Category AERONAUTICS

Another important area of spacecraft structural modeling is in the inter­action of control systems with flexible multibody structural systems. In a general sense, this is the spacecraft counterpart to aeroservoelasticity, although the driving mechanisms are very different. Dynamic Simulation of Controls & Structure (DISCOS) was developed in the late 1970s to perform this type of analysis. "The physical system undergoing analy­sis may be generally described as a cluster of contiguous flexible struc­tures (bodies) that comprise a mechanical system, such as a spacecraft. The entire system (spacecraft) or portions thereof may be either spin­ning or nonspinning. Member bodies of the system may undergo large relative excursions, such as those of appendage deployment or rotor/ stator motion. The general system of bodies is, by its inherent nature, a feedback system in which inertial forces (such as those due to centrifu­gal and Coriolis acceleration) and the restoring and damping forces are motion-dependent. . . . The DISCOS program can be used to obtain non­linear and linearized time response of the system, interaction constant forces in the system, total system resonance properties, and frequency domain response and stability information for the system. DISCOS is probably the most powerful computational tool to date for the computer simulation of actively controlled coupled multi-flexible-body systems,” according to the computer program abstract. The program was made available to approved licensees (for $1,000, in 1994) with the caveat that DISCOS " . . . is not easy to understand and effectively apply, but is not

intended for simple problems. The DISCOS user is expected to have extensive working knowledge of rigid-body and flexible-body dynamics, finite-element techniques, numerical methods, and frequency-domain analysis.” DISCOS was used extensively at least into the 1990s for spacecraft modeling.[989] In 1983, a program for bridging DISCOS, NASTRAN, and SAMSAN—a large order control system design program—was also publicly released.[990] A 1987 NASA-funded study by Honeywell (Space and Strategic Avionics Division) out­lined some limitations of DISCOS and other contemporary multi­body dynamics programs and made recommendations for future work in the field.[991] Also in the late 1980s, GSFC began collaborat­ing with a research group at the University of Iowa that was devel­oping similar multibody modeling capabilities for mechanical engineering applications. The National Science Foundation (NSF), U. S. Army Tank Automotive Command, and about 30 other Government and industry laboratories were involved in this project through the Industry/University Cooperative Research Center (I/UCRC) at the University of Iowa. Goals of the I/UCRC were to achieve mutual enhance­ment of capabilities in the modeling, simulation, and control of complex mechanical systems, including man/machine interactions applicable to manufacturing processes.[992]

Albert C. Piccirillo

Fly-by-wire (FBW) technology pioneered by NASA has enabled the design of highly unconventional airframe configurations. Continuing NASA FBW research has validated integrated digital propulsion and flight control systems. FBW has been applied to civil aircraft, improving their safety and efficiency. Lessons learned from NASA FBW research have been transferred to maritime design, and Agency experts have supported the U. S. Navy in developing digital electronic ship control.

HE EVOLUTION OF ADVANCED AIRCRAFT equipped with comput­erized flight and propulsion control systems goes back a long way and involved many players, both in the United States and interna­tionally. During the Second World War, use of electronic sensors and subsystems began to become pervasive, adding new mission capabili­ties as well as increasing complexity. Autopilots were coupled to flight control systems, and electric trim was introduced. Hydraulically or electrically boosted flight control surfaces, along with artificial feel and stability augmentation systems, soon followed as did electronic engine control systems. Most significantly, the first uses of airborne computers emerged, a trend that soon resulted in their use in aircraft and missiles for their flight and mission control systems. Very early on, there was a realization that traditional mechanical linkages between the cockpit flight controls and the flight control surfaces could be replaced by a computer – controlled fly-by-wire (FBW) approach in which electric signals were trans­mitted from the pilot’s controls to the control surface actuators by wire. This approach was understood to have the potential to reduce mechani­cal complexity, lower weight, and increase safety and reliability. In addi­tion, the processing power of the computer could be harnessed to enable unstable aircraft designs to be controlled. Properly tailored, these unsta­ble designs could enable new aircraft concepts to be implemented that could fully exploit the advantages of active flight control. Such aircraft

would be more maneuverable and lighter, have better range, and also allow for the fully integrated control of aircraft, propulsion, navigation, and mission systems to optimize overall capability.

Two significant events unfolded during the 1960s that fostered the move to electronic flight control systems. The space race was a major influence, with most space systems relying on computerized fly-by-wire control systems for safe and effective operation. The Vietnam war pro­vided a strong impetus for the development of more survivable aircraft systems as well for new aircraft with advanced performance features. The National Aeronautics and Space Administration (NASA) and the Air Force aggressively responded to these challenges and opportunities, resulting in the rapid transition of digital computer technology from the space program into aircraft fly-by-wire applications as exemplified by the Digital Fly-By-Wire F-8 and the AFTI/F-16 research programs. Very quickly after, a variety of flight research programs were implemented. These programs provided the basis for development and fielding of numerous military and civil aircraft equipped with advanced digital fly­by-wire flight control systems. On the civil aviation side, safety has been improved by preventing aircraft flight envelope limitations from being exceeded. Operating efficiency has been greatly enhanced and major weight savings achieved from fly-by-wire and related electronic flight control system components. Integrated flight and propulsion control systems precisely adjust throttles and fuel tank selections. Rudder trim drag because of unbalanced engine thrust is reduced. Fuel is automati­cally transferred between tanks throughout the aircraft to optimize cen­ter of gravity during cruise flight, minimizing elevator trim drag. In the case of advanced military aircraft, electronically controlled active flight, propulsion, and mission systems have been fully integrated, providing revolutionary improvements in capabilities. Significantly, new highly unstable aircraft configurations are providing unprecedented levels of mission performance along with very low radar signatures, capabilities that have been enabled by exploiting digital fly-by-wire flight and pro­pulsion control systems pioneered in NASA.

