, noise abate­ment, or air traffic control considerations. Many concepts were explored in the IFR program, including flight director control/display concepts and signal smoothing techniques, which proved valuable in achieving fully automatic approach and landing capability.[293] Extensive flight demon­strations were conducted at Wallops Flight Center with the VALT CH-47 aircraft for Government and industry groups to demonstrate the new progress achieved in IFR approach and landing technology.

In structures technology, one of the important outcomes of the space program was the development and implementation of comprehensive computational finite element analyses. State-of-the-art finite element methodology was collected from among the large aerospace compa­nies and unified into the NASA Structural Analysis (NASTRAN) com­puter program. The basic development contract was managed by NASA’s Goddard Space Flight Center and then by Langley for improvements and distribution to approximately 260 installations. During the early 1980s, Langley played a key role in bringing advanced structural design capability into the helicopter industry. The contribution here was the onsite assignment of an experienced structural dynamics specialist at a prime manufacturer’s facility to guide the integration of the prelimi­nary static structural design methodology with rotor dynamic analysis methodology.[294] This avoided the tedious process of repeatedly freez­ing an airframe structural design effort and each time doing a separate dynamic analysis to determine if an acceptable dynamic response cri­terion was achieved.

During this period, the Army added to its already extensive helicop­ter crash-test activities by joining with NASA to crash-test the Boeing Vertol CH-47C helicopter in the Impact Dynamics Research Facility at Langley, which accommodated aircraft up to 30,000 pounds.[295] The facil­ity had been converted from a Lunar Landing Research Facility to a center for the study of crash effects on aircraft. A unique feature of this massive gantry structure was the capability to impact full-scale aircraft under free-flight conditions with precise control of attitude and velocity.

The ongoing rotary wing research began to expand in scope with the establishment of the Army co-located research groups at the three NASA Centers. At Ames, full-scale rotor wind tunnel testing continued at an increased pace in the 40- by 80-foot tunnel. In the 1970s, the wind tunnel tests included the Sikorsky Advancing Blade Concept (ABC) rotor. This rotor concept incorporated two counter-rotating coaxial rotors. The hingeless blades were very stiff to allow the advancing blades on both sides of the rotor disk to balance the opposing rolling moments thereby

maintaining aircraft trim as airspeed is increased. Forward thrust is supplied by auxiliary propulsion rather than by forward tilt of the main rotor as in conventional helicopter designs.

NASA also tested a full-scale semispan wing-pylon-rotor of the Bell Helicopter tilt rotor design.[296] This test was followed by a similar entry of a semispan setup of a Boeing Vertol tilt rotor concept. During this period, improvements were made in the 40- by 80-foot Full-Scale Tunnel to upgrade the research capability. Its online data capability was aug­mented by introducing a new Dynamic Analysis System for real-time analysis of critical test parameters. A new Rotor Test Apparatus (RTA) was added to facilitate full-scale rotor testing. With this new equip­ment in place, a Kaman Controllable Twist Rotor (CTR) was first inves­tigated in 1975.

In the early 1970s, the modest in-house research funding level for rotary wing projects led to seeking other sources within the new, more elaborate financial system of NASA. It turned out that contracting out – of-house research had become a staple of the rapidly growing procure-

ment system.[297] This offered the opportunity to begin to solicit, select, and fund small supporting research contracts to augment the in-house rotary wing work categorized as Research and Technology Base. In the Flight Research Branch at Langley between 1969 and 1974, over 77 con­tractor reports (CR) and related technical papers were published. The performing organizations included industry and university research departments. The research topics included analytical and experimental investigations of rotor-blade aeroelastic stability, blade-tip vortex aero­dynamics, rotor-blade structural loads prediction, free-wake geometry prediction, nonuniform swash-plate dynamic analysis program, rotor – blade dynamic stall, composite blade structures, and variable geome­try rotor concepts, In the mid 1970s, this entry into contracted research to augment in-house work was further augmented by teaming of NASA and Army rotary wing research at the three NASA Centers. Finally, proj­ects between NASA, the Army, and contractors evolved into major joint efforts in Systems Technology and Experimental Aircraft during the following decade.

The mid-1970s brought two major rotary wing experimental aircraft programs, both jointly funded and managed by NASA and the Army. At Langley, the Rotor Systems Research Aircraft (RSRA) program was launched. This was a new approach to conducting flight research on helicopter rotor systems.[298] Two vehicles were designed and fabricated by Sikorsky Aircraft. The basic airframe, propulsion, and control systems of the two RSRA vehicles were those of the Sikorsky S-61 Sea King heli­copter. In addition, the RSRA incorporated a unique rotor force balance system and isolation system, a programmable electronic control system, a variable incidence wing with a force balance system, drag brakes, and two TF34 auxiliary thrust turbofan engines. As a unique safety feature, the three-member-crew ejection system incorporated automatic bal­anced sequencing of explosive separation of the test rotor-blades as the first step in permitting the rapid ejection of the pilot, copilot, and test engineer. After design and fabrication at Sikorsky, the first of two RSRA vehicles made its first flight in 1976. After initial tests of the helicopter configuration, flight-testing was continued at the NASA Wallops Flight

Center with the Langley-Army project team and contractor onsite sup­port. Acceptance testing was completed by the Langley team, which was then joined by Ames flight-test representatives in anticipation of pend­ing transfer of the RSRA program to Ames.

