Category AERONAUTICS

As the computational capability to accurately model the handling quali­ties of an airplane improved, there was recognition that the lack of motion cues was a distraction to the realism of the simulation. An early attempt to simulate motion for the pilot consisted of mounting the entire simula­tor cockpit on a set of large hydraulic actuators. These actuators would generate a small positive or negative bump to simulate g onset, while any steady-state acceleration was washed out over time (i. e., back to 1 g). The actuators could also tilt the simulator cockpit to produce a side force, or fore and aft force, on the crew. When correlated with a horizon on a visual screen, the result was a quite realistic sensation of motion. These moving-base cockpit systems were rather expensive and difficult to main­tain compared with a simple fixed-base cockpit. Since both the magni­tude of the g vector and the rotational motion required were false, the systems were not widely accepted in the flight-testing community, where the goal is to evaluate the pilot’s response and capabilities in a true flight environment. They found ready acceptance as airline procedures train­ers when the maneuvers are slow and g forces are typically small and proved a source of entertainment in amusement parks, aerospace muse­ums, and science centers.

In the 1950s, the Naval Air Development Center (NADC) at Johnsville, PA, developed a large, powerful centrifuge to explore human tolerance to high g forces. The centrifuge consisted of a 182-ton electric DC motor turning a 50-foot arm with a gondola at the end of the arm. The motor could generate g forces at the gondola as high as 40 g’s. The gondola was mounted with two controllable gimbals that allowed the g vector to be oriented in different directions for the gondola occupant.[728]

Many detailed studies defining human tolerance to g forces were per­formed on the centrifuge using programmed g profiles. NADC devised a method for installing a cockpit in the gondola, connecting it to a large analog computer, and allowing the pilot to control the computer sim­ulation, which in turn controlled the centrifuge rotation rate and gim – bal angles. This allowed the pilot in the gondola to not only see the pilot displays of the simulated flight, but also to feel the associated transla­tional g levels in all three axes. Although the translational g forces were correctly simulated, the gimbal rotations necessary to properly align the total g vector with the cockpit were artificial and were not representa­tive of a flight environment.

One of the first applications of this closed-loop, moving base sim­ulation was in support of the X-15 program in 1958. There were two prime objectives of the X-15 centrifuge program associated with the high g exit and entry: assessment and validation of the crew station (side arm controller, head and arm restraints, displays, etc.), and evaluation of the handling qualities with and without the Stability Augmentation System. The g environment during exit consisted of a forward acceler­ation (eyeballs-in) increasing from 2 to 4 g, combined with a 2 g pullup (eyeballs-down). The entry g environment was more severe, consisting

of a deceleration (eyeballs-out) of 3 g combined with a simultaneous pullout acceleration of 6 g (eyeballs-down).

The results of the X-15 centrifuge program were quite useful to the X-15’s overall development; however, the pilots felt that the centrifuge did not provide a very realistic simulation of an aircraft flight environment. The false rotational movement of the gondola was apparent to the pilots and was a distraction to the piloting task during entry. T he exit phase of an X-15 flight was a fairly steady acceleration with little rotational motion, and the pilots judged the simulation a good representation of that environment.[729]

The NADC centrifuge was also used in support of the launch phase of the Mercury, Gemini, and Apollo space programs. These provided valuable information regarding the physiological condition of the astronauts and the crew station design but generally did not include closed-loop piloting tasks with the pilot controlling the simulated vehicle and trajectory.

A second closed-loop centrifuge simulation was performed in sup­port of the Boeing X-20 Dyna-Soar program. Dyna-Soar constituted an ambitious but feasible Air Force effort to develop a hypersonic lofted boost-glider capable of an orbital flight. Unfortunately, it was prematurely canceled in 1963 by then-Secretary of Defense Robert S. McNamara. The Dyna-Soar centrifuge study effort was similar to the X-15 centrifuge program, but the acceleration lasted considerably longer and peaked at 6 g (eyeballs-in) at burnout of the Titan III booster. The pilots were "flying” the vehicle in all three axes during these centrifuge runs, and valuable data were obtained relative to the pilot’s ability to function effec­tively during long periods of acceleration. Some of the piloting demon­strations included alleviating wind spikes during the early ascent phase and successfully guiding the booster to an accurate orbital insertion using simple backup guidance concepts in the event of a booster guid­ance failure.[730] The Mercury and Gemini programs used automatic guid­ance during the ascent phase, and the only piloting task during boost was to initiate an abort by firing the escape rockets. The Apollo program included a backup piloting mode during the boost based on the results of the X-20 and other centrifuge programs.

What constitutes computational fluid dynamics? The basic equations of fluid dynamics, the Navier-Stokes equations, are expressions of three fundamental principles: (1) mass is conserved (the continuity equation), (2) Newton’s second law (the momentum equation), and (3) the energy equation (the first law of thermodynamics). Moreover, these equations in their most general form are either partial differential equations (as we have discussed) or integral equations (an alternate form we have not discussed involving integrals from calculus).