By the mid-1970s, the Air Force Flight Dynamics Laboratory had initiated a Control Configured Vehicle flight research program to investigate the use of nonconventional (often called "decoupled”) movements of aircraft flight control surfaces to enable maneuvers in one plane without move­ment in another. An example of such a maneuver would be a wings-level turn without having to bank the aircraft. The very first General Dynamics YF-16 technology demonstrator aircraft (USAF serial No. 72-1567) was selected for modification. Rebuilt in December 1975 and fitted with twin
vertical canards underneath the air intake, it became known as the F-16/CCV. Its flight controls were modified to enable the wing trail­ing edge flaperons to move in combination with the all-moving stabi – lator. In addition, the fuel system in the YF-16 was modified to enable the aircraft center of gravity to be adjusted in flight by transferring fuel between tanks, thus allowing the stability of the aircraft to be varied. The YF-16/CCV flew for the first time on March 16, 1976, piloted by GD test pilot David J. Thigpin. On June 24, 1976, while being flown by David Thigpin, it was seriously damaged in a crash landing. Engine power had been lost on final approach, and the landing gear collapsed in the sub­sequent hard landing. Repairs to the aircraft would take over 6 months. The F-16 CCV returned to flight in the spring of 1977. It would complete 87 flights and 125 flying hours before the research program ended, with the last flight F-16 CCV flight on June 30, 1977.[1178]

The F-16 CCV anticipated CCV flight-test approaches that were soon undertaken by a number of foreign countries. Flight research projects in Germany, the U. K., and Japan converted existing military aircraft into CCV testbeds. Fitted with computer-controlled fly-by-wire flight control systems, these projects provided experience and insights that enabled these countries to incorporate fly-by-wire flight control into their next generations of advanced civil and military aircraft. These foreign CCV projects are discussed in a separate section of this report. Experience gained with the F-16 CCV served as the basis for the subsequent Flight Dynamics Laboratory AFTI/F-16 program, which would yield valuable insights into many issues associated with developing advanced DFBW flight control systems and result in significantly improved capabilities being incorporated into U. S. military aircraft.

The Grumman X-29 research aircraft played a very interesting role in the evolution of modern fly-by-wire flight control systems. Exotic in appearance, with its forward-swept wings and large movable canard control surfaces, the X-29 was highly unstable about the longitudinal axis with a static stability margin of -35 percent. This level of instability probably indicates that the X-29 represents the most unstable piloted aircraft that has ever been successfully flown. Not only did it fly, but it also demonstrated good controllability at very high angles of attack. This degree of success was only possible through the use of a very advanced fly-by-wire flight control system that employed a combination of both digital computers in the primary system and analog computers in the backup system. The program began in 1977, with the Defense Advanced Research Projects Agency (DARPA) and the Air Force Flight Dynamics Laboratory jointly soliciting industry proposals for a research aircraft designed to investigate the forward-swept wing concept in a high-per­formance aircraft application. In December 1981, Grumman Aircraft

was selected to build two aircraft, which were designated X-29. The most unique and visually obvious aspect of the design was the forward – swept wings that incorporated a thin supercritical airfoil, but there were many other areas of the X-29 design that embodied advanced tech­nology. The aircraft used advanced composite materials and unusual construction approaches. Its control system made use of variable cam­ber wing surfaces, aft fuselage-mounted strake flaps, fully movable canards mounted on the sides of the engine inlets, and a computerized fly-by-wire flight control system to maintain control of the otherwise highly unstable aircraft.[1221]

In constructing the X-29s, Grumman used the forward fuselage and nose landing gear from two Northrop F-5A fighters. Control surface actuators and the main landing gear came from the F-16. The unique aspect of the X-29 airframe, its forward-swept wing, was developed by Grumman. Because of the major differences between the X-29s and the F-5As, the modified aircraft were assigned new USAF serial numbers, becoming 82-0003 and 82-0049. The main difference between the two

X-29s was the emergency spin parachute system mounted at the base of the rudder on the second aircraft. The X-29 flight research program was conducted by the NASA Dryden Flight Research Center in two phases and included participation by the Air Force Flight Test Center and the Grumman Corporation.[1222] The joint NASA-Air Force portion of the X-29 test program extended from 1984 to 1991; the Air Force conducted a follow-on investigation of vortex flow control (VFC) that lasted into 1992.

The X-29’s thin supercritical forward-swept wing presented signif­icant design challenges. The typical stall pattern of an aft-swept wing, from wingtip to root, is reversed for a forward-swept wing, which stalls from the root to the tip. The aerodynamic lift forces on the outer por­tions of a forward-swept wing produce a twisting moment that tends to force the leading edge further upward. This increases the angle of attack at the wingtips, causing even further twisting that, if uncontrolled, can lead to structural failure of the wing, a phenomenon known as aero – elastic divergence. To deal with this problem, Grumman made use of state-of-the-art composite materials in designing the wing external skin, which it laminated in a way that produced an inherent coupling between wing bending and torsion loads, a concept known as aeroelastic tailor­ing. At increasing angles of attack (higher lift), the structural character­istics incorporated into composite laminates were designed to ensure that the wing twisted to counter the upward twist produced by aerodynamic loads. The key to the design was balancing the aerodynamic aspects of the wing’s configuration with the structural characteristics of the composite laminate to control potential aeroelastic divergence. The wing substruc­ture and the basic airframe itself were made of aluminum and titanium.[1223]

The X-29 featured an unusual combination of flight control sur­faces. These consisted of forward-mounted canards that contributed positive lift and provided primary control about the pitch axis. Wing flaperons (combination flaps and ailerons) could change the camber of the wing and also functioned as ailerons for roll control. The actua­tors used to control wing camber were mounted externally in stream­lined fairings at the trailing edge of the wing because of the thinness of the supercritical airfoil. The strake flaps on each side of aft fuselage augmented the canards, proving additional pitch control. The control sur­faces were electronically linked to a triple-redundant digital fly-by-wire
flight control system (with analog backup) that provided artificial sta­bility necessary for controlling the inherently unstable forward-swept wing, close-coupled canard design used on the X-29. Each of the three digital flight control computers had an analog backup. If one of the digital computers failed, the remaining two took over. If two of the digital computers failed, the flight control system switched to the ana­log mode. If one of the analog computers failed, the two remaining ana­log computers took over.[1224]