At Ames, a NASA-Army program of equal magnitude was launched to design and fabricate two XV-15 Tilt Rotor Research Aircraft (TRRA). In this case, the program focused on a proof-of-concept flight investiga­tion. This concept, pursued by rotary wing designers since the early 20th century, employs a low-disk-loading rotor at each wingtip that can tilt its axis from vertical, providing lift, to horizontal, providing propulsive thrust in wing-borne forward flight. The TRRA contract was awarded to Bell Helicopter Textron. Late in the program, as the XV-15 reached flight status, the United States Navy added funding for special mission – suitability testing. Eventually, XV-15 testing gave confidence to tilt rotor advocates who successfully pushed for development of an operational system, which emerged as the V-22 Osprey.

The RSRA and TRRA experimental aircraft programs together rep­resented a total initial investment of approximately $90 million, ($337 million in 2009 dollars), shared equally by NASA and the Army. The size and scope of these programs were orders of magnitude beyond previ­ous NACA-NASA rotary wing projects. This represented a new level of

resources in rotary wing research for NASA and with it came consider­ably more day-to-day visibility within the NASA aeronautics program.

The bicentennial year of 1976 also marked a year of major orga­nizational change in NASA rotary wing research. As part of an over­all Agency reassessment of the roles and missions of each Center, the Ames Research Center was assigned the lead Center responsibility for helicopter research. An objective of the lead Center concept was to con­solidate program lead in one Center and, wherever possible, combine research efforts of similar nature. As a result, all rotary wing flight test, guidance, navigation, and terminal area research were consolidated at Ames, which brought these research activities together with the exten­sive simulation and related flight research facilities. Langley retained supporting research roles in structures, noise, dynamics, and aero – elasticity. The realignment of responsibilities and transfer of flight research aircraft caused unavoidable turbulence in the day-to-day conduct of the rotary wing program from 1976 to 1978. However, the momentum of the program gradually returned, and the program grew to new levels with NASA and Army research teams at Ames, Langley, and Glenn working to carry out their responsibilities in rotary wing research.

At Ames, the testing of full-scale rotor systems continued at an increasing pace in the 40 by 80 Full-Scale Tunnel. In 1976, the Controllable Twist Rotor concept was tested again, this time with mul­ticyclic control. "Two-per-rev” (two control cycles per one rotor revolu­tion), "three-per-rev,” and "four-per-rev” cyclic control was added to the CTR’s servo flap system to evaluate the effectiveness in reducing blade stresses and vibration of the fuselage module. Both favorable effects were achieved with only minor effect on the rotor power requirements. The Sikorsky S-76 rotor system was tested in 1977 in a joint NASA – Sikorsky investigation of tip shapes. This was followed by a joint NASA – Bell investigation of the Bell Model 222 fuselage drag characteristics. In 1978, the NASA-Army XV-15 Tilt Rotor Research Aircraft arrived from Bell Helicopter for full-scale wind tunnel tests prior to initiation of its own flight tests. The wind tunnel tests revealed a potential tail struc­tural vibration problem that would be further explored in flight follow­ing the strengthening of the empennage attachment structure. The next rotor test was the Kaman Circulation Control Rotor (CCR) in 1978.[299]

A new concept was introduced based on technology developed at the David Taylor Ship Research and Development Center (since 1992 the Carderock Division of the Naval Surface Weapons Center). The Kaman rotor utilized elliptical-shaped airfoils with trailing edge slots. Lift was augmented by blowing compressed air from these slots. The need for mechanical cyclic blade feathering to provide rotor control was elimi­nated replaced by cyclic blowing. The wind tunnel testing investigated the amount of blowing control necessary to maintain forward flight. In 1979, the Lockheed X-Wing Stoppable Rotor was tested in the 40 by 80 Full-Scale Tunnel. This concept, funded by the Defense Advanced Research Projects Agency, also incorporated a circulation control con­cept. The X-Wing rotor was designed to be stoppable (and startable) at high forward flight speed while still carrying lift. Since two of the four blade trailing edges become leading edges when stopped, provisions were made to provide for separate blowing systems for the leading and trail­ing edges of the blades. When operating as a fixed X-Wing aircraft, air­craft roll and pitch control were provided by differential blowing from the aft edges of opposing, nonrotating blades serving as swept forward and aft wings. The wind tunnel tests of the 25-foot-diameter rotor suc­cessfully demonstrated the ability to start and stop the rotor at speeds of approximately 180 knots (maximum tunnel speed).

