Category AERONAUTICS

John D. Anderson, Jr.

The expanding capabilities of the computer readily led to its increasing application to the aerospace sciences. NACA-NASA researchers were quick to realize how the computer could supplement traditional test meth­odologies, such as the wind tunnel and structural test rig. Out of this came a series of studies leading to the evolution of computer codes used to undertake computational fluid dynamics and structural predictive studies. Those codes, refined over the last quarter century and available to the public, are embodied in many current aircraft and spacecraft systems.

HE VISITOR TO THE SMITHSONIAN INSTITUTION’S National Air and Space Museum (NASM) in Washington, DC, who takes the east escalator to the second floor, turns left into the Beyond the Limits exhibit gallery, and then turns left again into the gallery’s main bay is suddenly confronted by three long equations with a bunch of squiggly symbols neatly painted on the wall. These are the Navier-Stokes equa­tions, and the NASM (to this author’s knowledge) is the world’s only museum displaying them so prominently. These are not some introduc­tory equations drawn for a first course in algebra, with simple symbols like a + b = c. Rather, these are "partial derivatives” strung together from the depths of university-level differential calculus. What are the Navier – Stokes equations, why are they in a gallery devoted to the history of the computer as applied to flight vehicles, and what do they have to do with the National Aeronautics and Space Administration (which, by the way, dominates the artifacts and technical content exhibited in this gallery)?

The answers to all these questions have to do with computational fluid dynamics (CFD) and the pivotal role played by the National Aeronautics and Space Administration (NASA) in the development of CFD over the past 50 years. The role played by CFD in the study and understanding of fluid dynamics in general and in aerospace engineering

in particular has grown from a fledgling research activity in the 1960s to a powerful "third” dimension in the profession, an equal partner with pure experiment and pure theory. Today it is used to help design air­planes, study the aerodynamics of automobiles, enhance wind tunnel testing, develop global weather models, and predict the tracts of hurri­canes, to name just a few. New jet engines are developed with an exten­sive use of CFD to model flows and combustion processes, and even the flow field in the reciprocating engine of the average family automobile is laid bare for engineers to examine and study using the techniques of CFD.

The history of the development of computational fluid dynamics is an exciting and provocative story. In the whole spectrum of the his­tory of technology, CFD is still very young, but its importance today and in the future is of the first magnitude. This essay offers a capsule history of the development of theoretical fluid dynamics, tracing how the Navier-Stokes equations came about, discussing just what they are and what they mean, and examining their importance and what they have to do with the evolution of computational fluid dynamics. It then discusses what CFD means to NASA—and what NASA means to CFD. Of course, many other players have been active in CFD, in universities, other Government laboratories, and in industry, and some of their work will be noted here. But NASA has been the major engine that powered the rise of CFD for the solution of what were otherwise unsolvable prob­lems in the fields of fluid dynamics and aerodynamics.

The project to develop a general-purpose finite element structural analysis system was conceived in the midst of this rapid expansion of finite element research in the 1960s. The development, and subsequent management, enhancement, and distribution, of the NASA Structural Analysis System, or NASTRAN, unquestionably constitutes NASA’s great­est single contribution to computerized structural analysis—and argu­ably the single most influential contribution to the field from any source. NASTRAN is the workhorse of structural analysis: there may be more advanced programs in use for certain applications or in certain proprie­tary or research environments, but NASTRAN is the most capable general – purpose, generally available, program for structural analysis in existence today, even more than 40 years after it was introduced.

The AD-1 was a small and inexpensive demonstrator aircraft intended to investigate some of the issues of an oblique wing. It flew between 1979 and 1982. It had a maximum takeoff weight of 2,100 pounds and a max­imum speed of 175 knots. It is an interesting case because (1) NASA had an unusually large role in its design and integration—it was essentially a NASA aircraft—and (2) because it provides a neat illustration of the prosecution of a particular objective through design, analysis, wind tun­nel test, flight test, and planned follow-on development.[953]

The oblique wing was conceived by German aerodynamicists in the midst of the Second World War. But it was only afterward, through the

brilliance and determination of NASA aerodynamicist Robert T. Jones that it advanced to actual flight. Indeed, Jones, father of the American swept wing, became one of the most persistent proponents of the oblique wing concept.[954] The principal advantage of the oblique wing is that it spreads both the lift and volume distributions of the wing over a greater length than that of a simple symmetrically swept wing. This has the effect of reducing both the wave drag because of lift and the wave drag because of volume, two important components of supersonic drag. With this the­oretical advantage come practical challenges. The challenges fall into two broad categories: the effects of asymmetry on the flight character­istics (stability and handling qualities) of the vehicle, and the aeroelas – tic stability of the forward-swept wing. The research objectives of the AD-1 were primarily oriented toward flying qualities. The AD-1 was not intended to explore structural dynamics or divergence in depth, other

than establishing safety of flight. Mike Rutkowski analyzed the wing for flutter and divergence using NASTRAN and other methods.[955]

However, the project did make a significant accomplishment in the use of static aeroelastic tailoring. The fiberglass wing design by Ron Smith was tailored to bend just enough, with increasing g, to cancel out an aero­dynamically induced rolling moment. Pure bending of the oblique wing increases the incidence (and therefore the lift) of the forward-swept tip and decreases the incidence (and lift) of the aft-swept tip. In a pullup maneuver, increasing lift coefficient (CL), and load factor at a given flight condition, this would cause a rollaway from the forward-swept tip. At the same time, induced aerodynamic effects (the downwash/upwash distribu­tion) increase the lift at the tip of an aft-swept wing. On an aircraft with only one aft-swept tip, this would cause a roll toward the forward-swept side. The design intent for the AD-1 was to have these two effects cancel each other as nearly as possible, so that the net change in rolling moment because of increasing g at a given flight condition would be zero. The design condition was CL = 0.3 for 1-g flight at 170 knots, 12,500-foot alti­tude, and a weight of 1,850 pounds, with the wing at 60-degree sweep.[956]

