Category AERONAUTICS

Another option for thermal protection during entry was the use of exotic, high-temperature materials for the external surface that could re­radiate the heat back into space. This concept was proposed for the X-20 Dyna-Soar program, and the vehicle was well under construc­tion at the time of cancellation.[755] In parallel with the X-20 program, the Air Force Flight Dynamics Laboratory developed a small radia – tive-cooled hot structure vehicle (essentially the first 4 feet of the X-20 Dyna Soar’s nose), called the McDonnell Aerothermodynamic/elastic Structural Systems Environmental Tests (ASSET). The ASSET design used the same materials and thermal protection concepts as the X-20 and first flew in September 1963, 3 months before cancellation of the Dyna-Soar. The fourth ASSET vehicle successfully completed a Mach 18.4 entry from 202,000 feet in 1965. Postflight examination indicated

it survived the entry well, although the operational problems and man­ufacturing methods for these exotic materials were expensive and time­consuming. Since that time, joint NASA-Air Force-Navy-industry devel­opmental programs such as the X-30 National Aero-Space Plane (NASP) effort of the late 1980s to early 1990s have advanced materials and fabri­cation technologies that, in due course, may be applied to future hyper­sonic systems.[756]

Basic principles of structural analysis—static equilibrium, trusses, and beam theory—were known long before computers, or airplanes, existed. Bridges, towers and other buildings, and ships were designed by a combination of experience and some amount of analysis—more so as designs became larger and more ambitious during and after the Industrial Revolution.

With airplanes came much greater emphasis on weight minimiza­tion. Massive overdesign was no longer an acceptable means to achieve structural integrity. More rigorous analysis and structural sizing was required. Simplifications allowed the analysis of primary members under simple loading conditions:

• Slender beams: axial load, shear, bending, torsion.

• Trusses: members carry axial load only, joined to other such members at ends.

• Simple shells: pressure loading.

• Semi-monocoque (skin and stringer) structures: shear flow, etc.

• Superposition of loading conditions.

With these simplifications, primary structural members could be sized appropriately to the expected loads. In the days of wood, wire, and fabric, many aircraft structures could be analyzed as trusses: exter­nally braced biplane wings; fuselage structures consisting of longerons, uprights, and cross braces, with diagonal braces or wires carrying tor­sion; landing gears; and engine mounts. As early as the First World War and in the 1920s, researchers were working to cover every required aspect of the problem: general analysis methods, analysis of wings, horizontal and vertical tails, gust loads, test methods, etc. The National Advisory Committee for Aeronautics (NACA) contributed significantly to the build­ing of this early body of methodology.[787]

Structures with redundancy—multiple structural members capable of sharing one or more loading components—may be desirable for safety, but they posed new problems for analysis. Redundant structures cannot be analyzed by force equilibrium alone. A conservative simplification, often practiced in the early days of aviation, was to analyze the struc­ture with redundant members missing. A more precise solution would require the consideration of displacements and "compatibility” condi­tions: members that are connected to one another must deform in such a manner that they move together at the point of connection. Analysis was feasible but time-consuming. Large-scale solutions to redundant ("statically indeterminate”) structure problems would become practical with the aid of computers. Until then, more simplifications were made, and specific types of solutions—very useful ones—were developed.

While these analysis methods were being developed, there was a lot of airplane building going on without very much analysis at all. In the "golden age of aviation,” many airplanes were built in garages or at small companies that lacked the resources for extensive analysis. "In many cases people who flew the airplanes were the same people who car­ried out the analysis and design. They also owned the company. There was very little of what we now call structural analysis. Engineers were brought in and paid—not to design the aircraft—but to certify that the aircraft met certain safety requirements.”[788]

Through the 1930s, as aircraft structures began to be formed out of aluminum, the semi-monocoque or skin-and-stringer structure became prevalent, and analysis methods were developed to suit. "In the 1930s, ’40s, and ’50s, techniques were being developed to analyze specific struc­tural components, such as wing boxes and shear panels, with combined bending, torsion, and shear loads and with stiffeners on the skins.”[789] A number of exact solutions to the differential equations for stress and strain in a structural member were known, but these generally exist only for very simple geometric shapes and very limited sets of loading conditions and boundary conditions. Exact solutions were of little prac­tical value to the aircraft designer or stress analyst. Instead, "free body diagrams” were used to analyze structures at selected locations, or "sta­tions.” The structure was considered to be cut by a theoretical plane at the station of interest. All loads, applied and inertial, on the portion of the aircraft outboard of the cut had to be borne (reacted) by the struc­ture at the cut.

In principle, this allowed the stress at any point in the structure to be analyzed—given the time to make an arbitrarily large number of these theoretical cuts through the aircraft. In practice, free body dia­grams were used to analyze the structure at key locations—selected fuselage stations, the root, and selected stations of wings and tail sur­faces. Structural members were left constant, or tapered appropriately, according to experience and judgment, between the analyzed sections. For major projects such as airliners or bombers, the analysis would be more thorough, and consequently, major design organizations had rooms full of people whose jobs were to perform the required calculations.