The Boeing Vertol Bearingless Main Rotor (BMR) was tested in 1980.[300] The BMR used elastic materials in the construction of the rotor hub rather than mechanical bearings for articulation. Such designs have aeroelastic stability characteristics different from conventional mechan­ical systems. Therefore, the wind tunnel tests investigated the degree of stability present and established appropriate boundaries for safe flight. In addition, in 1980, the Sikorsky Advancing Blade Concept (ABC) coax­ial rotor was again tested in the 40 by 80 Full-Scale Tunnel.[301] In this entry, the full-scale rotor was tested atop a configuration replica of the actual XH-59A flight-test aircraft. This testing focused on an investiga­tion of the drag characteristics of the rotor shaft and hubs of the coaxial rotors. In an effort to reduce the drag, tests were made with the actual fuselage modeled and the actual flight demonstrator hardware compo-

nents utilized to explore several inter-rotor fairing configurations. (In 2008, Sikorsky Aircraft unveiled a new technology demonstrator aircraft incorporating the advancing blade concept identified as the X2. In this design forward thrust is provided by a pusher propeller installation.)

In 1984, Ames shut down the 40- by 80-foot facility for tunnel mod­ification to upgrade the 40- by 80-foot section to a speed capability of 250 knots and add a new 80 by 120 leg to the tunnel facility capable of speeds to 80 knots. The upgraded facility, known as the National Full – Scale Aerodynamics Complex (NFAC), reopened in 1987 and would have been operated by NASA until 2010. However, budgetary pressures forced its closure in 2003. Four years later, in 2007, the United States Air Force’s Arnold Engineering Development Center (AEDC) upgraded key operating systems and reopened the facility under a 25-year lease with NASA. The anticipated majority customer for this national asset was seen to be the United States Army, in collaboration with NASA, in support of rotary wing research.

A Helicopter Transmission Technology program was initiated at the Glenn Research Center to foster the application of an extensive tech­nology base in bearings, seals, gears, and new concepts specifically to helicopter propulsion systems.[302] Research continued at a growing pace. In order to upgrade the analytical methods for large spiral bevel gears, NASA supported the development and validation testing of finite ele­ment method computer programs by Boeing Vertol. The opportunity was taken to utilize the available aft transmission hardware assets, available from the canceled XCH-62 Heavy Lift Helicopter Program, for analyt­ical methods validation data. Another program at Glenn was the joint NASA-DARPA Convertible Engine Systems Technology (CEST) pro­gram. This program involved the modification of a TF34 turbofan engine to a fan/shaft engine configuration for use as a research test engine to investigate the performance, control, noise, and transient characteris­tics. The potential application of CEST was to the X-Wing vehicle con­cept by using a single-core engine to provide shaft power to a rotor in hover and low speed, and conversion capability to provide fan thrust for high speed, stopped rotor mode, and flight propulsion.

Ongoing research in helicopter handling qualities continued and expanded at the Ames Research Center. In 1978, one of these programs

provided essential simulation data on the effects of large variations in rotor design parameters on handling qualities and agility in helicopter nap-of – the-Earth (NOE) flight. The parameters investigated including flapping hinge offset, flapping hinge restraint, rotor blade inertia, and blade pitch – flap coupling. Experiments were carried out on the Ames piloted simula­tors to systematically study stability and control augmentation systems designed to improve NOE flying and handling qualities characteristics.

New efforts in computational analysis to increase rotor efficiency began at Ames. An analytical procedure was developed to predict rotor performance trends in relation to changes in the shape of the blade tips. The analytical procedure utilized two full potential flow-field computer programs developed for computation of the transonic flow field about fixed wings and airfoils. The analytical procedure rapidly became a use­ful tool for predicting aerodynamic performance improvements that may be achieved by modifying blade geometry. The procedure was guided by design studies and reduced the experimental testing required to select blade configurations. NASA continued the long-established tradition of fur­nishing excellent references for technical practice when, in 1980, research scientist Wayne Johnson, a member of the Army-NASA research team at Ames, published his book Helicopter Theory, a comprehensive state-of – the-art coverage of the fundamentals of helicopter theory and engineering analysis. The extensive bibliography of cited literature included an exten­sive listing of rotary wing technical publications authored by researchers from the NACA, NASA, the Army, industry, and academia.[303]

Research accelerated on advancing the ability of a helicopter to execute a radar approach. Civil weather/mapping radar could be used to provide approach guidance under instrument meteorological con­ditions (IMC) to select safe landing environments. Onboard radar sys­tems were widely used by helicopter operators to provide approach guidance to offshore oil rigs without the need for electronic naviga­tion aids at the landing site. For use over the water, the radar provided guidance and ensures obstacle awareness and avoidance, but involved very high pilot workload and limited guidance accuracy. For use over land, the ground clutter return made these approaches infeasible with­out more advanced radar systems. Two programs at Ames resulted from major advances in radar approaches. One program involved the

use of a video data processor in conjunction with the weather radar for overwater approaches. This system automatically tracked a des­ignated radar target and displayed a pilot-selected approach course. The second radar program involved the development of an innovative technique to suppress ground clutter radar returns in order to locate simple, low-cost radar reflectors near the landing site. This program was extended to provide the pilot with precision localizer and glide – slope information using airborne weather radar and a ground-based beacon or reflector array.