An aeroelastically scaled one-sixth model was tested at full-scale Reynolds number in the Ames 12-Foot Pressure Wind Tunnel. A stiff alu­minum wing was used for preliminary tests, then two fiberglass wings. The two fiberglass wings had zero sweep at the 25- and 30-percent chord lines, respectively, bracketing the full-scale AD-1 wing, which had zero sweep at 27-percent chord. The wings were tested at the design scaled dynamic pressure and at two lower values to obtain independent varia­tion of wing load because of angle of attack and dynamic pressure at a constant angle of attack. Forces and moments were measured, and deflec­tion was determined from photographs of the wing at test conditions.[957]

Subsequently, ". . . the actual wing deflection in bending and twist was verified before flight through static ground loading tests.” Finally, in-flight measurements were made of total force and moment coeffi­cients and of aeroelastic effects. Level-flight decelerations provided angle-of-attack sweeps at constant load, and windup turns provided angle-of-attack sweeps at constant "q” (dynamic pressure). Results were interpreted and compared with predictions. The simulator model, with aeroelastic effects included, realistically represented the dynamic responses of the flight vehicle.[958]

Provision had been made for mechanical linkage between the pitch and roll controls, to compensate for any pitch-roll coupling observed in flight. However, the intent of the aeroelastic wing was achieved closely enough that the mechanical interconnect was never used.[959] Roll trim was not needed at the design condition (60-degrees sweep) nor at zero sweep, where the aircraft was symmetric. At intermediate sweep angles, roll trim was required. The correction of this characteristic was not pur­sued because it was not a central objective of the project. Also, the air­plane experienced fairly large changes in rolling moment with angle of attack beyond the linear range. Vortex lift, other local flow separa­tions, and ultimately full stall of the aft-swept wing, occurred in rapid succession as angle of attack was increased from 8 to approximately 12 degrees. Therefore, it would be a severe oversimplification to say that the AD-1 had normal handling qualities.[960]

The AD-1 flew at speeds of 170 knots or less. On a large, high-speed aircraft, divergence of the forward-swept wing would also be a consid­eration. This would be addressed by a combination of inherent stiffness, aeroelastic tailoring to introduce a favorable bend-twist coupling, and, potentially, active load alleviation. The AD-1 project provided initial cor­relation of measured versus predicted wing bending and its effects on the vehicle’s flight characteristics. NASA planned to take the next step with a supersonic oblique wing aircraft, using the same F-8 airframe that had been used for earlier supercritical wing tests. These studies delved deeper into the aeroelastic issues: "Preliminary studies have been performed to identify critical DOF [Degree of Freedom] for flut­ter model tests of oblique configurations. An ‘oblique’ mode has been identified with a 5 DOF model which still retains its characteristics with the three rotational DOF’s. An interdisciplinary analysis code (STARS), which is capable of performing flutter and aeroservoelastic analyses, has been developed. The structures module has a large library of elements and in conjunction with numerical analysis routines, is capable of effi­ciently performing statics, vibration, buckling, and dynamic response analysis of structures. . . . ” The STARS code also included supersonic (potential gradient method) and subsonic (doublet lattice) unsteady aero­dynamics calculations. " . . . Linear flutter models are developed and transformed to the body axis coordinate system and are subsequently augmented with the control law. Stability analysis is performed using hybrid techniques. The major research benefit of the OWRA [Oblique Wing Research Aircraft] program will be validation of design and anal­ysis tools. As such, the structural model will be validated and updated based on ground vibration test (GVT) results. The unsteady aero codes will be correlated with experimentally measured unsteady pressures.”[961] While the OWRA program never reached flight, (NASA was ready to begin wing fabrication in 1987, expecting first flight in 1991), these comments illustrate the typical interaction of flight programs with ana­lytical methods development and the progressive validation process that takes place. Such methods development is often driven by unconven­tional problems (such as the oblique wing example here) and eventually finds its way into routine practice in more conventional applications. For example, in the design of large passenger aircraft today, the loads process is typically iterated to include the effects of static aeroelastic deflections on the aerodynamic load distribution.[962]

X-29

The Grumman X-29 aircraft was an extraordinarily ambitious and pro­ductive flight-test program run between 1984 and 1992. It demonstrated a large (approximately 35 percent) unstable static margin in the pitch axis, a digital active flight control system utilizing three-surface pitch control (all-moving canards, wing flaps, and aft-mounted strake flaps), and a thin supercritical forward-swept wing, aeroelastically tailored to prevent struc­tural divergence. The X-29 was funded by the Defense Advanced Research Projects Agency (DARPA) through the USAF Aeronautical Systems Division (ASD). Grumman was responsible for aircraft design and fabrication, including the primary structural analyses, although there was exten­sive involvement of NASA and the USAF in addressing the entire realm of unique technical issues on the project. NASA Ames Research Center/ Dryden Flight Research Facility was the responsible test organization.[963]

Careful treatment of aeroelastic stability was necessary for the thin FSW to be used on a supersonic, highly maneuverable aircraft. According to Grumman, "Automated design and analysis procedures played a major role in the development of the X-29 demonstrator aircraft.” Grumman used one of its programs, called FASTOP, to optimize the X-29’s structure to avoid aeroelastic divergence while minimizing the weight impact.[964]

In contrast to the AD-1, which allowed the forward-swept wing to bend along its axis, thereby increasing the lift at the forward tip, the X-29’s forward-swept wings were designed to twist when bending, in a manner that relieved the load. This was accomplished by orienting the primary spanwise fibers in the composite skins at a forward "kick angle” relative to the nominal structural axis of the wing. The optimum angle was found in a 1977 Grumman feasibility study: "Both beam and coarse- grid, finite-element models were employed to study various materials and laminate configurations with regard to their effect on divergence and flutter characteristics and to identify the weight increments required to avoid divergence.”[965] While a pure strength design was optimum at zero kick angle, an angle of approximately 10 degrees was found to be best for optimum combined strength and divergence requirements.