Grumman chief test pilot Charles A. "Chuck” Sewell flew the first X-29 at Edwards AFB on December 14, 1984.[1225] During the Phase I research effort, X-29 aircraft No. 1 was used exclusively, flying 242 times. Its wingtips remained unstalled up to the 21-degree angle of attack allowed in Phase I testing. This limitation was due to the fact that an anti-spin parachute was not installed on the aircraft. The aeroelas – tic tailored wing prevented structural divergence of the wing, and the digital flight control system functioned safely and reliably. Flight con­trol laws and control surface effectiveness combined to provide good pilot handling qualities during maneuvering flight. The aircraft’s supercritical airfoil contributed to enhanced cruise and maneuver performance in the transonic regime.[1226] However, overall drag reduc­tion was not as great as had been predicted for the configuration. On December 13, 1985, with NASA test pilot Steve Ishmael at the con­trols, the X-29 became the first aircraft with a forward-swept wing to fly beyond the speed of sound, reaching Mach 1.03 in level flight.[1227] Other test pilots who flew the X-29 during Phase I of the joint test program were NASA test pilot Rogers Smith, Lt. Col. Theodore "Ted” Wierzbanowski and Maj. Harry Walker from the Air Force, and Navy Cdr. Ray Craig.

The second X-29 aircraft, modified to incorporate an anti-spin para­chute and its deployment mechanism, was used during Phase II test­ing to investigate the aircraft’s high-angle-of-attack characteristics and the potential usefulness of the forward-swept wing and canard config­uration on military fighter plane designs. First flown on May 23, 1989, it would eventually fly 120 research flights and demonstrate control and maneuvering qualities that were better in many cases than the pre­dictions derived from computational methods and simulation models. NASA, Air Force, and Grumman project pilots reported that the X-29 had excellent control response up to angles of attack of 45 degrees, with limited controllability still available at up to 67 degrees angle of attack. Phase II flight-testing defined an allowable X-29 flight envelope that extended to Mach 1.48, an altitude of just over 50,000 feet, an angle of attack of up to 50 degrees at 1 g and 35 degrees at airspeeds up to 300 knots. Much of the X-29’s high-angle-of-attack capability was attributed to the quality of the flight control laws that were cooperatively devel­oped by NASA and the Air Force. These had initially been developed using results obtained from extensive wind tunnel testing and predic­tions derived from radio-controlled flight tests of a 22-percent scale drop model at NASA Langley Research Center.[1228] Flight control system engi­neers at NASA Dryden and the Air Force Flight Test Center at Edwards used these as the basis for detailed flight control system design. This design used a combination of pitch rate and angle of attack in develop­ing the longitudinal control laws. Selectable gain was included in the flight control system design, and this was used by the X-29 test pilots during flying qualities assessments and evaluations of the effects of control law gain changes at higher angles of attack. Prior to the start of flight-testing, wing rock was estimated to restrict the available angle of attack to less than about 35 degrees. However in flight-testing, wing rock amplitude was found to be less than half of what had been predicted, allowing the roll rate to aileron gain to be lowered to one-fourth of the value that had been derived from preflight data using the subscale free flight model. The available flight envelope was extended to 67 degrees angle of attack at 1 g. Maneuvering flight about all axes was cleared up to an angle of attack of 45 degrees in 1 g flight. The reduced wing rock that had been observed in flight was apparently due to higher roll
damping and increased aileron control power for large aileron deflections.[1229]

Preflight predictions of the X-29’s pitch capabilities matched flight – test results up through 40 degrees angle of attack. Differences in nose – up pitching moment above an angle of attack of 40 degrees were found to require more canard deflection than predicted. Large yaw asymme­tries led to several instances in which the aircraft tended to stabilize at very high nose-up pitch angles during maneuvers at angles of attack above 50 degrees, with the aircraft at an aft center of gravity. Modifications to pro­vide additional nose-down pitch authority were not possible because of physical limits on canard deflection. Maximum pitch rates were limited by the high level of static instability inherent in the X-29 design and control surface rate limits. New actuators with at least a 50-percent higher actu­ation rate would have been required to achieve pitch rates comparable to those of an operational fighter like the F/A-18. The full wingspan flaper – ons were found to provide good roll control that was not affected by the fact that the X-29 did not use wing leading-edge maneuvering flaps. Pilot – selectable variable gain capability was used during examination of air­plane stability and maneuverability. Basic fighter maneuvers were flown, and roll and yaw gains were increased to improve roll performance. A gain that provided maximum rudder authority resulted in the best pilot com­ments. Roll coordination was better than anticipated, with rudder effec­tiveness also higher than preflight predictions at angles of attack between 20 and 40 degrees. Yaw asymmetries developed above 40 degrees angle of attack. Diminished aileron and rudder power was not sufficient to over­power these asymmetries. Increasing gain further produced rudder sat­uration, resulting in uncoordinated turns, a result that was disliked by test pilots, even though this actually resulted in better roll performance.[1230]

Flight-test data from Phase II, the high-angle-of-attack and mili­tary utility phase of the X-29 program, satisfied the program’s primary objective. The technologies demonstrated in the program had poten­tial to improve future fighter aircraft mission performance, and the forward-swept wings/movable canard configuration provided excel­lent control response at up to 45 degrees angle of attack. Very impor­tantly, the X-29A program provided a significant pool of knowledge that
was very useful background for fine-tuning ground-based predictive techniques for high-angle-of-attack aircraft.[1231] One significant poten­tial safety issue with the flight control system, the danger of sensor selection thresholds being set too wide, was discovered during the X-29 test program. The flight control system used three sources of air data in its computations. These air data sources were the nose probe and two probes mounted one on each side of the forward fuselage. The selection algorithm in the flight control system used the data from the nose probe as the primary source, provided it was within some threshold of the data from both side probes. However, the data selection threshold was inten­tionally large to accommodate known data errors in certain flight modes because of the position of the side sensors in the airflow. Long after the start of flight-testing, ground simulation revealed that the nose probe could, in some circumstances, furnish erroneous information at very low flight speeds, causing the X-29 to go out of control. Although this fault was successfully identified through ground simulation, 162 flights had already been flown before it was detected and corrected.[1232]