The NACA also utilized this brute-force approach to large calcu­lations, and the people who performed the calculations—overwhelm­ingly women—were called "computers.” Annie J. Easley, who worked at the NASA Lewis (now Glenn) Research Center starting in 1955, recalls:

. . . we were called computers until we started to get the machines, and then we were changed over to either math tech­nicians or mathematicians. . . . The engineers and the scien­tists are working away in their labs and their test cells, and they come up with problems that need mathematical compu­tation. At that time, they would bring that portion to the com­puters, and our equipment then were the huge calculators, where you’d put in some numbers and it would clonk, clonk, clonk out some answers, and you would record them by hand. Could add, subtract, multiply, and divide. That was pretty much what those big machines, those big desktop machines, could do. If we needed to find a logarithm or an exponential, we then pulled out the tables.[790]

After World War II, with jet engines pushing aircraft into ever more demanding flight regimes, the analytical community sought to keep up. The NACA continued to improve the methodologies for calculating loads on various parts of an aircraft, and some of the reports generated during that time are still used by industry practitioners today. NACA Technical Report (TR) 1007, for horizontal tail loads in pitch maneuvers, is a good example, although it does not cover all of the conditions required by recent airworthiness regulations.[791]

For structural analysis, energy methods and matrix methods began to receive more attention. Energy methods work as follows: one first expresses the deflection of a member as a set of assumed shape func­tions, each multiplied by an (initially unknown) coefficient; expresses the total strain energy in terms of these unknown coefficients; and finally, finds the values of the coefficients that minimize the strain energy. If the shape functions, from which the solution is built, satisfy the boundary conditions of the problem, then so does the final solution.

Energy methods were not new. The concept of energy minimization was introduced by Lord Rayleigh in the late 19 th century and extended by Walter Ritz in two papers of 1908 and 1909.[792] Rayleigh and Ritz were par­ticularly concerned with vibrations. Carlo Alberto Castigliano, an Italian engineer, published a dissertation in 1873 that included two important theorems for applying energy principles to forces and static displace­ments in structures.[793] However, in the early works, the shape functions were continuous over the domain of interest. The idea of breaking up (discretizing) a complex structure into many simple elements for numer­ical solution would lead to the concept of finite elements, but for this to be useful, computing technology needed to mature.

We now turn to an area of activity that provides, for aviation, the ulti­mate proof of design techniques and predictive capabilities: flight-test­ing. While there are many fascinating projects that could be discussed, we will consider only five that had particular relevance to the subject at hand, either because they collected data that were specifically intended to provide validation of computational predictions of structural behav­ior, or because they demonstrated unique structural design approaches.

Two of these are the YF-12 Thermal Loads project and the Rotor Aerodynamic Limits survey, both of which collected data for validat­ing and improving predictive methods. The remaining three are the Highly Maneuverable Aircraft Technology (HiMAT) digital fly-by-wire (DFBW) enhanced agility composite-structured canard demonstrator, the AD-1 oblique wing demonstrator, and the Grumman X-29 forward- swept wing (FSW) research aircraft. These three projects exercised, in progressively more challenging ways, the concept of aeroelastic tailor­ing: that is, predicting airframe flexibility and having enough confidence in those predictions to design an airplane that takes advantage of elas­tic deformation, rather than just trying to minimize it. In all of these, NASA-rooted computational structural prediction proved of great, and even occasionally, critical, significance.

The investigation of aircraft structural mechanics or, indeed, of almost any discipline, can be considered to include the following activ­ities: investigation by basic theory, computational analysis or simula­tion, laboratory test, and flight test (or, more generally, any test of the final product in its actual operating environment). Many arguments have been had over which is the most valuable. This author is of the opinion—based on his experience in the practice of engineering, on a certain amount of historical research, and on the teaching and example of mentors and peers—that theory, computation, laboratory test, and flight test all constitute imperfect but complementary views of reality. Thus, until someone comes up with a way to know the exact state of stress and deflection in every part of a vehicle under actual operating conditions, we must form our understanding of reality as a composite image, using what information we can gain from each available source:

• Flight test, obviously, is the best representation we have of an aircraft in actual operational conditions. However, our ability to interrogate the system is most severely compromised in this activity. Many data parameters are not available unless special instrumentation is installed, if at all, and this is the most difficult environment in which to obtain stable, high-quality data.

• Laboratory test offers better visibility into the opera­tion of specific parts of the system and better control of experimental parameters, at the price of some separa­tion from true operational conditions.

• Computation offers even greater opportunity to inter­rogate the value of any data parameter at any time(s) and to simulate conditions that might be impossible, difficult, or dangerous to test. Computation also elimi­nates all physical complications of running the experi­ment and all physical sources of noise and uncertainty.

But in stepping out of the physical world and into the analytical world, the researcher also becomes subject to the limited fidelity of his computational method: what effects are and are not included in the computation and how well the computation represents physical reality.

• Theory is sometimes the best source of insight and of understanding what parameters might be changed to obtain some desired effect, but it does not provide the detailed quantitative data necessary to implement the solution.

In this light, the following flight programs are discussed. Much more could be said about each of them. The present discussion is necessarily confined to their significance to the development or validation of loads and structural computation methods.

Test facilities have an important role in verifying and improving anal­ysis methods. A few test facilities that had a lot to do with the devel­opment and validation of structural analysis methods are described below. In addition to those described, other "landmark” test facilities include large-scale launch vehicle structural test facilities at Johnson and Marshall Space Centers, and the crash dynamics test facility at Langley Research Center.