When the program reached the flight-test phase, hardware-in-the – loop simulation was integral to the flight program. During the func­tional and envelope expansion phases, every mission was flown on the simulator before it was flown in the airplane.[966] In flight, the X-29 No. 1 aircraft (of two that were built) carried extensive and somewhat unique instrumentation to measure the loads and deflections of the air­frame, and particularly of the wing. This consisted of pressure taps on the left wing and canard, an optical deflection measurement system on the right wing, strain gages for static structural load measurement, and accelerometers for structural dynamic and buffet measurement.

The most unusual element of this suite was the optical system, which had been developed and used previously on the HiMAT demonstrator (see preceding description). Optical deflection data were sampled at a

rate of 13 samples per channel per second. Data quality was reported to be very good, and initial results showed good match to predictions. In addition, pressure data from the 156 wing and 17 canard pressure taps was collected at a rate of 25 samples per channel per second. One hundred six strain gages provided static loads measurement as shown. Structural dynamic data from the 21 accelerometers was measured at 400 samples per channel per second. All data was transmitted to ground station and, during limited-envelope phase, to Grumman in Calverton, NY, for analysis.[967] "Careful analyses of the instrumentation requirements, flight test points, and maneuvers are conducted to ensure that data of sufficient quality and quantity are acquired to validate the design, fab­rication, and test process.”[968] The detailed analysis and measurements provided extensive opportunities to validate predictive methods.

The X-29 was used as a test case for NASA’s STARS structural anal­ysis computer program, which had been upgraded with aeroservoelas- tic analysis capability. In spite of the exhaustive analysis done ahead of time, there were, as is often the case, several "discoveries” made during flight test. Handling qualities at high alpha were considerably better than predicted, leading to an expanded high-alpha control and maneuverability investigation in the later phases of the project. The X-29

No. 1 was initially limited to 21-degree angle of attack, but, during sub­sequent Phase II envelope expansion testing, its test pilots concluded it had "excellent control response to 45 deg. angle of attack and still had limited controllability at 67 deg. angle of attack.”[969]

There were also at least two distinct types of aeroservoelastic phe­nomena encountered: buffet-induced modes and a coupling between the canard position feedback and the aircraft’s longitudinal aerody­namic and structural modes were observed.[970] The modes mentioned involved frequencies between 11 and 27 hertz (Hz). Any aircraft with an automatic control system may experience interactions between the aircraft’s structural and aerodynamic modes and the control system. Typically, the aeroelastic frequencies are much higher than the charac­teristic frequencies of the motion of the aircraft as a whole. However, the 35-percent negative static margin of the X-29A was much larger than any unstable margin designed into an aircraft before or since. As a consequence, its divergence timescale was much more rapid, making it particularly challenging to tune the flight control system to respond quickly enough to aircraft motions, without being excited by structural dynamic modes. Simply stated, the X-29A provided ample opportunity for aeroservoelastic phenomena to occur, and such were indeed observed, a contribution of the aircraft that went far beyond simply demonstrat­ing the aerodynamic and maneuver qualities of an unstable forward – swept canard planform.[971]

In sum, each of these five advanced flight projects provides impor­tant lessons learned across many disciplines, particularly the validation of computer methods in structural design and/or analysis. The YF-12 project provided important correlation of analysis, ground-test data, and flight data for an aircraft under complex aerothermodynamic load­ing. The Rotor Aerodynamic Limits survey collected important data on helicopter rotors—a class of system often taken for granted yet one that represent an incredibly complex interaction of aerodynamic, aeroelas – tic, and inertial phenomena. The HiMAT, AD-1, and X-29 programs each advanced the state of the art in aeroelastic design as applied to nontra­ditional, indeed exotic, planforms featuring unstable design, compos­ite structures, and advanced flight control concepts. Finally, the data required to validate structural analysis and design methods do not auto­matically come from the testing normally performed by aircraft devel­opers and users. Special instrumentation and testing techniques are required. NASA has developed the facilities and the knowledge base needed for many kinds of special testing and is able assign the required priority to such testing. As these cases show, NASA therefore plays a key role in this process of gaining knowledge about the behavior of aircraft in flight, evaluating predictive capabilities, and flowing that experience back to the people who design the aircraft.

During World War II the whole of aeronautics used aluminum. There was no hypersonics; the very word did not exist, for it took until 1946 for the investigator Hsue-shen Tsien to introduce it. Germany’s V-2 was flying at Mach 5, but its nose cone was of mild steel, and no one expected that this simple design problem demanded a separate term for its flight regime.[1018]

A decade later, aeronautics had expanded to include all flight speeds because of three new engines: the liquid-fuel rocket, the ramjet, and the variable-stator turbojet. The turbojet promised power beyond Mach 3, while the ramjet proved useful beyond Mach 4. The Mach 6 X-15 was under contract. Intermediate-range missiles were in development, with ranges of 1,200 to 1,700 miles, and people regarded intercontinental missiles as preludes to satellite launchers.