The partial differential equations are those exhibited at the NASM. Computational fluid dynamics is the art and science of replacing the partial derivatives (or integrals, as the case may be) in these equations with discrete algebraic forms, which in turn are solved to obtain num­bers for the flow-field values (pressure, density, velocity, etc.) at discrete points in time and or space.[765] At these selected points in the flow, called grid points, each of the derivatives in each of the equations are simply replaced with numbers that are advanced in time or space to obtain a solution for the flow. In this fashion, the partial differential equations are replaced by a large number of algebraic equations, which can then be solved simultaneously for the flow variables at all the grid points.

The end product of the CFD process is thus a collection of numbers, in contrast to a closed-form analytical solution (equations). However, in the long run, the objective of most engineering analyses, closed-form or otherwise, is a quantitative description of the problem: that is, num­bers. Along these lines, in 1856, the famous British scientist James Clerk Maxwell wrote: "All the mathematical sciences are founded on relations between physical laws and laws of numbers, so that the aim of exact
science is to reduce the problems of nature to the determination of quantities by operations with numbers.”[766] Well over a century later, it is worth noting how well Maxwell captured the essence of CFD: operations with numbers.

Note that computational fluid dynamics results in solutions for the flow only at the distinct points in the flow called grid points, which were identified earlier. In a CFD solution, grid points are either initially dis­tributed throughout the flow and/or generated during the course of the solution (called an "adaptive grid”). This is in theoretical contrast with a closed-form analytical solution for the flow, where the solution is in the form of equations that allow the calculation of the flow variables at any point of one’s choosing, that is, an analytical solution is like a con­tinuous answer spread over the whole flow field. Closed-form analyti­cal solutions may be likened to a traditionalist Dutch master’s painting consisting of continuous brush strokes, while a CFD solution is akin to a French pointillist consisting of multicolored dots made with a brush tip.

Generating a grid is an essential part of the art of CFD. The spacing between grid points and the geometric ways in which they are arrayed is critical to obtaining an accurate numerical CFD solution. Poor grids almost always ensure poor CFD solutions. Though good grids do not guarantee good CFD solutions, they are essential for useful solutions. Grid genera­tion is a discipline all by itself, a subspecialty of CFD. And grid generation can become very labor-intensive—for some flows over complex three­dimensional configurations, it may take months to generate a proper grid.

To summarize, the Navier-Stokes equations, the governing equations of fluid dynamics, have been in existence for more than 160 years, their creation a triumph of derivative insight. But few knew how to analyti­cally solve them except for a few simple cases. Because of their complex­ity, they thus could not serve as a practical widely employed tool in the engineer’s arsenal. It took the invention of the computer to make that pos­sible. And because it did so, it likewise permitted the advent of computa­tional fluid dynamics. So how did the idea of numerical solutions to the Navier-Stokes equations evolve?

The first components of NASTRAN became operational at Goddard in May 1968. Distribution to other Centers, training, and a debugging period followed through 1969 and into 1970.[816] With completion of the initial devel­opment, "the management of NASTRAN was transferred to the Langley Research Center. The NASTRAN Systems Management Office (NSMO) was established in the Structures Division at Langley October 4, 1970.”[817] Initial public release followed just 1 month later, in November 1970.

NSMO responsibilities included:[818]

• Centralized program development (advisory committees).

• Coordinating user experiences (bimonthly NASTRAN Bulletin and annual Users’ Colloquia).

• System maintenance (error correction and essential improvements).

• Development and addition of new capability.

• NASTRAN-focused research and development (R&D).

The actual distribution of NASTRAN to the public was handled by the Computer Software Management and Information Center (COSMIC), NASA’s clearinghouse for software distribution (which is described in a subsequent section of this paper). The price at initial release was $1,700, "which covers reproducing and supplying the necessary system tapes and documentation.”[819] Documentation was published in four vol­umes, each with a distinct purpose: one for users, one for programmers who would be involved in maintenance and subsequent development, a theory manual, and finally a volume of demonstration problems. (The 900-page user’s manual could be obtained from COSMIC for $10, if purchased separately from the program itself. The author assumes that the other volumes were similarly priced.)[820]

CATEGORY	COMPUTERS
NASA CENTERS
DOC
OTHER GOV T
8
AEROSPACE
AIRCRAFT
COMPUTER CORF	MANUFACTURING ENGR CONSIT AUTO UNIVERSITIES
OTHERS

2,272 USERS

NASTRAN user community profile in 1974. NASA.