The 1980s brought several major accomplishments in the tilt rotor program.[304] The second XV-15 aircraft was brought to flight status and accepted by the Government after check flights and acceptance cere­monies at NASA’s Dryden Flight Research Center on October 28, 1980. It was then used for flight tests aimed at verifying aeroelastic stabil­ity, evaluating fatigue load reduction modifications, and expanding the maneuver envelope. Subsequently, this aircraft was ferried to Ames, where tests continued in the areas of handling qualities, flight con­trol, and expansion of the landing approach envelope. The first XV-15 aircraft was brought to flight status in late 1980, and initial work was

done on a ground tiedown rig to measure the downwash field and noise environment. Meanwhile, the second XV-15 participated in the Paris Air Show. The aircraft performed daily, on schedule, and received wide acclaim as a demonstration of new aeronautical technological achieve­ment. The XV-15 crew concluded each daily performance with a cour­teous "bow,” the hovering tilt rotor ceremoniously dipping its nose to the audience. After the flight demonstration in France and subsequent flights in Farnborough, England, the aircraft was returned to Ames for continued flight demonstration and proof-of-concept testing. The two vehicles achieved a high level of operational reliability, not the usual attribute of highly specialized research aircraft. One of the vehicles was returned to Bell Helicopter under a cooperative arrangement that made the aircraft available to the contractor at no cost in exchange for doing a number of program flight-test tasks, particularly in the mission suit­ability category. The overall success of the NASA-Army XV-15 (with a rotor diameter of 25 feet and a gross weight of 13,428 pounds) proof – of-concept program contribution is reflected in the application of the proven technology to the design and production of the new joint-ser­vice V-22 Osprey, (rotor diameter: 38 feet; gross weight: 52,000 pounds). The classic claim of research results having to endure a 20-year shelf life before actual engineering design application begins did not apply. It took only 5 years to move from achieving proof-of-concept with the XV-15 research aircraft to initiation of preliminary design of the oper­ational V-22 Osprey.

There has been over a half century of an unbroken series of NACA – NASA research contributions to tilt rotors since early XV-3 flight assess­ments and wind tunnel testing in the mid-1950s.[305] Since that beginning, NACA-NASA researchers have pursued many subject areas, includ­ing tilt rotor analytical investigations to solve a rotor/pylon aeroelas – tic stability problem, dynamic model aeroelastic testing in the Langley Transonic Dynamics Tunnel, analytical method development and verifi­cation, wind tunnel tests of full-scale rotor/wing/pylon assembles, XV-15 vehicle wind tunnel tests and flight tests, and detailed investigation of many other potential problem areas. This sustained effort and the robust demonstration and advocacy of the technology’s potential resulted in the XV-15 program being cited in 1993 as "the program that wouldn’t

die” in a University of California at Berkeley School of Engineering case study in a course on "The Political Process in Systems Architecture.”[306]

During the early 1980s, the rotary wing activity at Glenn Research Center increased with the addition of new transmission test facilities rated at 500 and 3,000 horsepower. Research progressed on traction drive, hybrid drive, and other advanced technology concepts. The prob­lem of efficient engine operation at partial power settings was addressed with initial studies indicating turbine bypass engine concepts offered potential solutions. Similar studies on contingency power for one – engine-inoperative (OEI) emergency operation focused on water injection and cooling flow modulation. Renewed efforts in aircraft icing included rotary wing icing research. A broad scope program was launched to study the icing environment, develop basic ice accretion prediction methods, acquiring in-flight icing data for comparison with wind tun­nel data from airfoil icing tests to verify rotor performance prediction methods. In addition, flight tests of a pneumatic deicing boot system were conducted using the Ottawa spray rig and the United States Army CH-47 in-flight icing spray system. In 1983, research testing began on the NASA-DARPA Convertible Engine System Technology program.[307] TF34 fan/shaft engine hardware with variable fan inlet guide vanes for thrust modulation was used to evaluate fan hub design and map the steady-state and transient performance and stability of the concept. New rotary wing efforts were also started in the areas of transmission noise, and flight/propulsion control integration technology.

Langley Research Center activity in rotary wing research increased substantially within the Structures Directorate, with focused programs in acoustics, dynamics, structural materials, and crashworthiness. This research was carried out in close association with the Army Structures Laboratory, now known as the Vehicle Technology Directorate (VTD). NASA and Army joint use of the Langley 4- by 7-meter tunnel for aero­dynamic and acoustic model testing became an important feature of the rotary wing program. Confirmed progress was achieved in airframe dynamic analysis methodology addressing the engineering manage­ment and execution of the efficient use of finite element methods for

simultaneous tasks of static and dynamic airframe preliminary design.[308] These techniques were demonstrated, publicly documented, and verified by comparison with shake test data for the CH-47 helicopter airframe. Other research related to helicopter dynamics included participation with the Army in a program to demonstrate the use of closed-loop multicyclic control of rotor-blade pitch motion for vibration reduction. The program involved flight-testing of an Army OH-6 helicopter by Hughes Helicopters.[309]