In 1992, the Air Force began a follow-on program with X-29 No. 2 that investigated the use of vortex flow control as a means of providing increased aircraft control at very high angles of attack, at which normal rudder control is ineffective. Wind tunnel tests had showed that injec­tion of air into the vortexes coming off the nose of the aircraft would change the direction of vortex flow. The forces created on the nose of the aircraft could be used to control directional (yaw) stability. The second X-29 aircraft was modified to incorporate two high-pressure nitrogen tanks, related control valves with two small nozzle jets on the forward upper portion of the nose. The nozzles injected air into the vortexes, which flowed off the nose of the aircraft at high angles of attack. From May to August 1992, 60 test flights were flown. Data from these flights were used to determine that VFC was more effective than expected in generating yaw forces, but it was less successful in providing control when sideslip was present, and it did little to decrease roll oscillations.[1233] The two X-29 aircraft flew a total of 436 flights, 254 by the first, and 182 by the second. The former is exhibited at the National Museum of the

United States Air Force, while the latter remained at Edwards and is on exhibit at the Dryden Flight Research Center.[1234]

Approved in 1975 and begun in 1976, the Aircraft Energy Efficiency (ACEE) program was managed by NASA and funded through 1983, as yet another round of research and development activities were put in work to improve the state of the art of aircraft structural and propulsion design. And once again, the program was aimed at pushing the technological envelope to see what might be possible. Then, based on that information, new Government regulations could be enacted, and the airline industry could decide if the improvements would offer a good return on its investment. The answer, as it turned out, was an enthusiastic yes, as the overall results of the pro­gram led directly to the introduction of the Boeing 757 and 767.[1307]

Driving this particular program was the rapid increase in fuel costs since 1973 and the accompanying energy crisis, which was brought on by the Organization of Arab Petroleum Exporting Countries’ decision to embargo all shipments of oil to the United States. This action began in October 1973 and continued to March 1974. As a result of this and other economic influences, the airlines saw their fuel prices as a per­centage of direct operating costs rise from 25 percent to as high as 50 percent within a few weeks. With the U. S. still vulnerable to a future oil embargo, along with general concerns about an energy shortage, the

Federal Government reacted by ordering NASA to lead an effort to help find ways for airlines to become more profitable. Six projects were ini­tiated under the ACEE program, three of which had to do with the air­craft structure and three of which involved advancing engine technology. The aircraft projects included Composite Structures, Energy Efficient Transport, and Laminar Flow Control. The propulsion technology proj­ects included Engine Component Improvement, Energy Efficient Engine, and Advanced Turboprop—all three of which are detailed next.[1308]

While most of NASA’s aircraft fuel-efficiency research grew out of the reality jolt of the 1970s oil crisis and environmental concerns, there is at least one notable exception. Researchers first began to investigate liq­uid hydrogen as an alternative to hydrocarbon-based fuel in the mid – 1950s because they suspected major performance efficiencies could be gained.[1468] The Lewis Flight Propulsion Laboratory issued a seminal report in April 1955—although it was not declassified until September 1962—suggesting that liquid hydrogen might have a positive impact on the performance of high-altitude military aircraft (subsonic and super­sonic bombers, fighters, and reconnaissance aircraft flying at 75,000 to 85,000 feet).[1469] While the report raised the aviation community’s aware­ness of the potential for hydrogen as a fuel source, it did not lead to widespread use in aircraft because of technical problems with using

hydrogen inside an aircraft. Again, this early interest in hydrogen did not reflect environmental or conservation concerns, but rather, a desire to achieve much higher flight vehicle performance.

In the report, two NACA researchers—Abe Silverstein and Eldon Hall—argued the case for using liquid hydrogen, noting it has a spe­cial advantage as an aviation fuel: a high heating value. This means that it takes less hydrogen fuel than hydrocarbon fuel to achieve the same thrust. That advantage could prove particularly important at high alti­tudes, the researchers noted, where maximizing fuel efficiency is criti­cal to make up for other penalties associated with flying high.

Indeed, one of the downsides of a high-altitude flight regime—in which atmospheric pressure is low—is that it generally requires heavy, high-pressure-ratio engines to ensure combustion is sustained and thrust levels are adequate. The NACA report speculated that it might be possi­ble to use lighter engines—albeit less-efficient ones, with lower pressure ratios—if liquid hydrogen were used for fuel. Liquid hydrogen requires less combustion volume than hydrocarbons do, making shorter and lighter engines feasible. And, with its high heating value, liquid hydro­gen fuel generates more thrust per pound than hydrocarbons do, even if it’s being used in a light engine running at a lower pressure ratio. The report posited that if every pound of weight saved by using a lighter

engine could be replaced by a pound of liquid hydrogen fuel, an aircraft could be over twice as effective in extending its range at high altitudes.

After Silverstein and Hall issued their report, the NACA conducted experiments showing that hydrogen had a high combustion efficiency in a turbojet combustor even in low-pressure conditions. In 1956, NACA researchers at Lewis made "three completely successful flights” using liquid hydrogen in one engine of a modified Martin B-57B jet bomber, thereby effectively demonstrating that liquid hydrogen could be handled and jettisoned safely and was feasible for use in aircraft.[1470] Meanwhile, from 1956 to 1958, the U. S. Air Force began work on a secret project, known as Suntan, to develop a high-altitude, hydrogen-fueled aircraft with performance superior to the secret U-2 spy plane of the Counter Intelligence Agency (CIA).

The use of liquid hydrogen in aircraft would have marked a major breakthrough in terms of high altitude flight because engine weight is

"the single most important variable determining the height at which an airplane can fly.”[1471] Liquid hydrogen offered the potential to fly at high altitudes at an extended range. Despite its potential, however, nei­ther the NACA nor the Air Force was able to convince enough stake­holders in Government and industry that liquid hydrogen was a viable candidate for aviation fuel.