Things were happening quickly. Within the first year after public release, NASTRAN was installed on over 60 machines across the United States. There were urgent needs requiring immediate attention. "When NSMO was established in October 1970, there existed a dire need for maintenance of the NASTRAN system. With the cooperation of Goddard Space Flight Center, an interim maintenance contract was negotiated with Computer Sciences Corporation through a contract in effect at GSFC. This contract provided for the essential function of error cor­rection until a contract for full time maintenance could be negotiated through an open competition. The interim maintenance activity was restricted to the correction of over 75 errors reported to the NSMO, together with all associated documentation changes. New thermal bend­ing and hydroelastic elements previously developed by the MacNeal- Schwendler Corporation under contract to GSFC were also installed. Levels 13 and 14 were created for government testing and evaluation. The next version of NASTRAN to be released to the public. . . will be built upon the results of this interim maintenance activity and will be designated Level 15,” according to a status report to the user community in 1971.[821]

In June 1971, the contract for full-time maintenance was awarded to MacNeal Schwendler Corporation, which then opened an office near Langley. A bug reporting and correction system was established. Bell Aerospace Company received a contract to develop new elements and a thermal analysis capability. Other efforts were underway to improve efficiency and execution time. A prioritized list of future upgrades was started, with input from all of the NASA Centers. However, for the time being, the pace of adding new capability would be limited by the need to also keep up with essential maintenance.[822]

By 1975, NASTRAN was installed on 269 computers. The estimated composition of the user community (based on a survey taken by the NSMO) is illustrated here.

By this time, the NSMO was feeling the pressure of trying to keep up with maintenance, bug fixes, and requested upgrades from a large and rapidly growing user community. There was also a need to keep up with changing hardware technology. Measures under consideration included improvements to the Error Correction Information System (ECIS); more user involvement in the development of improvements (although this would also require effort to enforce coding standards and interface requirements, and to develop procedures for verification and implementation); and a price increase to help support the NSMO’s maintenance costs and also possibly recoup some of the NASTRAN development costs. COSMIC eventually changed its terms for all soft­ware distribution to help offset the costs of maintenance.

An annual NASTRAN Users’ Colloquium was initiated, the first of which occurred approximately 1 year after initial public release. Each Colloquium usually began with an overview from the NSMO on NASTRAN status, including usage trends, what to expect in the next release, and planned changes in NASTRAN management or terms of distribution. Other papers covered experiences and lessons learned in deployment, installation, and training; technical presentations on new types of elements or new solution capabilities that had recently been, were being, or could be, implemented; evaluation and comparison of NASTRAN with test data or other analysis methods; and user experiences and applications. (The early NASTRAN Users’ Colloquia proceedings were available from the National Technical Information Service for $6.)

The first Colloquia were held at Langley and hosted by the NSMO staff. As the routine became more established and the user community grew, the Colloquia were moved to different Centers and cochaired, usu­ally by the current NSMO Manager and a representative from the host­ing Center. There were 21 Users’ Colloquia, at which 429 papers were presented. The breakdown of papers by contributing organization is shown here.(Note: collaborative papers are counted under each contrib­uting organization, so the sum of the subtotals exceeds the overall total.)

Computing companies were typically involved in theory, modeling technique, resolution of operational issues, and capability improvements (sometimes on contracts to NASA or other agencies), but also collab­orated with "user” organizations assisting with NASTRAN application to problems of interest. All participants were actively involved in the improvement of NASTRAN, as well as its application.

Major aircraft companies—Boeing, General Dynamics, Grumman, Lockheed, McDonnell-Douglas, Northrop, Rockwell, and Vought—were frequent participants, presenting a total of 70 papers. Smaller aerospace companies also began to use NASTRAN. Gates Learjet modeled the Lear 35/36 wing as a test case in 1976 and then used NASTRAN in the design phase of the Lear 28/29 and Lear 55 business jets.[823] Beechcraft used NASTRAN in the design of the Super King Air 200 twin turboprop and the T-34C Mentor military trainer.[824] Dynamic Engineering, Inc., (DEI) began using NASTRAN in the design and analysis of wind tunnel models in the 1980s.[825]

Nonaerospace applications appeared almost immediately. By 1972, NASTRAN was being used in the automotive industry, in architectural engineering, by the Department of Transportation, and by the Atomic Energy Commission. The NSMO had "received expressions of interest in NASTRAN from firms in nearly every West European country, Japan, and Israel.”[826] That same year, "NASTRAN was chosen as the principal analytical tool” in the design and construction of the 40-story Illinois Center Plaza Hotel building.[827]

Other nonaerospace applications included:

• Nuclear power plants.

• Automotive industry, including tires as well as primary structure.

• Ships and submarines.

• Munitions.

• Acoustic and electromagnetic applications.

• Chemical processing plants.

• Steam turbines and gas turbines.

• Marine structures.

• Electronic circuitry.