One of the more innovative approaches to research teaming was developed in the area of rotary wing noise. In 1982, discussions between the American Helicopter Society and NASA addressed the industry con­cern that the proposed rulemaking by Federal Aviation Administration would place the helicopter industry at a considerable disadvantage. The issue was based on the point that the state-of-the-art noise prediction did not allow the prediction of noise for new designs with acceptable confidence levels. As a result, NASA and the Society, joined by the FAA and the Helicopter Association International (HAI)—an organization of helicopter commercial operators—embarked on a joint program in noise research. Through the AHS, American helicopter manufacturers pooled their research with that of NASA under a 5-year plan leading to improved noise prediction capability. All research results were shared among the Government and industry participants in periodic techni­cal exchanges. Langley managed the program with full participation by Ames and Glenn Research Centers in their areas of expertise.

After delivery of the two RSRA vehicles to the Ames Research Center in the late 1970s, the helicopter and compound (with wing and TF34 turbofan engines installed) configurations were involved in an extended period of ground – and flight-testing to document the characteristics of the basic vehicles. This included extensive calibrations of the onboard load measurement systems for the rotor forces and moments; wing lift, drag, and pitching moment; and TF34 engine thrust. This work was fol­lowed by the initiation the research flight program utilizing the deliv­ered S-61 rotor system. In 1983, NASA and DARPA launched a major research program to design, fabricate and flight-test an X-Wing rotor on the new RSRA. The RSRA was ideally suited to the testing of new rotor

concepts, being specifically design for the purpose. One RSRA vehicle was returned to Sikorsky Aircraft for installation of an X-Wing rotor. This aircraft was eventually moved to NASA Dryden Flight Research Center at Edwards Air Force Base, CA, where final preparations were made for flight-testing. The second vehicle embarked on fixed-wing flight testing at the Dryden Center to expand and document the flight enve­lope of the RSRA beyond 200 knots, the speed range of interest in the start-stop conversion testing for the X-Wing rotor.

Contributions were beginning to emerge from the NASA-American Helicopter Society Rotorcraft Noise Prediction Program, the joint Government-industry effort initiated in 1983.[310] The four major thrusts were: noise prediction, database development, noise reduction, and crite­ria development. Fundamental experimental and analytical studies were started in-house and under grants to universities. In order to obtain high – quality noise data for comparison with evolving prediction capability, a wind tunnel testing program was initiated. This NASA-sponsored pro­gram was performed in 1986 in the Dutch-German wind tunnel (Duits – Nederlandse wind tunnel, DNW) using a model-scale Bo 105 main rotor. This program was performed with the support of the Federal Aviation Administration and the collaboration of the German aerospace research establishment. In these tests and in subsequent tests of the model in the DNW tunnel in 1988, researchers gained valuable insight into the aero – acoustic mechanism of blade vortex interaction (BVI) noise.

In regard to rotor external noise reduction, Langley researchers investigated the possibility of BVI noise reduction using active control of blade pitch. A model-scale wind tunnel test was conducted in the Langley Transonic Dynamics Tunnel (TDT) using the Aeroelastic Rotor Experimental System (ARES).[311] Results were encouraging and demon­strated noise level reductions up to 5 decibels (dB) for low and moderate forward speeds. A major contribution of the NASA-AHS program was the development of a comprehensive rotorcraft system noise prediction capability. The primary objective of this capability, the computer code named ROTONET, was to provide industry with a reliable predictor for

use in design evaluation and noise certification efforts. ROTONET was developed in several phases, with each phase released to Noise Reduction Program participants for testing and evaluation. Validation data from flight test of production and experimental rotorcraft constituted a vital element of the program. The first was of the McDonnell-Douglas 500E helicopter, tested at NASA’s Wallops Flight Facility. The second flight – test effort at Wallops, a joint NASA-Army program, was performed in

1987 using an Aerospatiale Dauphine helicopter, which had a relatively advanced blade design and a Fenestron-type (ducted) tail rotor. The year

1988 saw a joint NASA-Bell Helicopter effort in flight investigation of the noise characteristics the NASA-Army XV-15 Tilt Rotor Research Aircraft. The results indicated that while the aircraft seemed very quiet in the airplane mode, significant blade-vortex interaction noise was evi­dent in the helicopter mode of flight. NASA benefited from the inter­action with and participation in the variety of industry noise programs, which helped set the groundwork for subsequent joint participation in rotary wing research.[312]

The ballistic missile and atomic bomb became realities within a year of each other. At a stroke, the expectation arose that one might increase the range of the former to intercontinental distance and, by installing an atomic tip, generate a weapon—and a threat—of almost incomprehen­sible destructive power. But such visions ran afoul of perplexing techni­cal issues involving rocket propulsion, guidance, and reentry. Engineers knew they could do something about propulsion, but guidance posed a formidable challenge. MIT’s Charles Stark Draper was seeking inertial guidance, but he couldn’t approach the Air Force requirement, which set an allowed miss distance of only 1,500 feet at a range of 5,000 miles for a ballistic missile warhead.[554]