Structural Dynamics Laboratory (Ames Research Center, 1965)

During the 1960s, Ames and Langley collaborated on some of the struc­tural dynamics and buffet problems of spacecraft during ascent. (This collaboration occurred through some of the same meetings at NASA Headquarters that led to the development of NASTRAN.) To help assess the structural dynamic characteristics of boosters, and to build confi­dence in predictive methods, a large structural dynamics test facility was built at Ames (completed in 1965). This facility was large enough to hold a full-size Atlas or Titan II, had provisions for exciting the struc­tural modes of the test article, and could be evacuated to test the struc­tural damping characteristics in zero or reduced ambient air density.[1005] The facility was also used for research on buffet during reentry and land­ing impacts.[1006] Much of the structural dynamics research at Ames was discontinued or relocated during the early 1970s. The laboratory is long since deactivated, but the large, pentagonal tower still stands, housing a machine shop and a wind tunnel that can simulate Mars’s atmosphere by evacuating the chamber and then filling to low pressure with CO2.[1007]

Thermal Loads Laboratory (Dryden Flight Research Center, 1960s)

A 1973 accounting of NASA research facilities listed only one major ground laboratory at Dryden: the High Temperature Loads Calibration Laboratory.[1008] High supersonic and hypersonic flight research created a need (1) to test airframes on the ground under simultaneous thermal and structural loading conditions and (2) to calibrate loads instrumen-

tation at elevated temperatures, so that the data obtained in flight could be reliably interpreted. These needs " . . . led to the construction of a laboratory for calibrating strain-gage installations to measure loads in an elevated temperature environment. The problems involved in mea­suring loads with strain gages. . . require the capability to heat and load aircraft under simulated flight conditions. . . . The laboratory has the capability of testing structural components and complete vehicles under the combined effects of loads and temperatures, and calibrating and evaluating flight loads instrumentation under [thermal] conditions expected in flight.”[1009]

The laboratory is housed in a hangarlike building with attached shop, offices, and control room. Capabilities included:

• Hangar-door opening 40 feet high by 136 feet wide.

• Unobstructed test area 150 by 120 by 40 feet allowed the testing of aircraft up to and including, for example, a YF-12 or SR-71.

• Ten megawatts of electrical heating power via quartz lamps and reflectors.

• Temperatures up to 3,000 °F.

• Hydraulic power of 4.5 gallons/minute at 3,000 pounds per square inch (psi) to apply loads.

• Fourteen channels closed-loop load or position control of up to 34 separate actuators.

• Sensors including strain gages, thermocouples, load cells, and position transducers.

Slots in the floor provided flexible locations for tiedown points, as well as routing for hydraulic and electrical power, instrumentation wir­ing, compressed air, or water (presumably for cooling). Closed-loop ana­log control of both mechanical load and heating was provided, to any desired preprogrammed time history.

The facility was used in the YF-12 thermal loads project (discussed elsewhere in this paper), in Space Shuttle structural verification at high

temperatures, and for a variety of other studies.[1010] The loads laboratory made contributions to the validation of computational methods by pro­viding the opportunity to compare computational predictions with test data obtained under known, controlled, thermal and structural load­ing conditions, applied together or independently as required. At time of this writing, the facility is still in use.[1011]

By the 1950s, fully boosted flight controls were common, and the potential benefits of fly-by-wire were becoming increasingly apparent. Beginning during the Second World War and continuing postwar, fly-by­wire and power-by-wire flight control systems had been fielded in var­ious target drones and early guided missiles.[1114] However, most aircraft designers were reluctant to completely abandon mechanical linkages to
flight control surfaces in piloted aircraft, an attitude that would undergo an evolutionary change over the next two decades as a result of a broad range of NACA-NASA, Air Force, and foreign research efforts.

Beginning in 1952, the NACA Langley Aeronautical Laboratory began an effort oriented to exploring various aspects of fly-by-wire, including the use of a side stick controller.[1115] By 1954, flight-testing began with what was perhaps the first jet-powered fly-by-wire research aircraft, a modified former U. S. Navy Grumman F9F-2 Panther carrier-based jet fighter used as an NACA research aircraft. The primary objective of the NACA effort was to evaluate various automatic flight control systems, including those based on rate and normal acceleration feedback. Secondary objectives were to evaluate use of fly-by-wire with a side stick controller for pilot inputs. The existing F9F-2 hydraulic flight control system, with its mechan­ical linkages, was retained with the NACA designing an auxiliary flight control system based on a fly-by-wire analog concept. A small, 4-inch-tall side stick controller was mounted at the end of the right ejection seat arm­rest. The controller was pivoted at the bottom and was used for both lat­eral (roll) and longitudinal (pitch) control. Only 4 pounds of force were required for full stick deflection. The control friction normally present in a hydromechanical system was completely eliminated by the electrically powered system. Additionally, the aircraft’s fuel system was modified to enable fuel to be pumped aft to destabilize the aircraft by moving the cen­ter of gravity rearward. Another modification was the addition of a steel container mounted on the lower aft fuselage. This carried 250 pounds of lead shot to further destabilize the aircraft. In an emergency, the shot could be rapidly jettisoned to restabilize the aircraft. Fourteen pilots flew the modified F9F-2, including NACA test pilots William Alford[1116] and Donald