Hydrogen’s excellent combustion qualities raised questions about whether it could be safely transported or carried inside aircraft. To be sure, NACA flight tests demonstrated that safe handling of hydrogen fuel on the ground and in the air was possible. Also, the Air Force conducted tests in which liquid hydrogen tanks under pressure were ruptured, and it found that in many cases the hydrogen quickly escaped without igni­tion. Yet concerns about safety persisted, and, in a tight budget climate, hydrogen-fuelled aircraft lost out to other priorities. After receiving a full briefing on Suntan, Gen. Curtis E. Lemay, the former head of Strategic Air Command who had moved up to Vice Chief of Staff in July 1957, raised concerns about safety. "What,” he said, "put my pilots up there with a bomb?”[1472]

Others questioned whether using liquid hydrogen would truly yield big gains in aircraft range at high altitudes. Hydrogen has a high vol­ume—10 times that of hydrocarbons—which means that the aircraft fuselage has to be bigger and weigh more to accommodate the fuel. Silverstein and Hall argued that there would be more room for hydro­gen fuel tanks in high-altitude aircraft, which would need larger wings and therefore a bigger fuselage to provide lift in the thin air of the upper atmosphere (a bigger fuselage would mean more room for hydrogen fuel tanks). But while it might have been possible to extend the range of air­craft because of the increased efficiency of liquid hydrogen, others ques­tioned whether hydrogen-fuelled aircraft would still be fairly limited by the tremendous amount of fuel storage capacity that hydrogen requires. Kelly Johnson, the Lockheed Martin engineer who designed the U-2 and the hydrogen-fueled CL-400 for the Air Force’s Suntan project, said he could see a range growth of only 3 percent from adding more hydrogen fuel storage capacity to his CL-400 design. "We have crammed the max­
imum amount of hydrogen in the fuselage that it can hold. You do not carry hydrogen in the flat surfaces of the wing,” he said.[1473]

While liquid hydrogen is highly energetic, it has far less energy density than hydrocarbon fuels. Thus, to get an equivalent amount of energy from hydrogen requires a much greater volume. Accordingly, a hydrogen airplane would have extremely large fuel tanks, which, hav­ing to be supercold as well, pose significant technical challenges to air­craft designers. Researchers have not yet found a way to overcome the challenges associated with hydrogen’s large volume, which forces air­craft design compromises and requires complex ground transportation, storage distribution, and vent capture system. Moreover, hydrogen is not a viable source of energy in itself; producing it requires the use of other sources of energy—such as electric power produced by nuclear fusion as well as a large source of clean water. However, in one respect, hydrogen could "pay back” this "debt,” for it could be used to enrich the production process of synthetic fuel, achieving similar production efficiencies while reducing the amount of water and coal traditionally required for enrichment.[1474]

Despite these technical challenges, NASA’s research on the use of hydrogen to power aircraft did lead to some important findings: namely, that hydrogen is a potentially promising turbojet fuel in a high-altitude, low-speed flight regime. These conditions favor a fuel that can operate efficiently in low-pressure conditions. High altitudes also favor a large – volume aircraft, helping to offset the disadvantage of hydrogen’s low den­sity. Given these attractive characteristics, the prospect of using hydrogen as an aircraft propellant has continued to resurface in the past decade, especially when the cost of hydrocarbon-based fuel rises. For example, NASA’s Zero CO2 research project sought to eliminate carbon dioxide and lower NOx emissions by converting propulsion systems to hydro­gen fuel.[1475] One new propulsion technology that NASA engineers consid­ered as part of Zero CO2 was the use of fuel cells, which are discussed below. A NASA Glenn Web page updated as recently as 2008 says that the Combustion Branch of NASA’s Propulsion Division is still studying
hydrogen combustion to demonstrate that hydrogen can be used as an aviation fuel to minimize emissions.[1476]

The NACA’s early research on hydrogen-fuelled aircraft also cre­ated an awareness of hydrogen as a potential fuel source that did not exist before Silverstein and Hall embarked on their study. This aware­ness helped lead to important breakthroughs in rocketry and fuel cell research. In particular, research on the use of hydrogen in air-breath­ing aircraft laid the groundwork for the successful development of hydrogen-fueled rockets in the mid-1950s. In fact, Silverstein and Hall’s research helped to inform NASA’s decision in 1959 to use liquid hydro­gen as a propellant in the upper stage of the Saturn launch vehicle. That decision was one of the keys to the success of the Apollo Moon landing missions of the 1960s and 1970s.[1477]

The NACA’s early efforts to draw attention to hydrogen as a power source also led to the development of fuel cells for the Apollo and Gemini capsules. Apollo employed the world’s first fuel cells, which used hydro­gen and oxygen to generate onboard power for Apollo command and ser­vice modules. Fuel cells are essentially plastic membranes treated with a special catalyst; hydrogen seeps into the membrane and meets up with the oxygen inside to generate electricity and water. The fuel cells used on the Apollo proved so successful that they were once again employed on the Space Shuttle orbiter.

The first-generation ERAST solar-test program HALE vehicle was the Pathfinder, which was designed and built by AeroVironment. AeroVironment’s earlier solar aircraft projects included the building of the piloted Gossamer Albatross and a scaled-down version known as the Gossamer Penguin. This experience assisted the company in building the Pathfinder UAV. In addition to Ray Morgan, who was vice president of AeroVironment, the company team included a number of experienced engineers and technicians, including William Parks, who was the com­pany’s chief engineer for the Centurion and Helios Prototype UAVs, and Robert Curtin and Kirk Flittie, who both served as project and later as program managers. The program brought honors for Bob Curtin and Ray Morgan, who both received the Aviation Week Laurel Award in 1996.[1536]