B. F. Goodrich, General Motors, Tennessee Eastman, and Texas Instruments were common presenters at the Colloquia. Frequent Government participants, apart from the NASA Centers, included the David Taylor Naval Ship Research & Development Center, the Naval Underwater Systems Cener, the U. S. Army Armament Research & Development Command, and several U. S. Army arsenals and laboratories.[828]

Technical improvements, too numerous to describe them all, were continually being made. At introduction (Level 12), NASTRAN offered linear static and dynamic analysis. There were two main classes of new capability: analysis routines and structural elements. Developments were often tried out on an experimental basis by users and reported on at the Colloquia before being incorporated into standard NASTRAN. Evaluations and further improvements to the capability would typically follow. In addi­tion, of course, there were bug fixes and operational improvements. A few key developments are identified below. Where dates are given, they rep­resent the initial introduction of a capability into standard NASTRAN, not to imply full maturity:

• Thermal analysis: initial capability introduced at Level 15 (1973).

• Pre – and post-processing: continuous.

• Performance improvements: continuous.

• Adaptation to new platforms and operating systems: con­tinuous. (The earliest mention the author has found of NASTRAN running on a PC is 1992.[829])

• New elements: continuous. Level 15 included a dummy structural element to facilitate user experimentation.

• Substructuring: the decomposition of a larger model into smaller models that could be constructed, manipulated, and/or analyzed independently. It was identified as an important need when NASTRAN was first introduced.

Initial substructuring capability was introduced at Level 15 in 1973.

• Aeroelastics and flutter: studies were conducted in the early 1970s. Initial capability was introduced in Level 16 by 1976. NASTRAN aeroelastic, flutter, and gust load anal­ysis uses a doublet-lattice aerodynamic method, which approximates the wing as an initially flat surface for the aerodynamic calculation (does not include camber or thickness). The calculation is much simpler than full – fledged computational fluid dynamics (CFD) analysis but neglects many real flow effects as well as configuration geometry details. Accuracy is provided by using correc­tion factors to match the static characteristics of the dou­blet-lattice model to higher fidelity data from flight test, wind tunnel test, and/or CFD. One of the classic references on correcting lifting surface predictions is a paper by J. P. Giesing, T. P. Kalman, and W. P. Rodden of McDonnell – Douglas, on contract to NASA, in 1976.[830]

• Automated design and analysis: automated fully stressed design was introduced in Level 16 (1976). Design automa­tion is a much broader field than this, however, and most attempts to further automate design, analysis, and/or opti­mization have taken the form of applications outside of, and interfacing with, NASTRAN. In many industries, auto­mated design has become routine; in others, the status of automated design remains largely experimental, primarily because of the inherent complexity of design problems.[831]

• Nonlinear problems: geometric and/or material. Geometric nonlinearity is introduced, for example, when displacements are large enough to change the geomet­ric configuration of the structure in significant ways. Material nonlinearity occurs when local stresses exceed the linear elastic limit. Applications of nonlinear analy­sis include engine hot section parts experiencing regions of local plasticity, automotive and aircraft crash simula­tion, and lightweight space structures that may experi­ence large elastic deformations—to name a few. Studies and experimental implementations were made dur­ing the 1970s. There are many different classes of non­linear problems encompassed in this category, requir­ing a variety of solutions, many of which were added to standard NASTRAN through the 1980s.

Note: This appendix includes a list of the dates and locations of the NASTRAN Users’ Colloquia and NASTRAN applications presented at the Colloquia by "nontraditional” users, i. e., industry other than aero­space, Government agencies other than NASA, and universities. Not every paper from these sources is listed, only those that represent applications. Many other papers were presented on modeling techniques, capability improvements, etc., which are not listed.

NASTRAN Reference Sources

At time of writing, these are available from the NASA Technical Reports Server at http://ntrs. nasa. gov:

NASTRAN: Users’ Experiences, NASA TM-X-2378, 1971.

NASTRAN: Users’ Experiences (2nd), NASA TM-X-2637, 1972.

NASTRAN: Users’ Experiences (4th), NASA TM-X-3278, 1975.

NASTRAN: Users’ Experiences (5th), NASA TM-X-3428, 1976.^^^^^^ Sixth NASTRAN Users’ Colloquium, NASA CP-2018, 1977.

Seventh NASTRAN Users’ Colloquium, NASA CP-2062, 1978.

Eight NASTRAN Users’ Colloquium, NASA CP-2131, 1979.

Ninth NASTRAN Users’ Colloquium, NASA CP-2151, 1980.

Tenth NASTRAN Users’ Colloquium, NASA CP-2249, 1982.

Eleventh NASTRAN Users’ Colloquium, NASA CP-2284, 1983.

Twelfth NASTRAN Users’ Colloquium, NASA CP-2328, 1984.

Thirteenth NASTRAN Users’ Colloquium, NASA CP-2373, 1985.

Fourteenth NASTRAN Users’ Colloquium, NASA CP-2419, 1986.

Fifteenth NASTRAN Users’ Colloquium, NASA CP-2481, 1987.