L. Mallick.[1117] [1118] Using only the side stick controller, the pilots conducted
takeoffs, stall approaches, acrobatics, and rapid precision maneuvers that included air-to-air target tracking, ground strafing runs, and pre­cision approaches and landings. The test pilots quickly became used to flying with the side stick and found it comfortable and natural to use.18

In mid-1956, after interviewing aircraft flight control experts from the Air Force Wright Air Development Center’s Flight Control Laboratory, Aviation Week magazine concluded:

The time may not be far away when the complex mechani­cal linkage between the pilot’s control stick and the airplane’s control surface (or booster valve system) is replaced with an electrical servo system. It has long been recognized that this"fly-by-wire” approach offered attractive possibilities for reducing weight and complexity. However, airplane designers and pilots have been reluctant to entrust such a vital function to electronics whose reliability record leaves much to be desired.[1119]

Even as the Aviation Week article was published, several noteworthy aircraft were under development that would incorporate various fly-by­wire approaches in their flight control systems. In 1956, the British Avro Vulcan B.2 bomber flew with a partial fly-by-wire system that operated in conjunction with hydraulically boosted, mechanically activated flight controls. The supersonic North American A-5 Vigilante Navy carrier – based attack bomber flew in 1958 with a pseudo-fly-by-wire flight control system. The Vigilante served the fleet for many years, but its highly com­plex design proved very difficult to maintain and operate in an aircraft carrier environment. By the mid-1960s, the General Dynamics F-111 was flying with triple-redundant, large-authority stability and command aug­mentation systems and fly-by-wire-controlled wing-mounted spoilers.[1120]

On the basic research side, the delta winged British Short S. C.1, first flown in 1957, was a very small, single-seat Vertical Take-Off and Landing
(VTOL) aircraft. It incorporated a triply redundant fly-by-wire flight con­trol system with a mechanical backup capability. The outputs from the three independent fly-by-wire channels were compared, and a failure in a single channel was overridden by the other two. A single channel failure was relayed to the pilot as a warning, enabling him to switch to the direct (mechanical) control system. The S. C. 1 had three flight control modes, as described below, with the first two only being selectable prior to takeoff.[1121]

• Full fly-by-wire mode with aerodynamic surfaces and noz­zles controlled electrically via three independent servo motors with triplex fail-safe operation in conjunction with three analog autostabilizer control systems.

• A hybrid mode in which the reaction nozzles were servo/ autostabilizer (fly-by-wire) controlled and the aerodynamic surfaces were linked directly to the pilot’s manual controls.

• A direct mode in which all controls were mechanically linked to the pilot control stick.

The S. C. 1 weighed about 8,000 pounds and was powered by four ver­tically mounted Rolls-Royce RB.108 lift engines, providing a total ver­tical thrust of 8,600 pounds. One RB.108 engine mounted horizontally in the rear fuselage provided thrust for forward flight. The lift engines were mounted vertically in side-by-side pairs in a central engine bay and could be swiveled to produce vectored thrust (up to 23 degrees for­ward for acceleration or -12 degrees for deceleration). Variable thrust nose, tail, and wingtip jet nozzles (powered by bleed air from the four lift engines) provided pitch, roll, and yaw control in hover and at low speeds during which the conventional aerodynamic controls were inef­fective. The S. C.1 made its first flight (a conventional takeoff and land­ing) on April 2, 1957. It demonstrated tethered vertical flight on May 26, 1958, and free vertical flight on October 25, 1958. The first transi­tion from vertical flight to conventional flight was made April 6, 1960.[1122]

During 10 years of flight-testing, the two S. C.1 aircraft made hun­dreds of flights and were flown by British, French, and NASA test pilots. A Royal Aircraft Establishment (RAE) report summarizing flight-test experience with the S. C. 1 noted: "Of the visiting pilots, those from NASA [Langley’s John P. "Jack” Reeder and Fred Drinkwater from Ames] flew the aircraft 6 or 7 times each. They were pilots of very wide experience, including flight in other VTOL aircraft and variable stability helicopters, which was of obvious assistance to them in assessing the S. C.1.”[1123] On October 2, 1963, while hovering at an altitude of 30 feet, a gyro input malfunction in the flight control system produced uncontrollable pitch and roll oscillations that caused the second S. C. 1 test aircraft (XG 905) to roll inverted and crash, killing Shorts test pilot J. R. Green. The air­craft was then rebuilt for additional flight-testing. The first S. C. 1 (XG 900) was used for VTOL research until 1971 and is now part of the Science Museum aircraft collection at South Kensington, London. The second S. C.1 (XG 905) is in the Flight Experience exhibit at the Ulster Folk and Transport Museum in Northern Ireland, near where the air­craft was originally built by Short Brothers.