A common set of descriptions presents the flight environments within which designers must work. Well beyond Mach 3, engineers accommo­date aerodynamic heating through materials substitutions. The aircraft themselves continue to accelerate and cruise much as they do at lower speeds. Beyond Mach 4, however, cruise becomes infeasible because of heating. A world airspeed record for air-breathing flight (one that lasted for nearly the next half century) was set in 1958 with the Lockheed X-7, which was made of 4130 steel, at Mach 4.31 (2,881 mph). The airplane had flown successfully at Mach 3.95, but it failed structurally in flight at Mach 4.31, and no airplane has approached such performance in the past half century.[1019]

No aircraft has ever cruised at Mach 5, and an important reason involves structures and materials. "If I cruise in the atmosphere for 2 hours,” said Paul Czysz of McDonnell-Douglas, "I have a thousand times the heat load into the vehicle that the Shuttle gets on its quick transit of the atmosphere.”[1020] Aircraft indeed make brief visits to such speed regimes, but they don’t stay there; the best approach is to pop out of the atmosphere and then return, the hallmark of a true trans­atmospheric vehicle.

At Mach 4, aerodynamic heating raises temperatures. At higher Mach, other effects are seen. A reentering intercontinental ballistic mis­sile (ICBM) nose cone, at speeds above Mach 20, has enough kinetic energy to vaporize 5 times its weight in iron. Temperatures behind its bow shock reach 9,000 kelvins (K), hotter than the surface of the Sun. The research physicist Peter Rose has written that this velocity would be "large enough to dissociate all the oxygen molecules into atoms, dissociate about half of the nitrogen, and thermally ionize a consider­able fraction of the air.”[1021]

Aircraft thus face a simple rule: they can cruise up to Mach 4 if built with suitable materials, but they cannot cruise at higher speeds. This rule applies not only to entry into Earth’s atmosphere but also to entry into the atmosphere of Jupiter, which is far more demanding but which an entry probe of the Galileo spacecraft investigated in 1995, at Mach 50.[1022]

Other speed limits become important in the field of wind tunnel simulation. The Government’s first successful hypersonic wind tun­nel was John Becker’s 11-inch facility, which entered service in 1947. It approached Mach 7, with compressed air giving run times of 40 sec­onds.[1023] A current facility, which is much larger and located at the National Aeronautics and Space Administration (NASA) Langley Research Center, is the Eight-Foot High-Temperature Tunnel—which also uses compressed air and operates near Mach 7.

The reason for such restrictions involves fundamental limitations of compressed air, which liquefies if it expands too much when seeking higher speeds. Higher speeds indeed are achievable but only by creat­ing shock waves within an instrument for periods measured in milli­seconds. Hence, the field of aerodynamics introduces an experimental speed limit of Mach 7, which describes its wind tunnels, and an opera­tional speed limit of Mach 4, which sets a restriction within which cruis­ing flight remains feasible. Compared with these velocities, the usual definition of hypersonics, describing flight beyond Mach 5, is seen to describe nothing in particular.

In mid-1969, modifications began to convert the prototype McDonnell – Douglas YF-4E (USAF serial No. 62-12200) for the SFCS program. A quadruple-redundant analog computer-based three-axis fly-by-wire flight control system with integrated hydraulic servo-actuator packages was incorporated and side stick controllers were added to both the front and back cockpits. Roll control was pure fly-by-wire with no mechani­cal backup. For initial testing, the Phantom’s mechanical flight control system was retained in the pitch and yaw axes as a safety backup. On April 29, 1972, McDonnell-Douglas test pilot Charles P. "Pete” Garrison flew the SFCS YF-4E for the first time from the McDonnell-Douglas fac­tory at Lambert Field in St. Louis, MO. The mechanical flight control system was used for takeoff with the pilot switching to the fly-by-wire system during climb-out. The aircraft was then flown to Edwards AFB for a variety of additional tests, including low-altitude supersonic flights. After the first 27 flights, which included 23 hours in the full three-axis fly­by-wire configuration, the mechanical flight control system was disabled. First flight in the pure fly-by-wire configuration occurred January 22, 1973. The SFCS YF-4E flew as a pure fly-by-wire aircraft for the remain­der of its flight-test program, ultimately completing over 100 flights.[1138]

Whereas the earlier phases of the flight-test effort were primarily flown by McDonnell-Douglas test pilots, the next aspect of the SFCS

program was focused on an Air Force evaluation of the operational suitability of fly-by-wire and an assessment of the readiness of the tech­nology for transition into new aircraft designs. During this phase, 15 flights were accomplished by two Air Force test pilots (Lt. Col. C. W. Powell and Maj. R. C. Ettinger), who concluded that fly-by-wire was indeed ready and suitable for use in new designs. They also noted that flying qualities were generally excellent, especially during takeoffs and landings, and that the pitch transient normally encountered in the F-4 during rapid deceleration from supersonic to subsonic flight was nearly eliminated. Another aspect of the flight-test effort involved so – called technology transition and demonstration flights in the SFCS aircraft. At this time, the Air Force had embarked on the Lightweight Fighter (LWF) program. One of the two companies developing flight demonstrator aircraft (General Dynamics)had elected to use fly-by-wire in its new LWF design (the YF-16). A block of 11 flights in the SFCS YF-4E was allocated to three pilots assigned to the LWF test force at Edwards AFB (Lt. Col. Jim Ryder, Maj. Walt Hersman, and Maj. Mike Clarke). Based on their experiences flying the SFCS YF-4E, the LWF test force pilots were able to provide valuable inputs into the design, devel­opment, and flight test of the YF-16, directly contributing to the dra­matic success of that program. An additional 10 flights were allocated to another 10 pilots, who included NASA test pilot Gary E. Krier and USAF Maj. Robert Barlow.[1139] Earlier, Krier had piloted the first flight of a digital fly-by-wire (DFBW) flight control system in the NASA DFBW F-8C on May 25, 1972. That event marked the first time that a piloted aircraft had been flown purely using a fly-by-wire flight control system without any mechanical backup provisions. Barlow, as a colonel, would command the Air Force Flight Dynamics Laboratory during execution of several important fly-by-wire flight research efforts. The Air Force YF-16 and the NASA DFBW F-8 programs are discussed in following sections.