Reentry posed an even more daunting prospect. A reentering 5,000-mile-range missile would reach 9,000 kelvins, hotter than the solar surface, while its kinetic energy would vaporize five times its weight in iron.[555] Rand Corporation studies encouraged Air Force and industry mis­sile studies. Convair engineers, working under Karel J. "Charlie” Bossart, began development of the Atlas ICBM in 1951. Even with this seemingly rapid implementation of the ballistic missile idea, time scales remained long term. As late as October 1953, the Air Force declared that it would not complete research and development until "sometime after 1964.”[556]

Matters changed dramatically immediately after the Castle Bravo nuclear test on March 1, 1954, a weaponizable 15-megaton H-bomb, fully 1,000 times more powerful than the atomic bomb that devastated Hiroshima less than a decade previously. The "Teapot Committee,” chaired by the Hungarian emigree mathematician John von Neumann, had anticipated success with Bravo and with similar tests. Echoing Bruno Augenstein of the Rand Corporation, the Teapot group recom-

It ia well known that for any truly blunt body, the bow shock wave is detached and there exists a stagnation point at the nose. Consider conditions at this point and assume that the local radius of curvature of the body is a (see sketch).

The bow shock wave is normal to the stagnation streamline and converts the supersonic flow ahead of the shock to a low subsonic speed flow at high static temperature downstream of the shock. Thus, it is suggested that conditions near the stagnation point may be investigated by treating the nose section as if it were a segment of a sphere in a subsonic flow field.

Extract of text from NACA Report 1381 (1953), in which H. Julian Allen and Alfred J. Eggers

postulated using a blunt-body reentry shape to reduce surface heating of a reentry body. NASA.

mended that the Atlas miss distance should be relaxed "from the pres­ent 1,500 feet to at least two, and probably three, nautical miles.”[557] This was feasible because the new H-bomb had such destructive power that such a "miss” distance seemed irrelevant. The Air Force leadership con­curred, and only weeks after the Castle Bravo shot, in May 1954, Vice Chief of Staff Gen. Thomas D. White granted Atlas the service’s highest developmental priority.

But there remained the thorny problem of reentry. Only recently, most people had expected an ICBM nose cone to possess the needle – nose sharpness of futurist and science fiction imagination. The realities of aerothermodynamic heating at near-orbital speeds dictated other­wise. In 1953, NACA Ames aerodynamicists H. Julian Allen and Alfred

Eggers concluded that an ideal reentry shape should be bluntly rounded, not sharply streamlined. A sharp nose produced a very strong attached shock wave, resulting in high surface heating. In contrast, a blunt nose generated a detached shock standing much further off the nose sur­face, allowing the airflow to carry away most of the heat. What heating remained could be alleviated via radiative cooling or by using hot struc­tures and high-temperature coatings.[558]

There was need for experimental verification of blunt body theory, but the hypersonic wind tunnel, previously so useful, was suddenly inad­equate, much as the conventional wind tunnel a decade earlier had been inadequate to obtaining the fullest understanding of transonic flows. As the slotted throat tunnel had replaced it, so now a new research tool, the shock tube, emerged for hypersonic studies. Conceived by Arthur Kantrowitz, a Langley veteran working at Cornell, the shock tube enabled far closer simulation of hypersonic pressures and temperatures. From the outset, Kantrowitz aimed at orbital velocity, writing in 1952 that: "it is possible to obtain shock Mach numbers in the neighborhood of 25 with reasonable pressures and shock tube sizes.”[559]

Despite the advantages of blunt body design, the hypersonic envi­ronment remained so extreme that it was still necessary to furnish ther­mal protection to the nose cone. The answer was ablation: covering the nose with a lightweight coating that melts and flakes off to carry away the heat. Wernher von Braun’s U. S. Army team invented ablation while working on the Jupiter intermediate-range ballistic missile (IRBM), though General Electric scientist George Sutton made particularly nota­ble contributions. He worked for the Air Force, which built and success­fully protected a succession of ICBMs: Atlas, Titan, and Minuteman.[560]

Flight tests were critical for successful nose cone development, and they began in 1956 with launches of the multistage Lockheed X-17. It rose high into the atmosphere before firing its final test stage back at Earth, ensuring the achievement of a high-heat load, as the test nose cone would typically attain velocities of at least Mach 12 at only 40,000 feet. This was half the speed of a satellite, at an altitude typically tra­versed by today’s subsonic airliners. In the pre-ablation era, the war­heads typically burned up in the atmosphere, making the X-17 effectively a flying shock tube whose nose cones only lived long enough to return data by telemetry. Yet out of such limited beginnings (analogous to the

rudimentary test methodologies of the early transonic and supersonic era just a decade previously) came a technical base that swiftly resolved the reentry challenge.[561]