Sixteenth NASTRAN Users’ Colloquium, NASA CP-2505, 1988.

Seventeenth NASTRAN Users’ Colloquium, NASA CP-3029, 1989.

Eighteenth NASTRAN Users’ Colloquium, NASA CP-3069, 1990.

Nineteenth NASTRAN Users’ Colloquium, NASA CP-3111, 1991.

Twentieth NASTRAN Users’ Colloquium, NASA CP-3145, 1992.

Twenty-First NASTRAN Users’ Colloquium, NASA CP-3203, 1993.

Appendix B:

Atmosphere entry of satellites takes place above Mach 20, only slightly faster than the speed of reentry of an ICBM nose cone. The two phe­nomena nevertheless are quite different. A nose cone slams back at a sharp angle, decelerating rapidly and encountering heating that is brief but very severe. Entry of a satellite is far easier, taking place over a number of minutes.

To learn more about nose cone reentry, one begins by considering the shape of a nose cone. Such a vehicle initially has high kinetic energy because of its speed. Following entry, as it approaches the ground, its kinetic energy is very low. Where has it gone? It has turned into heat, which has been transferred both into the nose cone and into the air that has been disturbed by passage of the nose cone. It is obviously of inter­est to transfer as much heat as possible into the surrounding air. During reentry, the nose cone interacts with this air through its bow shock. For effective heat transfer into the air, the shock must be very strong. Hence the nose cone cannot be sharp like a church steeple, for that would substantially weaken the shock. Instead, it must be blunt, as H. Julian Allen of the National Advisory Committee for Aeronautics (NACA) first recognized in 1951.[1030]

Now that we have this basic shape, we can consider methods for cooling. At the outset of the Atlas ICBM program, in 1953, the sim­plest method of cooling was the heat sink, with a thick copper shield absorbing the heat of reentry. An alternative approach, the hot struc­ture, called for an outer covering of heat-resistant shingles that were to radiate away the heat. A layer of insulation, inside the shingles, was to protect the primary structure. The shingles, in turn, overlapped and could expand freely.

A third approach, transpiration cooling, sought to take advantage of the light weight and high heat capacity of boiling water. The nose cone was to be filled with this liquid; strong g-forces during deceleration in the atmosphere were to press the water against the hot inner skin. The skin was to be porous, with internal steam pressure forcing the fluid

through the pores and into the boundary layer. Once injected, steam was to carry away heat. It would also thicken the boundary layer, reducing its temperature gradient and hence its rate of heat transfer. In effect, the nose cone was to stay cool by sweating.

Still, each of these approaches held difficulties. Transpiration cooling was poorly understood as a topic for design. The hot-structure concept raised questions of suitably refractory metals along with the prospect
of losing the entire nose cone if a shingle came off. Heat sinks appeared to promise high weight. But they seemed the most feasible way to proceed, and early Atlas designs specified use of a heat-sink nose cone.[1031]

Atlas was an Air Force program. A separate set of investigations was underway within the Army, which supported hot structures but raised problems with both heat sink and transpiration. This work antic­ipated the independent studies of General Electric’s George Sutton, with both efforts introducing an important new method of cooling: ablation. Ablation amounted to having a nose cone lose mass by flaking off when hot. Such a heat shield could absorb energy through latent heat, when melting or evaporating, and through sensible heat, with its tempera­ture rise. In addition, an outward flow of ablating volatiles thickened the boundary layer, which diminished the heat flow. Ablation promised all the advantages of transpiration cooling, within a system that could be considerably lighter and yet more capable, and that used no fluid.[1032]

Though ablation proved to offer a key to nose cone reentry, experi­ments showed that little if any ablation was to be expected under the rel­atively mild conditions of satellite entry. But satellite entry involved high total heat input, while its prolonged duration imposed a new require­ment for good materials properties as insulators. They also had to stay cool through radiation. It thus became possible to critique the useful­ness of ICBM nose cone ablators for the new role of satellite entry.[1033]

Heat of ablation, in British thermal units (BTU) per pound, had been a standard figure of merit. Water, for instance, absorbs nearly 1,000 BTU/lb when it vaporizes as steam at 212 °F. But for satellite entry, with little energy being carried away by ablation, head of ablation could be irrelevant. Phenolic glass, a fine ICBM material with a measured heat of 9,600 BTU/lb, was unusable for a satellite because it had an unac­ceptably high thermal conductivity. This meant that the prolonged ther­mal soak of a satellite entry could have enough time to fry a spacecraft.

Teflon, by contrast, had a measured heat only one-third as large. It nevertheless made a superb candidate because of its excellent proper­ties as an insulator.[1034]

Hence it became possible to treat the satellite problem as an exten­sion of the ICBM problem. With appropriate caveats, the experience and research techniques of the ICBM program could carry over to this new realm. The Central Intelligence Agency was preparing to recover satellite spacecraft at the same time that the Air Force was preparing to fly full-size Atlas nose cones, with both being achieved in April 1959.