During 1997, NASA Dryden had evaluated a single electrohydrostatic actuator installation on the NASA F-18 Systems Research Aircraft (SRA), with the primary goal being the flight demonstration of power-by-wire technology on a single primary flight control surface. The electrohydro­static actuator, provided by the Air Force, replaced the F-18’s standard left aileron actuator and was evaluated throughout the aircraft’s flight envelope out to speeds of Mach 1.6. Numerous mission profiles were accomplished that included a full series of aerobatic maneuvers. The electrohydrostatic actuator accumulated 23.5 hours of flight time on the F-18 SRA between January and July 1997. It performed as well as the standard F-18 actuator and was shown to have more load capabil­ity than required by the aileron actuator specification for the aircraft.[1188]

At about the same time, a Joint Strike Fighter/Integrated Subsystems Technology program had been formed to reduce the risk of selected

technology candidates, in particular the power-by-wire approach that was intended to replace cumbersome hydraulic actuation systems with all-electrical systems for flight surface actuation. A key to this effort was the AFTI F-16, which was modified to replace all of the standard hydrau­lic actuators on the primary flight control surfaces with electrohydro­static actuators (EHAs) to operate the flaperons, horizontal tails, and rudder. Each electrohydrostatic actuator uses an internal electric motor to drive an integral hydraulic pump, thus it relies on local hydraulics for force transmission (similar to the approach used with the Powered Flight Control Units on the Vickers VC10 aircraft discussed earlier).[1189]

In a conventional F-16, the digital fly-by-wire flight control system sends out electrical command signals to each of the flight control actu­ators. These electrical signals drive the control valves (located with the actuators) that schedule the fluid from the high-pressure hydraulic pump to position the flight control surfaces. Dual engine-driven 3,000 pounds per square inch (psi) hydraulic systems power each primary control sur­face actuator to drive the control surfaces to the desired position. The standard F-16 hydraulic actuators operate continuously at 3,000 psi, and power is dumped into the actuators, whether it is needed or not.[1190] In straight and level flight (where most aircraft operate most of their time, including even high-performance fighters), the actual electrical power requirement of the actuation system is low (only about 500 watts per actuator), and excess energy is dissipated as heat and is transferred into the fuel system.[1191]

With the electrohydrostatic power design tested in the AFTI/F-16, the standard fly-by-wire flight control system was relatively unchanged. However, the existing F-16 hydraulic power system was removed and replaced by a new power-by-wire system, consisting of an engine-driven Hamilton Sundstrand dual 270-volt direct current (DC) electrical power generation system (to provide redundancy) and Parker Aerospace elec­trohydrostatic actuators on the flaperons, rudder, and horizontal sta­bilizer. The new electrical system powers five dual power electronics units, one for each flight control surface actuator. Each power electron­ics unit regulates the DC electrical power that drives dual motor/pumps that are self-contained in each electrohydrostatic actuator. The dual
motor/pumps convert electrical power into hydraulic power, allowing the piston on the actuators to move the control surfaces. The electrohy­drostatic actuators operate at pressures ranging from 300 to 3,000 psi, providing power only on demand and generating much less heat. An electrical distribution and electrical actuation system simplifies second­ary power and thermal management systems, because the need to pro­vide secondary and emergency backup sources of hydraulic power for the flight control surfaces is eliminated. The electrohydrostatic system also provides more thermal margin, which can be applied to cooling other high-demand systems (such as avionics and electronic warfare), or, alternatively, the thermal management system weight and volume can be reduced making new aircraft designs smaller, lighter, and more affordable. Highly integrated electrical subsystems, including power-by­wire, reportedly could reduce takeoff weight by 6 percent, vulnerable area by 15 percent, procurement cost by 5 percent, and total life-cycle cost by 2 to 3 percent, compared with current fighters based on Air Force and industry studies. The power-by-wire approach is now being used in the Lockheed Martin F-35 Lightning II, with the company estimat­ing a reduction in aircraft weight of as much as 700 pounds because of weight reductions in the hydraulic system, the secondary power sys­tem, and the thermal management system, made possible because the electrical power-by-wire system produces less heat than the traditional hydraulic system that it replaces.[1192]

The modified power-by-wire AFTI/F-16 was the first piloted aircraft of any type to fly with a totally electric control surface actuation system with no hydraulic or mechanical backup flight control capability of any kind. It was designed to have the same flight control system responses as an unmodified F-16. After the first power-by-wire AFTI/F-16 flight on October 24, 2000, at Fort Worth, Lockheed Martin test pilot Steve Barter stated aircraft handling qualities with the power-by-wire modi­fications were indistinguishable from that of the unmodified AFTI/F-16. The aircraft was subsequently flown about 10 times, with flight control effectiveness of the power-by-wire system demonstrated during super­sonic flight. Test pilots executed various flying quality maneuvers, includ­ing high-g turns, control pulses (in pitch, roll, and yaw), doublet inputs, and sideslips. The tests also included simulated low-altitude attack

missions and an evaluation of the electrostatic actuator and generator subsystems and their thermal behavior under mission loads.[1193]