Tests followed with various Army and Air Force ballistic missiles. In August 1957, a Jupiter-C (an uprated Redstone) returned a nose cone after a flight of 1,343 miles. President Dwight D. Eisenhower subse­quently showed it to the public during a TV appearance that sought to bolster American morale a month after Sputnik had shocked the world. Two Thor-Able flights went to 5,500 miles in July 1958, though their nose cones both were lost at sea. But the agenda also included Atlas, which first reached its full range of 6,300 miles in November 1958. Two nose cones built by GE, the RVX-1 and -2, flew subsequently as payloads. An RVX-2 flew 5,000 miles in July 1959 and was recovered, thereby becom­ing the largest object yet to be brought back. Attention now turned to a weaponized nose cone shape, GE’s Mark 3. Flight tests began in October, with this nose cone entering operational service the following April.[562]

Success in reentry now was a reality, yet there was much more for the future. The early nose cones were symmetric, which gave good ballistic char­acteristics but made no provision for significant aerodynamic maneuver and cross-range. The military sought both as a means of achieving greater oper­ational flexibility. An Air Force experimental uncrewed lifting body design, the Martin SV-5D (X-23) PRIME, flew three flights between December 1966 and April 1967, lofted over the Pacific Test Range by modified Atlas boost­ers. The first flew 4,300 miles, maneuvering in pitch (but not in cross-range), and missed its target aim point by only 900 feet. The third mission demon­strated a turning cross-range of 800 miles, the SV-5D impacting within 4 miles of its aim point and subsequently was recovered.[563]

Other challenges remained. These included piloted return from the Moon, reusable thermal protection for the Shuttle, and planetary entry into the Jovian atmosphere, which was the most demanding of all. Even

so, by the time of PRIME in 1967, the reentry problem had been resolved, manifested by the success of both ballistic missile nose cone development and the crewed spacecraft effort. The latter was arguably the most sig­nificant expression of hypersonic competency until the return to Earth from orbit by the Space Shuttle Columbia in 1981.

Among their many distinctive attributes, birds possess a particularly unique characteristic not experienced by humans: they are continu­ously aware of the loads their wings and control feathers bear, and they can adjust the wing shape to alleviate or redistribute these loads in real time. This allows a bird to optimize its wing shape across its entire range of flight; for example, a different wing shape for low-speed soar­ing than for high-speed cruising. Humans are not so fortunate. In the earliest days of flight, most aircraft designers consciously emulated the design of birds for both the planform and airfoil cross section of wings. Indeed, the frail fabric and wood structure of thin wings used by pio­neers such as the Wright brothers, Louis Bleriot, the Morane brothers, and Anthony Fokker permitted use of aeroelastic wing-warping (twist­ing) of a wing to bank an airplane, until superseded by the invention of the pivoted aileron. Naturally, when thicker wings appeared, the option of wing-warping became a thing of the past, not revived until the far later jet age and the era of thin composite structures.

For human-created flight, structural loads can be measured via strain gages, and, indeed, the YF-16 utilized strain gages on the main wing spar to adjust the g limiter in the control laws for various fuel loadings and external store configurations. Though the system worked

and showed great promise, General Dynamics and the Air Force aban­doned this approach for the production F-16 out of concern over the relatively low reliability of the strain gages. The technology still has not yet evolved to the point where designers are willing to put the strain gage outputs directly into the flight control system in a load-feedback manner.[711] Certainly this technology will continue, and changing wing shapes based on load measurements will evolve.

The NASA-Air Force Transonic Aircraft Technology (TACT) program, a joint cooperative effort from 1969 to 1988 between the Langley, Ames, and Dryden Centers, and the Air Force Flight Dynamics Laboratory, led to the first significant test of a so-called mission-adaptive wing (MAW), one blending a Langley-designed flexible supercritical wing planform joined to complex hydraulic mechanisms that could vary its shape in flight. Installed on an F-111A testbed, the MAW could "recon­tour” itself from a thick supercritical low-speed airfoil section suitable for transonic performance to a thinner symmetrical section ideal for supersonic flight.[712] The MAW, a "first generation” approach to flexible skin and wing approaches, inspired follow-on work including tests by NASA Dryden on its Systems Research Aircraft, a McDonnell-Douglas (now Boeing) F/A-18B Hornet attack fighter using wing deformation as a means of achieving transonic and supersonic roll control.[713]

NASA DFRC is continuing its research on adaptive wing shapes and airfoils to improve efficiency in various flight environments. Thus, over a century after the Wrights first flew a bird-imitative wing­warping airplane at Kitty Hawk, wing-warping has returned to aero­nautics, in a "back to the future—back to nature” technique used by the Wright brothers (and birds) to bank, and to perform turns. This cutting-edge technology is not yet in use on any operational airplanes, but it is only a matter of time before these performance enhancement features will increase the efficiency of future military and civilian aircraft.