The Army flew a subscale nose cone to intermediate range in August 1958, which President Dwight Eisenhower displayed during a November news conference. The Air Force became the first to fly a nose cone to intercontinental range, in July 1958. Both flights carried a mouse, and both mice survived their reentry, but neither was recovered. Better success came the following April, when an Atlas launched the full – size RVX-l nose cone, and the Discoverer II reconnaissance spacecraft returned safely through the atmosphere—though it fell into Russian, not American, hands.[1035]

By the late 1960s, several European research aircraft using partial fly-by-wire flight control systems were in development. In Germany, the supersonic VJ-101 experimental Vertical Take-Off and Landing fighter technology demonstrator, with its swiveling wingtip mounted after­burning turbojet engines, and the Dornier Do-31 VTOL jet transport used analog computer-controlled partial fly-by-wire flight control sys­tems. American test pilots were intimately involved with both programs. George W. Bright flew the VJ-101 on its first flight in 1963, and NASA test
pilot Drury W. Wood, Jr., headed the cooperative U. S.-German Do-31 flight-test program that included representatives from NASA Langley and NASA Ames. Wood flew the Do-31 on its first flight in February 1967. He received the Society of Experimental Test Pilots’ Ivan C. Kinchloe Award in 1968 for his role on the Do-31 program.[1142] By that time, NASA test pilot Robert Innis was chief test pilot on the Do-31 program. The German VAK-191B VTOL fighter technology flight demonstrator flew in 1971. Its triply redundant analog flight control system assisted the pilot in operating its flight control surfaces, engines, and reaction control nozzles, but the aircraft retained a mechanical backup capability. Later in its flight-test program, the VAK-191B was used to support development of the partial fly-by-wire flight control system that was used in the multina­tional Tornado multirole combat aircraft that first flew in August 1974.[1143]

In the U. K., a Hawker Hunter T.12 two-seat jet trainer was con­verted into a fly-by-wire testbed by the Royal Aircraft Establishment. It incorporated a three-axis, quadruplex analog Integrated Flight Control System (IFCS) and a "sidearm” controller. The mechanical backup flight control system was retained.[1144] First flown in April 1972, the Hunter was eventually lost in a takeoff accident.

In the USSR, a Sukhoi Su-7U two-seat jet fighter trainer was mod­ified with forward destabilizing canards as the Projekt 100LDU fly-by­wire testbed. It first flew in 1968 in support of the Sukhoi T-4 supersonic bomber development effort. Fitted with a quadruple redundant fly-by­wire flight control system with a mechanical backup capability, the four – engine Soviet Sukhoi T-4 prototype first flew in August 1972. Reportedly, the fly-by-wire flight control system provided much better handling qual­ities than the T-4’s mechanical backup system. Four T-4 prototypes were built, but only the first aircraft ever flew. Designed for Mach 3.0, the T-4 never reached Mach 2.0 before the program was canceled after only 10 test flights and about 10 hours of flying time.[1145] In 1973-1974, the Projekt
100LDU testbed was used to support development of the fly-by-wire sys­tem flight control system for the Sukhoi T-10 supersonic fighter proto­type program. The T-10 was the first pure Soviet fly-by-wire aircraft with no mechanical backup; it first flew on May 27, 1977. On July 7, 1978, the T-10-2 (second prototype) entered a rapidly divergent pitch oscilla­tion at supersonic speed. Yevgeny Solovyev, distinguished test pilot and hero of the Soviet Union, had no chance to eject before the aircraft dis­integrated.[1146] In addition to a design problem in the flight control system, the T-10’s aerodynamic configuration was found to be incapable of pro­viding required longitudinal, lateral, and directional stability under all flight conditions. After major redesign, the T-10 evolved into the highly capable Sukhoi Su-27 family of supersonic fighters and attack aircraft.[1147]

NASA observations on some of the more serious issues encountered in early testing of the AFTI/F-16 asynchronous digital flight control sys­tem are worthy of note. For example, an unknown failure in the Stores Management System on flight No. 15 caused it to request DFCS mode changes at a rate of 50 times per second. The DFCS could not keep up and responded at a rate of 5 mode changes per second. The pilot reported that the aircraft felt like it was in severe turbulence. The flight was aborted, and the aircraft landed safely. Subsequent analysis showed that if the aircraft had been maneuvering at the time, the DFCS would have failed. A subsequent software modification improved the DFCS’s immunity to this failure mode.[1202]