NASA Dryden hosted the AFTI/F-16 program for 16 years, from 1982 to 1998. During that time, personnel from Dryden composed 50 percent of the AFTI joint test team. Dryden pilots who flew the AFTI/F-16 included Bill Dana, Dana Purifoy, Jim Smolka, Rogers Smith, and Steve Ishmael. Dryden responsibilities, in addition to its host role, included flight safety, operations, and maintenance. Mark Skoog, who served as the USAF AFTI/F-16 project manager for many years and later became a NASA test pilot, commented: "AFTI had the highest F-16 sortie success rate on base, due to Dryden maintenance personnel having tremendous exper­tise in tailoring their operations to the uniqueness of the vehicle. That includes all the other F-16s based at Edwards during those years, none of which were nearly as heavily modified as the AFTI.”[1194] A good summary of the AFTI/F-16’s accomplishments was provided by NASA test pilot Dana Purifoy: "Flying AFTI was a tremendous opportunity. The aircraft pineered many important technologies including glass cockpit human factors, automated ground collision avoidance, integrated night vision capability and on-board data link operations. All of these technologies are cur­rently being implemented to improve the next generation of both civil and military aircraft.”[1195] The AFTI F-16’s last flight at Dryden was on November 4, 1997. Over a period of 15 years, it made over 750 flights and was flown by 23 pilots from the U. S. Air Force, NASA, the U. S. Marine Corps, and the Swedish Air Force. The AFTI F-16 then served as an Air Force technol­ogy testbed. Experience and lessons learned were used to help develop the production DFBW flight control system used in the F-16. The F-16, the F-22, and the F-35, in particular, directly benefited from AFTI/F-16 research and technology maturation efforts. After 22 years as a research aircraft for NASA and the Air Force, the AFTI F-16 was flown to Wright-Patterson AFB, OH, on January 9, 2001, for display at the Air Force Museum.[1196]

The Self-Repairing Flight Control System (SRFCS) consists of software integrated into an aircraft’s digital flight control system that is used to detect failures or damage to the aircraft control surfaces. In the event of control surface damage, the remaining control surfaces are automat­ically reconfigured to maintain control, enabling pilots to complete their mission and land safely. The program, sponsored by the U. S. Air Force, demonstrated the ability of a flight control system to identify the failure of a control surface and reconfigure commands to other control devices, such as ailerons, rudders, elevators, and flaps, to continue the aircraft’s mission or allow it to be landed safely. As an example, if the horizontal elevator were damaged or failed in flight, the SRFCS would diagnose the failure and determine how the remaining flight control surfaces could be repositioned to compensate for the damaged or inoperable control surface. A visual warning to the pilot was used to explain the type of fail­ure that occurred. It also provided revised aircraft flight limits, such as reduced airspeed, angle of attack, and maneuvering loads. The SRFCS also had the capability of identifying failures in electrical, hydraulic, and mechanical systems. Built-in test and sensor data provided a diag­nostic capability and identified failed components or system faults for subsequent ground maintenance repair. System malfunctions on an air­craft with a SRFCS can be identified and isolated at the time they occur and then repaired as soon as the aircraft is on the ground, eliminating lengthy postflight maintenance troubleshooting.[1267]

The SRFCS was flown 25 times on the HIDEC F-15 at NASA Dryden between December 1989 and March 1990, with somewhat mixed results. The maintenance diagnostics aspect of the system was a general suc­cess, but there were frequent failures with the SRFCS. Simulated con­trol system failures were induced, with the SRFCS correctly identifying every failure that it detected. However, it only sensed induced control system failures 61 percent of the time. The overall conclusion was that the SRFCS concept was promising, but it needed more develop­ment if it was to be successfully implemented into production aircraft.

NASA test pilot Jim Smolka flew the first SRFCS flight, on December 12, 1989, with test engineer Gerard Schkolnik in the rear cockpit; other SRFCS test pilots were Bill Dana and Tom McMurtry.[1268]

While most eyes were on the big picture of making major engine advance­ments through the years, some very specific problems were addressed with programs that are just as interesting to consider as the larger research endeavors. The casings that surround the jet engine’s turbo­machinery are a case in point.

With the 1989 crash of United Airlines Flight 232 at Sioux City, IA, aviation safety officials became more interested in finding new materials capable of containing the resulting shrapnel created when a jet engine’s blade or other component breaks free. In the case of the DC-10 involved in this particular crash, the fan disk of the No. 2 engine—the one located in the tail—separated from the engine and caused the powerplant to explode, creating a rain of shrapnel that could not be contained within the engine casing. The sharp metal fragments pierced the body of the aircraft and cut lines in all three of the aircraft’s hydraulic systems. As previously mentioned in this case study, the pilots on the DC-10 were able to steer their aircraft to a nearly controlled landing. The incident inspired NASA pilots to refine the idea of using only jet thrust to maneuver an airplane and undertake the Propulsion Controlled Aircraft program, which took full advantage of the earlier Digital Electronic Engine Control research. The Iowa accident also sent structures and materials experts off on a hunt to find a way to prevent accidents like this in the future.

The United Flight 232 example notwithstanding, the challenge for structures engineers is to design an engine casing that will contain a failed fan blade within the engine so that it has no chance to pierce the passenger compartment wall and threaten the safety of passengers or cause a catastrophic tear in the aircraft wall. Moreover, not only does the casing have to be strong enough to withstand any blade or shrapnel impacts, it must not lose its structural integrity during an emergency

engine shutdown in flight. A damaged engine can take some 15 seconds to shut down, during which time cracks from the initial blade impacts can propagate in the fan case. Should the fan case totally fail, the result­ing breakup of the already compromised turbomachinery could be cat­astrophic to the aircraft and all aboard.[1360]