Pathfinder, which initially was battery-powered, was a remotely piloted flying wing that demonstrated a number of technologies, includ­ing lightweight composite structures, low wing loading flying-wing configuration, redundant and fault tolerant flight control systems,

high-efficiency electric motors, thermal control systems for high-alti­tude flight, and a high specific power solar array. The remotely piloted Pathfinder had 6 electric motors that each weighed 13 pounds and con­sisted of a fixed-pitch 79-inch propeller and a solid-state motor with inter­nal power electronics, nacelle, and cooling fins. Differential power to two wingtip motors on either side was used for lateral control. Wing dihedral (upsweep) provided roll stability, and 26 elevator control surfaces were attached to the wing’s trailing edge for pitch control. Pathfinder’s solar array generated approximately 8,000 watts near solar noon. The solar UAV could obtain an airspeed of between 15 and 25 mph and a cruising speed of between 17 to 20 mph. The vehicle had a length of 12 feet, a wingspan of 98.4 feet, a wing chord (front to rear distance) of 8 feet, a gross weight of approximately 560 pounds, a payload capacity of up to 100 pounds, a wing aspect ratio (the ratio between the wingspan and the wing chord) of 12 to 1, and a power-off glide ratio of 18 to 1. Pathfinder had a maxi­mum bank rate of 5 degrees and a maximum turn rate of 3 degrees per second at sea level and 1.7 degrees at 60,000 feet.[1537]

To gain some introductory understanding and experience with the challenges and nuances of solar cell operation, prior to the official start of the ERAST program and the transfer of Pathfinder from BMDO, the project team had arranged for some solar cell flight tests on local sorties over the Edwards dry lake. Pathfinder itself was not equipped with solar arrays on early test flights because the ERAST alliance did not want to risk any damage to the expensive solar cell arrays until Pathfinder’s fly­ing capabilities could be tested. Pathfinder’s gross weight of 560 pounds produced a very low wing-loading load distribution of less than 0.64 pounds per square foot that significantly increased sensitivity to winds during takeoff and landing. This necessitated special training for the ground controllers, especially during takeoff and landing of the UAV. Pathfinder’s first foray to high altitude took it to 50,500 feet and proved immensely productive. "We learned tons from that flight,” John Del Frate recalled afterward, noting, "There were a lot of naysayers that were qui­eted after that flight.”[1538] Unfortunately, afterward the vehicle dramati­cally demonstrated its sensitivity to wind, being seriously damaged in its hangar when ground crews opened both hangar doors, thus creat­ing a draft that blew Pathfinder into a jet that was in the same hangar.

After a number of developmental flights and further modifications at Dryden, Pathfinder was transported to the Island of Kauai, which offered a more favorable wind environment and a greater operational area with less competing air traffic. From testing at Edwards Air Force Base, the NASA-industry alliance learned that weather factors—including wind, turbulence, cloud cover, humidity, temperature, and pressure—were crit­ical in attempting to fly the wing-loaded Pathfinder at high altitudes. In addition, the team noted that the UAV s would probably not be flying in the same conditions as found in standard atmosphere reference tables because testing indicated a surprising variance in actual temperature in comparison with the tables. The team also noted that the higher a solar UAV flies, the greater the downwind drift distance if activation of a flight termination system (FTS) is required.[1539] These factors required careful study of historical weather patterns to determine the optimum

site to attempt to set a world record UAV altitude flight. Accordingly, NASA selected the Navy’s Pacific Missile Range Facility in Hawaii as the location to test the high-altitude capabilities of the solar UAVs.[1540]

In Hawaii, Pathfinder was flown for seven additional flights, one of which in 1997 set a world record of over 71,500 feet for a high-altitude flight by a propeller-driven aircraft. This broke a 1995 Pathfinder test alti­tude flight record of 50,500 feet, which had earned NASA recognition as 1 of the 10 most memorable record flights of 1995. Pathfinder test flights also demonstrated solar-powered HALE vehicles’ potential as platforms for environmental monitoring and technical demonstration missions by gaining additional information relating to the Island of Kauai’s terrestrial and coastal ecosystems. These science missions, which employed specially build lightweight sensor systems (see below), included detection of for­est nutrient status, forest regrowth after damage from Hurricane Iniki in 1992, sediment concentrations in coastal waters, and assessment of coral reef conditions. Experience from Pathfinder test flights, in combination with other UAV testing, also yielded a number of lessons regarding hard­ware reliability, including the following recommended procedures: (1) testing the airframe structure as much as possible before flight, particu­larly the composite airframe joint bondings; (2) testing the vehicle’s sys­tems in an altitude chamber because of the extreme cold and low-pressure conditions encountered by high-altitude science aircraft; (3) recognizing that UAVs, like aircraft, have a tendency to gain weight; (4) maintaining strict configuration control; and (5) ensuring that a redundant system is functional before switching from the primary system.[1541]

The transformation of the NACA into NASA in 1958 was marked by an inevitable subordination of the NACA’s aeronautical research char­ter to NASA’s mandated space mission work. The assigned aeronau­tics staff dropped over 80 percent, from 7,100 to 1,400, as the space program gained momentum in the early 1960s. In the new space – focused environment, aeronautics needed to be product-oriented to attract budget allocation support. In these circumstances, helicop­ter research lost ground as the focus shifted to new nonrotor Vertical Take-Off and Landing (VTOL) and Short Take-Off and Landing air­craft. In many cases, the rotary wing work formed the base for VTOL investigations. In the case of NACA-NASA rotor-flow studies, exper­imental and theoretical studies on rotor-time-averaged inflow led to extensive work on establishing wind tunnel jet-boundary layer (wall interference) correction methodology for other VTOL, as well as rotor – borne, lifting systems.[283]

In a sense, it became the U. S. Army’s turn to bolster NASA rotary wing endeavors in support of the Army’s need for continued helicop­ter development. In 1965, the Army was granted permission to reacti­vate, staff, and utilize the Ames 7- by 10-foot Tunnel No. 2. In addition, the Army provided personnel to assist Ames in carrying out projects of interest to the Army. A group of about 45 people was established by the Army and identified as the Army Aeronautical Activity at Ames (AAA – A).[284] In 1970, the working relationship between NASA and the Army was significantly enhanced. Co-located Army research organizations were established at Ames, Langley, and Lewis (now Glenn) Research Centers. They focused on the respective Center’s specialty of aeroflight dynamics, structures, and propulsion. This teaming laid the solid groundwork for major rotary wing programs that NASA and the Army jointly planned, executed, and funded in the 1970s and 1980s that influenced both mil­itary and civilian rotary wing aircraft development.