A highly significant flight control law anomaly was encountered on AFTI/F-16 flight No. 36. Following a planned maximum rudder "step and hold” input by the pilot, a 3-second departure from controlled flight occurred. Sideslip angle exceeded 20 degrees, normal acceleration fluc­tuated from -4 g to +7 g, angle of attack varied between -10 and +20 degrees, and the aircraft rolled 360 degrees. Severe structural loads were encountered with the vertical tailfin exceeding its design load. During the out-of-control situation, all control surfaces were operating at rate limits, and failure indications were received from the hydraulics and canard actuators. The failures were transient and reset after the pilot regained control. The problem was traced to a fault in the programmed flight control laws. It was determined that the aerodynamic model used to develop the control laws did not accurately model the nonlinear nature of yaw stability variations as a function of higher sideslip angles. The same inaccurate control laws were also used in the real-time AFTI/F-16 ground flight simulator. An additional complication was caused when the side fuselage-mounted air-data probes were blanked by the canard
at the high angles of attack and sideslip encountered. This resulted in incorrect air data values being passed to the DFCS. Operating asynchro­nously, the different flight control system channels took different paths through the flight control laws. Analysis showed these faults could have caused complete failure of the DFCS and reversion to analog backup.[1203] Subsequently, the canards were removed from the command path to prevent the AFTI/F-16 from obtaining higher yaw angles.

AFTI/F-16 flight-testing revealed numerous other flight control prob­lems of a similar nature. These prompted NASA engineer Dale Mackall to report: "The asynchronous design of the [AFTI/F-16] DFCS introduced a random, unpredictable characteristic into the system. The system became untestable in that testing for each of the possible time relation­ships between the computers was impossible. This random time rela­tionship was a major contributor to the flight test anomalies. Adversely affecting testability and having only postulated benefits, asynchronous operation of the DFCS demonstrated the need to avoid random, unpre­dictable, and uncompensated design characteristics.” Mackall also pro­vided additional observations that would prove to be highly valuable in developing, validating, and certifying future software-intensive digital fly-by-wire flight control system designs. Urging more formal approaches and rigorous control over the flight control system software design and development process, Mackall reported:

The criticality and number of anomalies discovered in flight and ground tests owing to design oversights are more significant than those anomalies caused by actual hardware failures or software errors. . . . As the operational requirements of avionics systems increase, complexity increases. . . . If the complexity is required, a method to make system designs more understandable, more visible, is needed. . . qualification of such a complex system as this, to some given level of reliability, is difficult. . . the number of test conditions becomes so large that conventional testing methods would require a decade for completion. The fault – tolerant design can also affect overall system reliability by being made too complex and by adding characteristics which are ran­dom in nature, creating an untestable design.[1204]

NASA Dryden used an NF-15B research aircraft on various research proj­ects from 1993 through early 2009. Originally designated the TF-15, it was the first two-seat F-15 Eagle built by McDonnell-Douglas, the sixth F-15 off the assembly line, and the oldest F-15 flying up to its retire­ment. First flown in July 1973, the aircraft was initially used for F-15 developmental testing and evaluation as part of the F-15 combined test force at Edwards AFB in the 1970s. In the 1980s, the aircraft was exten­sively modified for the Air Force’s Short Takeoff and Landing Maneuver Technology Demonstrator (S/MTD) program. Modifications included the integration of a digital fly-by-wire control system, canards mounted on the engine inlets ahead of the wings,[1273] and two-dimensional thrust­vectoring, thrust-reversing nozzles. The vectoring nozzles redirected

engine exhaust either up or down, giving greater pitch control and addi­tional aerodynamic braking capability. Designated NF-15B to reflect its status as a highly modified research aircraft, the aircraft was used in the S/MTD program from 1988 until 1993. During Air Force S/MTD testing, a 25-percent reduction in takeoff roll was demonstrated with thrust-reversing, enabling the aircraft to stop in just 1,650 feet. Takeoffs using thrust-vectoring produced nose rotation speeds as low as 40 knots, resulting in greatly reduced takeoff distances. Additionally, thrust-revers­ing produced extremely rapid in-flight decelerations, a feature valuable during close-in combat.[1274]

It’s a noisy world out there, especially around the Nation’s busiest air­ports, so NASA is pioneering new technologies and aircraft designs that could help quiet things down a bit. Every source of aircraft noise, from takeoff to touchdown, is being studied for ways to reduce the racket, which is expected to get worse as officials predict that air traffic will double in the next decade or so.

"It’s always too noisy. You have to always work on making it quieter,” said Edmane Envia, an aerospace engineer at NASA’s Glenn Research Center in Cleveland. "You always have to stay a step ahead to fulfill the needs and demands of the next generation of air travel.”[1366]

Noise reduction research is part of a broader effort by NASA’s Aeronautics Research Mission Directorate in Washington to lay a tech­nological foundation for a new generation of airplanes that are not as noisy, fly farther on less fuel, and may operate out of airports with much shorter runways than exist today. There are no clear solutions yet to these tough challenges, neither is there a shortage of ideas from NASA researchers who are confident positive results eventually will come.[1367]

"Our goal is to have the technologies researched and ready, but ulti­mately it’s the aircraft industry, driven by the market, that makes the deci­sion when to introduce a particular generation of aircraft,” Envia said.