As engineers considered the use of composite materials, two methods for containing blade damage within the engine casing were now available: the new softwall and the traditional hardwall. In the softwall concept, the casing was made of a sandwich-type aluminum structure overwound with dry aramid fibers. (Aramid fibers were introduced commercially by DuPont during the early 1960s and were known by the trade name Nomex.) The design allows broken blades and other shrapnel to pass through the "soft” aluminum and be stopped and contained within the aramid fiber wrap. In the hardwall approach, the casing is made of aluminum only and is built as a rigid wall to reflect blade bits and other collateral damage back into the casing interior. Of course that vastly increases the risk that the shrap­nel will be ingested through the engine and cause even greater damage, perhaps catastrophic. While that risk exists with the softwall design, it is not as substantial. Another benefit of the hardwall is that it maintains its structural soundness, or ductility, during a breakup of an engine. A softwall also features some amount of ductility, but the energy-absorb­ing properties of the aramid fibers is the major draw.[1361]

In 1994, NASA engineers at the Lewis Research Center began look­ing into better understanding engine fan case structures and conducted impact tests as part of the Enabling Propulsion Materials program. Various metallic materials and new ideas for lightweight fan contain­ment structures were studied. By 1998, the research expanded to include investigations into use of polymer composites for engine fan casings. As additional composite materials were made available, NASA researchers sought to understand their properties and the appropriateness of those materials in terms of containment capability, damage tolerance, com­mercial viability, and understanding any potential risk not yet identi­fied for their use on jet engines.[1362]

In 2001, NASA awarded a Small Business Innovation Research (SBIR) grant to A&P Technology, Inc., of Cincinnati to develop a dam­age-tolerant fan casing for a jet engine. Long before composites came along, the company’s expertise was in braiding materials together, such as clotheslines and candlewicks. A&P—working together with the FAA, Ohio State University, and the University of Akron—was able to rapidly develop a prototype composite fan case that could be compared to the metal fan case. Computer simulations were key to the effort and seren­dipitously provided an opportunity to grow the industry’s understand­ing and ability to use those very same simulation capabilities. First, well understood metallic casings undergoing a blade-out scenario were modeled, and the computer tested the resulting codes to reproduce the already-known results. Then came the trick of introducing code that would represent A&P’s composite casing and its reaction to a blade-out situation. The process was repeated for a composite material wrapped with a braided fiber material, and results were very promising.[1363]

The composite casing proposed by A&P used a triaxial carbon braid, which has a toughness superior to aluminum but is lighter, which helps ease fuel consumption. In tests of debris impact, the braided laminate performed better than the metal casing, because in some cases, the com­posite structure absorbed the energy of the impact as the debris bounced off the wall, and in other cases where the shrapnel penetrated the mate­rial, the damage to the wall was isolated to the impact point and did not spread. In a metal casing that was pierced, the resulting hole would instigate several cracks that would continue to propagate along the cas­ing wall, appearing much like the spiderweb of cracks that appear on an automobile windshield when it is hit with a small stone on the freeway.

NASA continues to study the use of composite casings to better understand the potential effects of aging and/or degradation following the constant temperature, vibration, and pressure cycles a jet engine experiences during each flight. There also is interest in studying the effects of higher operating temperatures on the casing structure for pos­sible use on future supersonic jets. (The effect of composite fan blades on casing containment also has been studied.)[1364]

While composites have found many uses in commercial and military aviation, the first use of an all-composite engine casing, provided by A&P, is set to be used on GE’s GEnx turbojet designed for the Boeing 787. The braided casing weighs 350 pounds less per engine, and, when other engine installation hardware to handle the lighter powerplants is considered, the 787 should weigh 800 pounds less than a similarly equipped airliner using aluminum casings. The weight reduction also should provide a savings in fuel cost, increased payload, and/or a greater range for the aircraft.[1365]

The primary objective of the Federal Wind Energy Program and the specific objectives of NASA’s portion of the program were outlined in a followup technical paper presented in 1975 by Thomas, Savino, and Richard L. Puthoff. The paper noted that the overall objective of the
program was "to develop the technology for practical cost-competitive wind-generator conversion systems that can be used for supplying sig­nificant amounts of energy to help meet the nation’s energy needs.”[1499] The specific objectives of NASA Lewis’s portion of the program were to: (1) identify cost-effective configurations and sizes of wind-conversion systems; (2) develop the technology needed to produce cost-effective, reliable systems; (3) design wind turbine generators that are compati­ble with user applications, especially with electric utility networks; (4) build up industry capability in the design and fabrication of wind tur­bine generators; and (5) transfer the technology from the program to industry for commercial application. To satisfy these objectives, NASA Lewis divided the development function into the three following areas: (1) design, fabrication, and testing of a 100-kilowatt experimental wind turbine generator; (2) optimizing the wind turbines for selected user operation; and (3) supporting research and technology for the systems.

The planned workload was divided further by assignment of dif­ferent tasks to different NASA Research Centers and industry partici­pants. NASA Lewis would provide project management and support in aerodynamics, instrumentation, structural dynamics, data reduction, machine design, facilities, and test operations. Other NASA Research Centers would provide consulting services within their areas of expertise. For example, Langley worked on aeroelasticity matters, Ames consulted on rotor dynamics, and Marshall provided meteorology support. Initial industry participants included Westinghouse, Lockheed Corporation, General Electric, Boeing, and Kaman Aerospace.