Another potential method for disbursing heat during high-speed flight was the application of an "ablation” material to the outer surface of the structure. An ablator is a material that is applied to the outside of a vehi­cle that burns or chars when exposed to high temperature, thus carry­ing away much of the associated heat and hot gases. Ablators are quite efficient for short duration, one-time entries such as an intercontinen­tal ballistic missile (ICBM) nose cone. Ablators were also used on the early crewed orbiting capsules (Mercury, Gemini, and Apollo), which used ballistic or semiballistic entry trajectories with relatively short peak heating exposure times. They seemed to offer special promise for lifting bodies, with developers hoping to build classes of aluminum-structured spacecraft that could have a cheap, refurbishable ablative coating re­applied after each flight. Indeed, on April 19, 1967, the Air Force did fly and recover one such subscale experimental vehicle, the Mach 27 Martin SV-5D (X-23) Precision Recovery Including Maneuvering Entry (PRIME) lofted over the Pacific Test Range by a modified Atlas ballistic missile.[751]

But for all their merits, ablators are hardly a panacea. Subsonic and transonic testing of several rocket-powered aluminum lifting bodies at NASA’s Flight Research Center showed that this class of vehicle could be landed; however, later analysis indicated that the rough surface of an exposed ablator would probably have reduced the lift and increased the drag so that successful landings would have been questionable.[752]

Flight-test experience with the X-15 confirmed such conclusions. When the decision was made to rebuild the second X-15 after a crash landing, it seemed a perfect opportunity to demonstrate the potential of ablative coatings as a means of furnishing refurbishable thermal pro­tection to hypersonic aircraft and spacecraft. The X-15A-2 was designed to reach Mach 7, absorbing the additional heat load it would experience via MA-25S, a thin Martin-developed silica ablative coating. Coating the aircraft with the MA-25S proved surprisingly time-consuming, as did the refurbishment between flights.

During a flight to Mach 6.7 by Maj. William J. "Pete” Knight, unantic­ipated heating actions severely damaged the aircraft, melting a scramjet boilerplate test module off the airplane and burning holes in the exter­nal skin. Though Knight landed safely—a great tribute to his piloting skills—the X-15A-2 was in no condition to fly without major repairs. Although the ablator did provide the added protection needed for most of the airplane, the tedious process of applying it and the operational problems associated with repairing and protecting the soft coating were quite time-consuming and impracticable for an operational military or civilian system.[753] The postentry ablated surface also increased the drag of the airplane by about the same percentage that was observed on the PRIME vehicle. Clearly the X-15A-2’s record flight emphasized, as NASA engineer John V. Becker subsequently wrote, "the need for maximum attention to aerothermodynamic detail in design and preflight testing.”[754] The "lifting body” concept evolved as a means of using ablative protec­tion for entries of wingless, but landable, vehicles. As a result of the X-15 and lifting body testing by NASA, an ablative coating has not been seriously considered for any subsequent reusable lifting entry vehicle.

David C. Aronstein

NASA research has been pivotal in its support of computational ana­lytical methods for structural analysis and design, particularly through the NASTRAN program. NASA Centers have evolved structural analy­sis programs tailored to their own needs, such as assessing high-tem­perature aerothermodynamic structural loading for high-performance aircraft. NASA-developed structural tools have been adopted through­out the aerospace industry and are available on the Agency Web site.

HE FIELD OF COMPUTER METHODS in structural analysis, and the contributions of the National Aeronautics and Space Administration (NASA) to it, is wide-ranging. Nearly every NASA Center has a struc­tural analysis group in some form. These groups conduct research and assist industry in grappling with a broad spectrum of problems. This paper is an attempt to show both aspects: the origins, evolution, and application of NASA Structural Analysis System (NASTRAN), and the variety and depth of other NASA activities and contributions to the field of computational structural methods.

In general terms, the goal of structural analysis is to establish that a product has the required strength and stiffness—structural integrity— to perform its function throughout its intended life. Its strength must exceed the loads to which the product is subjected, by some safety mar­gin, the value of which depends on the application.

With aircraft, loads derive from level flight, maneuvering flight, gusts, landings, engine thrust and torque, vibration, temperature and pressure differences, and other sources. Load cases may be specified by regula­tory agency, by the customer, and/or by the company practice and expe­rience. Many of the loads depend on the weight of the aircraft, and the weight in turn depends on the design of the structure. This makes the structural design process iterative. Because of this, and also because a large fraction of an aircraft’s weight is not actually accounted for by pri­mary structure, initial weight estimates are usually based on experience

rather than on a detailed buildup of structural material. A sizing pro­cess must be performed to reconcile the predicted empty weight and its relationship to the assumed maximum gross weight, with the required payload, fuel, and mission performance.[786]

After the sizing process has converged, the initial design is docu­mented in the form of a three-view drawing with supporting data. From there, the process is approximately as follows:

• The weights group generates an initial estimate of the weights of the major airframe components.

• The loads group analyzes the vehicle at the defined condition(s) to determine forces, bending moments, etc., in the major components and interfaces.

• The structures group defines the primary load paths and sizes the primary structural members to provide the required strength.

• Secondary load paths, etc., are defined to the required level of detail.

Process details vary between different organizations, but at some point, the structural definition reaches a level of maturity to enable a check of the initial weight estimate. Then the whole designmay be iterated, if required. Iteration may also be driven by maturing requirements or by evolution in other aspects of the design, e. g., aerodynamics, propulsion, etc.