The Canadian Avro CF-105 Arrow supersonic interceptor flew for the first time in 1958. Revolutionary in many ways, it featured a dual channel, three-axis fly-by-wire flight control system designed without any mechanical backup flight control capability. In the CF-105, the pilot’s control inputs were detected by pressure-sensitive transducers mounted in the pilot’s control column. Electrical signals were sent from the transducers to an electronic control servo that operated the valves in the hydraulic system to move the various flight control surfaces. The CF-105 also incorporated artificial feel and stability augmentation sys­tems.[1124] In a highly controversial decision, the Canadian government can­celed the Arrow program in 1959 after five aircraft had been built and flown. Although only about 50 flight test hours had been accumulated, the Arrow had reached Mach 2.0 at an altitude of 50,000 feet. During its development, NACA Langley Aeronautical Laboratory assisted the CF-105 design team in a number of areas, including aerodynamics, performance, stability, and control. After the program was terminated,

many Avro Canada engineers accepted jobs with NASA and British or American aircraft companies.[1125] Although it never entered production and details of its pioneering flight control system design were reportedly lit­tle known at the time, the CF-105 presaged later fly-by-wire applications.

NACA test data derived from the F9F-2 fly-by-wire experiment were used in development of the side stick controllers in the North American X-15 rocket research plane, with its adaptive flight control system.[1126] First flown in 1959, the X-15 eventually achieved a speed of Mach 6.7 and reached a peak altitude of 354,200 feet. One of the two side stick con­trollers in the X-15 cockpit (on the left console) operated the reaction thruster control system, critical to maintaining proper attitude control at high Mach numbers and extreme altitudes during descent back into the higher-density lower atmosphere. The other controller (on the right cockpit console) operated the conventional aerodynamic flight control surfaces. A CALSPAN NT-33 variable stability test aircraft equipped with a side stick controller and an NACA-operated North American F-107A (ex-USAF serial No. 55-5120), modified by NACA engineers with a side stick flight control system, were flown by X-15 test pilots during 1958— 1959 to gain side stick control experience prior to flying the X-15.[1127]

Interestingly, the British VC10 jet transport, which first flew in 1962, has a quad channel flight control system that transmits electrical signals directly from the pilot’s flight controls or the aircraft’s autopilot via elec­trical wiring to self-contained electrohydraulic Powered Flight Control Units (PFCUs) in the wings and tail of the aircraft, adjacent to the flight control surfaces. Each VC10 PFCU consists of an individual small self – contained hydraulic system with an electrical pump and small reservoir. The PFCUs move the control surfaces based on electrical signals pro­vided to the servo valves that are electrically connected to the cockpit flying controls.[1128] There are no mechanical linkages or hydraulic lines between the pilot and the PFCUs. The PFCUs drive the primary flight
control surfaces that consist of split rudders, ailerons, and elevators on separate electrical circuits. Thus, the VC10 has many of the attributes of fly-by-wire and power-by-wire flight control systems. It also features a backup capability that allows it to be flown using the hydraulically boosted variable incidence tail plane and differential spoilers that are operated via conventional mechanical linkages and separate hydraulic systems.[1129] The VC10K air refueling tanker was still in Royal Air Force (RAF) service as of 2009, and the latest Airbus airliner, the A380, uses the PFCU concept in its fly-by-wire flight control system.

The Anglo-French Concorde supersonic transport first flew in 1969 and was capable of transatlantic sustained supercruise speeds of Mach 2.0 at cruising altitudes well above 50,000 feet. In support of the Concorde development effort, a two-seat Avro 707C delta winged flight research aircraft was modified as a fly-by-wire technology testbed with a side stick controller. It flew 200 hours on fly-by-wire flight trails at the U. K. at Farnborough until September 1966.[1130] Concorde had a dual channel analog fly-by-wire flight control system with a backup mechanical capa­bility. The mechanical system served in a follower role unless problems developed with the fly-by-wire control elements of the system, in which case it was automatically connected. Pilot movements of the cockpit con­trols operated signal transducers that generated commands to the flight control system. These commands were processed by an analog electri­cal controller that included the aircraft autopilot. Mechanically operated servo valves were replaced by electrically controlled ones. Much as with the CF-105, artificial feel forces were electrically provided to the Concorde pilots based on information generated by the electronic controller.[1131]

Phase I flight-testing was conducted by the AFTI/F-16 Joint Test Force from the NASA Dryden Flight Research Facility at Edwards AFB, CA, from July 10, 1982, through July 30, 1983. During this phase, five test pilots from NASA, the Air Force, and the U. S. Navy flew the aircraft. Initial flights checked out the aircraft’s stability and control systems. Handling qualities were assessed in air-to-air and air-to-ground scenarios, as well
as in-formation flight and during approach and landing. The Voice Command System allowed the pilot to change switch positions, display formats, and modes simply by saying the correct word. Initial tests were of the system’s ability to recognize words, with later testing conducted under increasing levels of noise, vibrations, and g-forces. Five pilots flew a total of 87 test sorties with the Voice Command System, with a gen­eral success rate approaching 90 percent. A prototype helmet-mounted sight was also evaluated. On July 30, 1983, the AFTI/F-16 aircraft was flown back to the General Dynamics facility at Fort Worth, TX, for mod­ification for Phase II. During the Phase I test effort, 118 flight-test sor­ties were flown, totaling about 177 flight hours. In addition to evaluating the DFCS, the potential operational utility of task-tailored flight modes (that included decoupling of aircraft attitude and flight path) was also assessed. During these unconventional maneuvers, the AFTI/F-16 dem­onstrated that it could alter its nose position without changing flight path and change its flight path without changing aircraft attitude. The air­craft also performed coordinated horizontal turns without banking or sideslip.[1183] NASA test pilot Bill Dana recounted: "In Phase I we evaluated non-classic flight control modes. By deflecting the elevators and flaps in various relationships, it was possible to translate the aircraft vertically without changing pitch attitude or to pitch-point the airplane without changing your altitude. You could also translate laterally without using bank and yaw-point without translating the aircraft, by using rudder and canard inputs programmed together in the flight control computer.”[1184]