In order to undertake its project management role, NASA Lewis established the Center’s Wind Power Office, which consisted initially of three operational units—one covering the development of an experi­mental 100-kilowatt wind turbine, one handling the industry-built util­ity-operated wind turbines, and one providing supporting research and technology. The engineers in these offices basically worked together in a less formal structure, crossing over between various operational areas. Also, the internal organization apparently underwent several changes during the program’s existence. For example, in 1976, the program was

directed by the Wind Power Office as part of the Solar Energy Branch. The first two office managers were Ronald Thomas and William Robbins. By 1982, the organization consisted of a Wind Energy Project Office, which was once again under the supervision of Thomas and was part of the Wind and Stationary Power Division. The office consisted of a proj­ect development and support section under the supervision of James P. Couch (who managed the Mod-2 project), a research and technology sec­tion headed by Patrick M. Finnegan, and a wind turbine analysis section under the direction of David A. Spera. By 1984, the program organiza­tion had changed again with the Wind Energy Project Office, which was under the supervision of Darrell H. Baldwin, becoming part of the Energy Technology Division. The office consisted of a technology section under Richard L. Puthoff and an analysis section headed by David A. Spera. The last NASA Lewis wind energy program manager was Arthur Birchenough.

Whitcomb joined the research community at Langley in 1943 as a mem­ber of Stack’s Transonic Aerodynamics Branch working in the 8-foot High-Speed Tunnel (HST). Initially, NACA managers placed him in the Flight Instrument Research Division, but Whitcomb’s force of person­ality ensured that he would be working directly on problems related to aircraft design. As many of his colleagues and historians would attest, Whitcomb quickly became known for an analytical ability rooted in mathematics, instinct, and aesthetics.[145]

In 1945, Langley increased the power of the 8-foot HST to gener­ate Mach 0.95 speeds, and Whitcomb was becoming increasingly famil­iar with transonic aerodynamics, which helped him in his developing investigation into the design of supersonic aircraft. The onset of drag created by shock waves at transonic speeds was the primary challenge. John Stack, Ezra Kotcher, and Lawrence D. Bell proved that breaking the sound barrier was possible when Chuck Yeager flew the Bell X-1 to Mach 1.06 (700 mph) on October 14, 1947. Designed in the style of a.50- caliber bullet with straight wings, the Bell X-1 was a successful super­sonic airplane, but it was a rocket-powered research airplane designed specifically for and limited to that purpose. The X-1 would not offer designers the shape of future supersonic airplanes. Operational turbojet – powered aircraft designed for military missions were much heavier and would use up much of their fuel gradually accelerating toward Mach 1 to lessen transonic drag.[146] The key was to get operation aircraft through the transonic regime, which ranged from Mach 0.9 to Mach 1.1.

A very small body of transonic research existed when Whitcomb undertook his investigation. British researchers W. T. Lord of the Royal Aeronautical Establishment and G. N. Ward of the University of Manchester and American Wallace D. Hayes attempted to solve the problem of transonic drag through mathematical analyses shortly after World War II in 1946. These studies generated mathematical symbols that did not lend themselves to the design and shape of transonic and supersonic aircraft.[147]

Whitcomb’s analysis of available data generated by the NACA in ground and free-flight tests led him to submit a proposal for testing swept wing and fuselage combinations in the 8-foot HST in July 1948. There had been some success in delaying transonic drag by addressing the relationship between wing sweep and fuselage shape. Whitcomb believed that careful attention to arrangement and shape of the wing and fuselage would result in their counteracting each other. His goal was to reach a milestone in supersonic aircraft design. The tests, conducted from late 1949 to early 1950, revealed no significant decrease in drag at high subsonic (Mach 0.95) and low supersonic (Mach 1.2) speeds. The wing-fuselage combinations actually generated higher drag than their individual values combined. Whitcomb was at an impasse and realized he needed to refocus on learning more about the fundamental nature of transonic airflow.[148]

Just before Whitcomb had submitted his proposal for his wind tun­nel tests, John Stack ordered the conversion of the 8-foot HST in the spring of 1948 to a slotted throat to enable research in the transonic regime. In theory, slots in the tunnel’s test section, or throat, would enable smooth operation at very high subsonic speeds and at low supersonic speeds. The initial conversion was not satisfactory because of uneven flow. Whitcomb and his colleagues, physicist Ray Wright and engineer Virgil S. Ritchie, hand-shaped the slots based on their visualization of smooth transonic flow. They also worked directly with Langley wood­workers to design and fabricate a channel at the downstream end of the test section that reintroduced air that traveled through the slots. Their painstaking work led to the inauguration of transonic operations within the 8-foot HST 7 months later, on October 6, 1950.[149] Whitcomb,

as a young engineer, was helping to refine a tunnel configuration that was going to allow him to realize his potential as a visionary experimen­tal aeronautical engineer.

The NACA distributed a confidential report on the new tunnel during the fall of 1948, which was distributed to the military services and select manufacturers. By the following spring, rumors had been circulating about the new tunnel throughout the industry. Initially, the call for secrecy evolved into outright public acknowledgement of the NACAs new tran­sonic tunnels (including the 16-foot HST) with the awarding of the 1951 Collier Trophy to John Stack and 19 of his associates at Langley for the slotted wall. The Collier Trophy specifically recognized the importance of a research tool, which was a first in the 40-year history of the award. The NACA claimed that its slotted-throat transonic tunnels gave the United States a 2-year lead in the design of supersonic military aircraft.[150]

With the availability of the 8-foot HST and its slotted throat, the com­bined use of previously available wind tunnel components—the tunnel bal­ance, pressure orifice, tuft surveys, and schlieren photographs—resulted in a new theoretical understanding of transonic drag. The schlieren photo­graphs revealed three shock waves at transonic speeds. One was the famil­iar shock wave that formed at the nose of an aircraft as it pushed forward through the air. The other two were, according to Whitcomb, "fascinating new types” of shock waves never before observed, in which the fuselage and wings met and at the trailing edge of the wing. These shocks contributed to a new understanding that transonic drag was much larger in proportion to the size of the fuselage and wing than previously believed. Whitcomb speculated that these new shock waves were the cause of transonic drag.[151]