One of the unique research facilities authorized in 1939 and oper­ated by the NACA, and then NASA, was the 40- by 80-foot Full-Scale Tunnel at Ames. This research facility also provided the opportunity to work directly with industry on vehicle development programs. In the case of rotary wing aircraft, the tunnel was utilized for investigating new vehicle and rotor system concepts and for thoroughly documenting the basic aerodynamic behavior of prototype and production articles. By the 1960s, numerous in-house and industry full-scale rotary wing hardware were tested. Examples include the Bell XV-1 "convertiplane” in 1953­1954, followed by many other projects, including a modified production rotor incorporating leading edge camber and boundary-layer control; the Bell UH-1 "Huey” helicopter (tested to assist in the development of a high-performance flight-test helicopter); a folded rotor with test data obtained in start-stop and folding conditions at forward speeds; and a four-bladed rotor investigation with extensive rotor-blade pressure mea­surements taken as a followup to prior flight test measurements made at Langley Research Center.[285]

The pressure-instrumented blade used in the latter tests had an extremely limited operating life of only 10 hours. This was because of the installation of nearly 50 miniature differential pressure transduc­ers inside the rotor blade. This required that a total of almost 100 small holes be drilled in the upper and lower surface of the primary structure D-spar—normally an absolute "safety of flight” violation. The conserva­tive 10-hour limit was based upon conservative crack-growth-rate limits determined from blade specimen cyclic load tests. The earlier flight test investigation of blade pressure distributions produced a very significant contribution as a primary database for the understanding of basic rotor unsteady aerodynamics. The tabulated pressure data provided time his­tories of individual differential pressures and simultaneous blade bend­ing moments around the rotor azimuth in a wide assortment of steady and maneuvering flight conditions.[286] This database became the standard experimental data reference source for advancing theoretical comparison work for many years. As an aside, in working with the original flight data to hand-digitize the detailed recordings of differential pressure time-his­tory traces, it became possible, in time, to visually recognize the specific flight-test condition by the periodic pressure trace signature.[287] It was pos­sible to identify the rotor’s actual flight condition relative to the surround­ing airmass. This still raises the question of the possibility of applying modern signal recognition technology to provide on-board safety-of-flight and noise abatement operating boundary displays for the pilot.

Flying qualities flight investigations emphasized the importance of ample damping of angular velocity and of control power (rotor-gener­ated aircraft pitch and roll control moments) and their interaction. This work at Langley and similar work at Ames provided a significant portion of the helicopter flying qualities criteria. This early work was extended to the use of in-flight simulation using Langley’s YHC-1A tandem rotor helicopter with special onboard computing and recording equipment.[288]

The flight operations of most interest were terminal area instrument flight on steep approaches to vertical touchdown landings. The results of this work were initially oriented to nonrotor VTOL operations, but the results were found to be equally applicable to helicopters.

In the area of structural dynamics, investigations addressing the problems of aeroelastic stability of rotor-powered aircraft were con­ducted utilizing new analytical methods and experimental studies by Langley and Ames researchers. Emphasis was placed on tilt rotor and tilt propeller (i. e., tilt wing) aircraft concepts. Two-degree-of-freedom "air resonance” (akin to rotor-fuselage "ground resonance”) and prop – rotor/propeller whirl instability were among the problems investigated.[289] Rotor-pylon-wing aeroelastic instability problems for tilt rotor designs were explored in the Ames 40 by 80 Full-Scale Tunnel in this period. The aeroelastic stability problems of the tilt rotor and tilt-stopped rotor designs were also investigated at model scale in the unique Freon atmo­sphere of the Langley Transonic Dynamics Tunnel, which provided full – scale Mach number and Reynolds number scaling.[290] These research

investigations resulted in significant contributions to the development of the validated design tools for advanced rotorcraft.

With the increased interest in hingeless rotor concepts, NASA obtained and quickly accomplished flight research with a copy of an experimental Bell Helicopter three-bladed hingeless rotor installed on an H-13 helicopter.[291] Early experience with "rigid” rotors had led the NACA to encourage interest in exploring the possibilities of removing conventional blade-root hinges and substituting instead blade struc­tural flexibility. Another manufacturer, Lockheed Aircraft, made a major commitment to the hingeless rotor concept coupled to a mast-mounted mechanical gyro introduced into the pitch control linkage.[292] The root regions of the blades in this innovative design were "matched stiffness” or "soft in-plane,” which meant that the blade chord-wise, or horizon­tal, structural bending stiffness was matched to the flap-wise, or vertical, bending stiffness. Dynamic model tests of this concept were conducted in the Langley 30 by 60 Full-Scale Tunnel and in the Freon atmosphere of the Langley Transonic Dynamics Tunnel. The use of Freon gas facilitated the testing of the 10-foot-diameter rotor model at full-scale Reynolds number and Mach numbers. This work began the establishment of a documented database for hingeless rotor design. These dynamic model tests were part of a cooperative NASA-Army AVLABS program.

To further explore the problems and practical means for realizing the potential of the hingeless rotor concept, Langley Research Center purchased the Lockheed XH-51N, a high-speed research helicopter. The flight investigation focused on the tendency for hingeless rotors to encounter high in-plane blade loads in roll maneuvers, coupling between the response to longitudinal and lateral control input, ride quality, and pilot handling qualities. In general, it was demonstrated with the flight tests and model tests that the hingeless rotor system was different from the conventional hinged systems. Inherently, the hingeless designs pro­duced increased control moments, quicker response to pilot input and superior handling qualities. It turned out that later rotor designs incor­porating elastomeric bearings to replace conventional hinges could pro­vide a practical option to some of the fully hingeless designs.