The Highly Integrated Digital Electronic Control (HIDEC) evolved from the earlier DEEC research effort. Major elements of the HIDEC were
a Digital Electronic Flight Control System (DEFCS), engine-mounted DEECs, an onboard general-purpose computer, and an integrated archi­tecture that provided connectivity between components. The HIDEC F-15A (USAF serial No. 71-0287) was modified to incorporate DEEC – equipped F100 engine model derivative (EMD) engines. A dual chan­nel Digital Electronic Flight Control System augmented the standard hydromechanical flight control system in the F-15A and replaced its ana­log control augmentation system. The DEFCS was linked to the aircraft data buses to tie together all other electronic systems, including the air­craft’s variable geometry engine inlet control system.[1261] Over a span of about 15 years, the HIDEC F-15 would be used to develop several modes of integrated propulsion and flight control systems. These integrated modes were Adaptive Engine Control System, Performance Seeking Control, Self-Repairing Flight Control System, and the Propulsion-Only Flight Control System. They are discussed separately in the following sections.[1262]

The third engine-related effort to design a more fuel-efficient powerplant during this era did not focus on another idea for a turbojet configura­tion. Instead, engineers chose to study the feasibility of reintroducing a jet-powered propeller to commercial airliners. An initial run of the numbers suggested that such an advanced turboprop promised the larg­est reduction in fuel cost, perhaps by as much as 20 to 30 percent over turbofan engines powering aircraft with a similar performance. This compared with the goal of a 5-percent increase in fuel efficiency for the Engine Component Improvement program and a 10- to 15-percent increase in fuel efficiency for the E Cubed program.[1316]

But the implementation of an advanced turboprop was one of NASA’s more challenging projects, both in terms of its engineering and in secur­ing public acceptance. For years, the flying public had been conditioned to see the fanjet engine as the epitome of aeronautical advancement. Now they had to be "retrained” to accept the notion that a turbopropel­ler engine could be every bit as advanced, indeed, even more advanced, than the conventional fanjet engine. The idea was to have a jet engine
firing as usual with air being compressed and ignited with fuel and the exhaust expelled after first passing through a turbine. But instead of the turbine spinning a shaft that turned a fan at the front of the engine, the turbines would be spinning a shaft, which fed into a gearbox that turned another shaft that spun a series of unusually shaped propeller blades exterior to the engine casing.[1317]

Begun in 1976, the project soon grew into one of the larger NASA aeronautics endeavors in the history of the Agency to that point, eventu­ally involving 4 NASA Field Centers, 15 university grants, and more than 40 industrial contracts.[1318]

Early on in the program, it was recognized that the major areas of concern were going to be the efficiency of the propeller at cruise speeds, noise both on the ground and within the passenger cabin, the effect of the engine on the aerodynamics of the aircraft, and maintenance costs. Meeting those challenges were helped once again by the computer-aided, three-dimensional design programs created by the Lewis Research Center. An original look for an aircraft propeller was devised that changed the blade’s sweep, twist, and thickness, giving the propellers the look of a series of scimitar-shaped swords sticking out of the jet engine. After much development and testing, the NASA-led team eventually found a solution to the design challenge and came up with a propeller shape and engine configuration that was promising in terms of meeting the fuel-efficiency goals and reduced noise by as much as 65 decibels.[1319]

In fact, by 1987, the new design was awarded a patent, and the NASA-industry group was awarded the coveted Collier Trophy for creat­ing a new fuel-efficient turboprop propulsion system. Unfortunately, two unexpected variables came into play that stymied efforts to put the design into production.[1320]

The first had to do with the public’s resistance to the idea of flying in an airliner powered by propellers—even though the blades were still

being turned by a jet engine. It didn’t matter that a standard turbofan jet also derived most of its thrust from a series of blades—which did, in fact, look more like a fan than a series of propellers. Surveys showed passengers had safety concerns about an exposed blade letting go and sending shrapnel into the cabin, right where they were sitting. Many passengers also believed an airliner equipped with an advanced turbo­prop was not as modern or reliable as pure turbojet engine. Jets were in; propellers were old fashioned. The second thing that happened was that world fuel prices dropped to the lower levels that preceded the oil embargo and the very rationale for developing the new turboprop in the first place. While fuel-efficient jet engines were still needed, the "extra mile” in fuel efficiency the advanced turboprop provided was no lon­ger required. As a result, NASA and its partners shelved the technology and waited to use the archived files another day.[1321]