During the Second World War, Germany held global leadership in high­speed aerodynamics. The most impressive expression of its technical interest and competence in high-speed aircraft and missile design was the V-2 terror weapon, which introduced the age of the long-range rocket. It had a range of over 200 miles at a speed of approximately Mach 5.[547] A longer-range experimental variant tested in 1945, the A-4b, sported swept wings and flew at 2,700 mph, reentering and leveling off in the upper atmosphere for a supersonic glide to its target. In its one semi­successful flight, it completed a launch and reentry, though one wing broke off during its terminal Mach 4+ glide.[548] One appreciates the ambi­tious nature and technical magnitude of the German achievement given that the far wealthier and more technically advantaged United States pursued a vigorous program in piloted rocket planes all through the 1950s without matching the basic performance sought with the A-4b.

Key to the German success was a strong academic-industry part­nership and, particularly, a highly advanced complex of supersonic wind tunnels. The noted tunnel designer Carl Wieselsberger (who died of cancer during the war) introduced a blow-down design that initially operated at Mach 3.3 and later reached Mach 4.4. The latter instrument supported supersonic aerodynamic and dynamic stability studies of var­ious craft, including the A-4b. German researchers had ambitious plans for even more advanced tunnels, including an Alpine complex capable of attaining Mach 10. This tunnel work inspired American emulation after the war and, in particular, stimulated establishment of the Air Force’s Arnold Engineering Development Center at Tullahoma, TN.[549]

At war’s end, America had nothing comparable to the investment Germany had made in high-speed flight, either in rockets or in wind tun­nels and other specialized research facilities. The best American wartime tunnel only reached Mach 2.5. As a stopgap, the Navy seized a German facility, transported it to the United States, and ran it at Mach 5.18, but

HIGH PRESSURE TANK	HIGH VACUUM -PUMP
LOW PRESSURE TANK

DIFFUSER

The layout of the Langley 11-inch hypersonic tunnel advocated by John V. Becker. NASA.

it did this only beginning in 1948.[550] Even so, aerodynamicist John Becker, a young and gifted engineer working at the National Advisory Committee for Aeronautics (NACA) Langley Laboratory, took the initiative in intro­ducing Agency research in hypersonics. He used the V-2 as his rationale. In an August 1945 memo to Langley’s chief of research, written 3 days before the United States atom-bombed Hiroshima, he noted that planned NACA facilities were to reach no higher than Mach 3. With the V-2 having already flown at Mach 5, he declared, this capability was clearly inadequate.

He outlined an alternative design concept for "a supersonic tunnel having a test section four-foot square and a maximum test Mach number of 7.0.”[551] A preliminary estimate indicated a cost of $350,000. This was no mean sum. It was equivalent six decades later to approximately $4.2 mil­lion. Becker sweetened his proposal’s appeal by suggesting that Langley

begin modestly with a small demonstration wind tunnel. It could be built for roughly one-tenth of this sum and would operate in the blow-down mode, passing flow through a 1-foot-square test section. If it proved suc­cessful and useful, a larger tunnel could follow. His reasoned idea received approval from the NACA’s Washington office later in 1945, and out of this emerged the Langley 11-Inch Hypersonic Tunnel. Slightly later, Alfred J. Eggers began designing a hypersonic tunnel at the NACAs West Coast Ames Aeronautical Laboratory, though this tunnel, with a 10-inch by 14-inch test section, used continuous, not blow-down, flow. Langley’s was first. When the 11-inch tunnel first demonstrated successful operation (to Mach 6.9) on November 26, 1947, American aeronautical science entered the hyper­sonic era. This was slightly over a month after Air Force test pilot Capt. Charles E. Yeager first flew faster than sound in the Bell XS-1 rocket plane.[552]

Though ostensibly a simple demonstration model for a larger tun­nel, the 11-inch tunnel itself became an important training and research tool that served to study a wide range of topics, including nozzle devel­opment and hypersonic flow visualization. It made practical contribu­tions to aircraft development as well. Research with the 11-inch tunnel led to a key discovery incorporated on the X-15, namely that a wedge­shaped vertical tail markedly increased directional stability, eliminat­ing the need for very large stabilizing surfaces. So useful was it that it remained in service until 1973, staying active even with a successor, the larger Continuous Flow Hypersonic Tunnel (CFHT), which entered ser­vice in 1962. The CFHT had a 31-inch test section and reached Mach 10 but took a long time to become operational. Even after entering ser­vice, it operated much of the time in a blow-down mode rather than in continuous flow.[553]

Nearly all high-speed airplanes use hydraulic actuators to operate the control surfaces. This provides a significant boost to the pilot’s ability to move a large control surface, which is experiencing very high aero­dynamic loads. The computers and other electronic devices mentioned above merely provided signals to servos, which in turn commanded movement of hydraulic actuators. The hydraulic system provided the real muscle to move the surfaces. When Lockheed Martin’s "Skunk Works” was designing the planned X-33 Research Vehicle (intended to explore one possible design for a single-stage-to-orbit logistical spacecraft), keeping gross lift-off weight (GLOW) as low as possible was a crucial design goal. Because the hydraulic system would have been employed only during boost and entry, the entire hydraulic system would have been dead weight while the vehicle was in the space environment. Thus, control system designers elected to use electro-mechanical actuators to move the control surfaces, eliminating any need for a hydraulic system. Though X-33 was canceled for a variety of other reasons, its provision for electrical actuators clearly pointed toward future design practice.

Following up on this were a series of three flight-test projects dur­ing 1997-1998 as part of the Electrically Powered Actuator Design (EPAD) program sponsored by the Naval Air Warfare Center and Air Force Research Laboratory. Each project tested a different advanced flight control actuator for the left aileron of NASA Dryden’s F/A-18 Systems Research Aircraft (SRA). The first was the "smart actuator” that used fiber optics instead of the normal fly-by-wire system to con-

trol an otherwise conventional hydraulic actuator.[708] The next project flew an electro-hydrostatic actuator that used an electric motor to drive a small hydraulic pump that actuated the left aileron; the actuator was independent of the normal aircraft hydraulic system.[709] The third project used an electro-mechanical actuator (EMA) that used electrical power generated by the F/A-18’s engines to power the left aileron actuator. A fiber-optic controller, self-contained control-surface actuator promises a significant reduction in weight and complexity over conventional actu­ation systems for future advanced air and space vehicles.[710]