The story of the Advanced Turboprop project had one more twist to it. While NASA and its team of contractor engineers were working on their new turboprop design, engineers at GE were quietly working on their own design, initially without NASA’s knowledge. NASA’s engine was distinguished by the fact that it had one row of blades, while GE’s ver­sion featured two rows of counter-rotating blades. GE’s design, which became known as the Unducted Fan (UDF), was unveiled in 1983 and demonstrated at the 1985 Paris Air Show. A summary of the UDF’s tech­nical features is described in a GE-produced report about the program:

The engine system consists of a modified F404 gas generator engine and counterrotating propulsor system, mechanically decoupled, and aerodynamically integrated through a mixing frame structure. Utilization of the existing F404 engine min­imized engine hardware, cost, and timing requirements and provided an engine within the desired thrust class. The power turbine provides direct conversion of the gas generator horse­power into propulsive thrust without the requirement for a gearbox and associated hardware. Counterrotation utilizes the full propulsive efficiency by recovering the exit swirl between blade stages and converting it into thrust.[1322]

Although shelved during the late 1980s, the Alternate Turboprop and UDF technology and concept is being explored again as part of programs such as the Ultra-High Bypass Turbofan and Pratt & Whitney’s Geared Turbofan. Neither engine is routinely flying yet on commercial airlin­ers. But both concepts promise further reductions in noise, increases in fuel efficiency, and lower operating costs for the airline—goals the aero­space community is constantly working to improve upon.

Several concepts are under study for an Ultra-High Bypass Turbofan, including a modernized version of the Advanced Turboprop that takes advantage of lessons learned from GE’s UDF effort. NASA has teamed with GE to start testing an open-rotor engine. For the NASA tests at Glenn Research Center, GE will run two rows of counter-rotating fan blades, with 12 blades in the front row and 10 blades in the back row. The composite fan blades are one-fifth subscale in size. Tests in
a low-speed wind tunnel will simulate low-altitude aircraft speeds for acoustic evaluation, while tests in a high-speed wind tunnel will simulate high-altitude cruise conditions in order to evaluate blade efficiency and performance.[1323]

"The tests mark a new journey for GE and NASA in the world of open rotor technology. These tests will help to tell us how confident we are in meeting the technical challenges of an open-rotor architecture. It’s a journey driven by a need to sharply reduce fuel consumption in future aircraft,” David Joyce, president of GE Aviation, said in a statement.[1324]

In an Ultra-High Bypass Turbofan, the amount of air going through the engine casing but not through the core compressor and combustion chamber is at least 10 times greater than the air going through the core. Such engines promise to be quieter, but there can be tradeoffs. For exam­ple, an Ultra-High Bypass Engine might have to operate at a reduced thrust or have its fan spin slower. While the engine would meet all the goals, it would fly slower, thus making passengers endure longer trips.

In the case of Pratt & Whitney’s Geared Turbofan engine, the idea is to have an Ultra-High Bypass Ratio engine, yet spin the fan slower (to reduce noise and improve engine efficiency) than the core compressor blades and turbines, all of which traditionally spin at the same speed, as they are connected to the same central shaft. Pratt & Whitney designed a gearbox into the engine to allow for the central shaft to turn at one speed yet turn a second shaft connected to the fan at another speed.[1325]

Alan H. Epstein, a Pratt & Whitney vice president, testifying before the House Subcommittee on Transportation and Infrastructure in 2007, explained the potential benefits the company’s Geared Turbofan might bring to the aviation industry:

The Geared Turbofan engine promises a new level of very low noise while offering the airlines superior economics and envi­ronmental performance. For aircraft of 70 to 150 passenger size, the Geared Turbofan engine reduces the fuel burned,
and thus the CO2 produced, by more than 12% compared to today’s aircraft, while reducing cumulative noise levels about 20dB below the current Stage 4 regulations. This noise level, which is about half the level of today’s engines, is the equiva­lent difference between standing near a garbage disposal run­ning and listening to the sound of my voice right now.[1326]

Pratt & Whitney’s PW1000G engine incorporating a geared turbo­fan is selected to be used on the Bombardier CSeries and Mitsubishi Regional Jet airliners beginning in 2013. The engine was first flight-tested in 2008, using an Airbus A340-600 airliner out of Toulouse, France.[1327]

Bruce I. Larrimer

Consistent with its responsibilities to exploit aeronautics technology for the benefit of the American people, NASA has pioneered the develop­ment and application of alternative energy sources. Its work is argu­ably most evident in wind energy and solar power for high-altitude remotely piloted vehicles. Here, NASA’s work in aerodynamics, solar power, lightweight structural design, and electronic flight controls has proven crucial to the evolution of novel aerospace craft.

HIS CASE STUDY REVIEWS two separate National Aeronautics and Space Administration (NASA) programs that each involved research and development (R&D) in the use of alternative energy. The first part of the case study covers NASA’s participation in the Federal Wind Energy Program from 1974 through 1988. NASA’s work in the wind energy area included design and fabrication of large horizontal-axis wind turbine (HAWT) generators, and the conduct of supporting research and technology projects. The second part of the case study reviews NASA’s development and testing of high-altitude, long-endurance solar – powered unmanned aerial vehicles (UAVs). This program, which ran from 1994 through 2003, was part of the Agency’s Environmental Research and Aircraft Sensor Technology Program.

NASA had first acquired solar cells from Spectralab but chose cells from SunPower Corporation of Sunnyvale, CA, for the ERAST UAVs. These photovoltaic cells converted sunlight directly into electricity and were lighter and more efficient than other commercially available solar cells at that time. Indeed, after NASA flew Helios, SunPower was selected to fur­nish high-efficiency solar concentrator cells for a NASA Dryden ground solar cell test installation, spring-boarding, as John Del Frate recalled subsequently, "from the technology developed on the PF+ and Helios solar cells.”[1546] The Dryden solar cell configuration consisted of two fixed – angle solar arrays and one sun-tracking array that together generated up to 5 kilowatts of direct current. Field-testing at the Dryden site helped SunPower lower production costs of its solar cells and identify uses and performance of its cells that enabled the company to develop large-scale commercial applications, resulting in the mass-produced SunPower A-300 series solar cells.[1547] SunPower’s solar cells were selected for use on the Pathfinders, Centurion, and Helios Prototype UAVs because of their high – efficiency power recovery (more than 50-percent higher than other com­mercially available cells) and because of their light weight. The solar cells designed for the last generation of ERAST UAVs could convert about 19 percent of the solar energy received into 35 kilowatts of electrical current at high noon on a summer day. The solar cells on the ERAST vehicles were bifacial, meaning that they could absorb sunlight on both sides of the cells, thus enabling the UAV s to catch sunrays reflected upward when flying above cloud covers, and were specially developed for use on the aircraft.

While solar cell technology satisfied the propulsion problem during daylight hours, a critical problem relating to long-endurance backup sys­tems remained to be solved for flying during periods of darkness. Without solving this problem, solar UAV flight would be limited to approximately 14 hours in the summer (much less, of course, in the dark of winter), plus whatever additional time could be provided by the limited (up to 5 hours for the Pathfinder) backup batteries. Although significant improvements had been made, batteries failed to satisfy both the weight limitation and long duration power generation requirements for the solar-powered UAVs.

As an alternative to batteries, the ERAST alliance tested a number of different fuel cells and fuel cell power systems. An initial problem to overcome was how to develop lightweight fuel cells because only 440 pounds of Helios’s takeoff weight of 1,600 pounds were originally planned to be allocated to a backup fuel cell power system. Helios required approximately 120 kilowatthours of energy to power the craft for up to 12 hours of flight during darkness, and, fortunately, the state of fuel cell technology had advanced far enough to permit attaining this; ear­lier efforts back to the early 1980s had been frustrated because fuel cell technology was not sufficiently developed at that time. The NASA – industry team later determined, as part of the ERAST program, that a hydrogen-oxygen regenerative fuel cell system (RFCS or regen system) was the hoped for solution to the problem, and substantial resources were committed to the project.

RFCSs are closed systems whereby some of the electrical power pro­duced by the UAV’s solar array during daylight hours is sent to an electro­lyzer that takes onboard water and disassociates the water into hydrogen gas and oxygen gas, both of which are stored in tanks aboard the vehicle. During periods of darkness, the stored gases are recombined in the fuel cell, which results in the production of electrical power and water. The power is used to maintain systems and altitude. The water is then stored for reuse the following day. This cycle theoretically would repeat on a 24-hour basis for an indefinite time period. NASA and AeroVironment also considered, but did not use, a reversible regen system that instead of having an electrolyzer and a fuel cell used only a reversible fuel cell to do the work of both components.[1548]

As originally planned, Helios was to carry two separate regen fuel cell systems contained in two of four landing gear pods. This not only disbursed the weight over the flying wing, but also was in keeping with the plan for redundant systems. If one of the two fuel cells failed, Helios could still stay aloft for several days, albeit at a lower altitude. Contracts to make the fuel cell and electrolyzer were given to two companies—Giner of Waltham, MA, and Lynntech, Inc., of College Station, TX. Each of the two systems was planned to weigh 200 pounds, including 27 pounds for the fuel cell, 18 pounds for the electrolyzer, 40 pounds for oxygen and hydrogen tanks, and 45 pounds for water. The remaining 70 pounds con­sisted of plumbing, controls, and ancillary equipment.[1549]

While the NASA-AeroVironment team made a substantial invest­ment in the RFCS and successfully demonstrated a nearly closed system in ground tests, it decided that the system was not yet ready to satisfy the planned flight schedule. Because of these technical difficulties and time and budget deadlines, NASA and AeroVironment agreed in 2001 to switch to a consumable hydrogen-air primary fuel cell system for the Helios Prototype’s long-endurance ERAST mission. The fuel cells were already in development for the automotive industry. The hydrogen-air fuel cell system required Helios to carry its own supply of hydrogen. In periods of darkness, power for the UAV would be produced by combining gaseous hydrogen and air from the atmosphere in a fuel cell. Because of the low air density at high altitudes, a compressor needed to be added to the sys­tem. This system, however, would operate only until the hydrogen fuel was consumed, but the team thought that the system could still provide multiple days of operation and that an advanced version might be able to stay aloft for up to 14 days. The installation plan was likewise changed. The fuel cell was now placed in one pod with the hydrogen tanks attached to the lower surface of the wing near each wingtip. This modification, of course, dramatically changed Helios’s structural loadings, transforming it from a span-loaded flying wing to a point-loaded vehicle.[1550]