NASA organized its research to look three generations into the future, with conceptual aircraft designs that could be introduced 10, 20, or 30 years from now. The generations are called N+1, N+2, and N+3. Each generation represents a design intended to be flown a decade or so later than the one before it and is to feature increasingly sophisticated meth­ods for delivering quieter aircraft and jet engines.[1368]

"Think of the Boeing 787 Dreamliner as N and the N+1 as the next generation aircraft after that,” Envia said.

The N+1 is an aircraft with familiar parts, including a conventional tube-shaped body, wings, and a tail. Its jet engines still are attached to the wings, as with an N aircraft, but those engines might be on top of the wings, not underneath. Conceptual N+2 designs throw out con­vention and basically begin with a blank computer screen, with design engineers blending the line between the body, wing, and engines into a more seamless, hybrid look. What an N+3 aircraft might look like is anyone’s guess right now. But with its debut still 30 years away, NASA is sponsoring research that will produce a host of ideas for consid­eration. The Federal Aviation Administration’s current guidelines for overall aircraft noise footprints constitute the design baseline for all of NASA’s N aircraft concepts. That footprint summarizes in a single number, expressed as a decibel, the noise heard on the ground as an airplane lands, takes off, and then cuts back on power for noise abate­ment. The noise footprint extends ahead and behind the aircraft and to a certain distance on either side. NASA’s design goal is to make each new aircraft generation quieter than today’s airplanes by a set number of decibels. The N+1 goal is 32 decibels quieter than a fully noise compliant Boeing 737, while the N+2 goal is 42 decibels quieter than a Boeing 777. So far, the decibel goal for the N+1 aircraft has been elusive.[1369]

"What makes our job very hard is that we are asked to reduce noise but in ways that do not adversely impact how high, far or fast an air­plane is capable of flying,” Envia said.

NASA researchers have studied changes in the operation, shape, or materials from which key noise contributors are made. The known suspects include the airframe, wing flaps, and slats, along with components of the jet engine, such as the fan, turbine, and exhaust noz­zle. While some reductions in noise can be realized with some design changes in these components, the overall impact still falls short of the N+1 goal by about 6 decibels. Envia said that additional work with design and operation of the jet engine’s core may make up the difference, but that a lot more work needs to be done in the years to come. Meanwhile, reaching the N+2 goals may or may not prove easier to achieve.[1370]

"We’re starting from a different aircraft configuration, from a clean sheet, that gives you the promise of achieving even more aggressive goals,” said Russell Thomas, an aerospace engineer at Langley Research Center. "But it also means that a lot of your prior experience is not directly appli­cable, so the problem gets a lot harder from that point of view. You may have to investigate new areas that have not been researched heavily in the past.”[1371]

Efforts to reduce noise in the N+2 aircraft have focused on the air­frame, which blends the wing and fuselage together, greatly reducing the number of parts that extend into the airflow to cause noise. Also, according to Thomas, the early thinking on the N+2 aircraft is that the jet engines will be on top of the vehicle, using the airplane body to shield most of the noise from reaching the ground.

"We’re on course to do much more thorough research to get higher quality numbers, better experiments, and better prediction methods so we can really understand the acoustics of this new aircraft configura­tion,” Thomas said.

As for the N+3 aircraft, it remains too early to say how NASA researchers will use technology not yet invented to reduce noise levels to their lowest ever.

"Clearly significant progress has been made over the years and air­planes are much quieter than they were 20 years ago,” Envia said, not­ing that further reductions in noise will require whole new approaches to aircraft design. "It is a complicated problem and so it is a worthy challenge to rise up to.”

The Mod-0 testbed wind turbine system was upgraded from 100 kilo­watts to a 200-kilowatt system that became the Mod-0A. Installation of the first Mod-0A system was completed in November 1977, with one additional machine installed each year through 1980 at four locations: Clayton, NM; Culebra, PR; Block Island, RI; and Oahu, HI. This first generation of wind turbines completed its planned experimental oper­ations in 1982 and was removed from service.

The basic components and systems of the Mod-0A consisted of the rotor – and pitch-change mechanism, drive train, nacelle equipment, yaw drive mechanism and brake, tower and foundation, electrical sys­tem and components, and control systems. The rotor consisted of the blades, hub, pitch-change mechanism, and hydraulic system. The drive train included the low-speed shaft, speed increaser, high-speed shaft, belt drive, fluid coupling, and rotor blades. The electrical system and components were the generator, switchgear, transformer, utility con­nection, and slip rings. The control systems were the blade pitch, yaw, generator control, and safety system.11 [1502]

Similar to the Mod-0 testbed, the Mod-0A horizontal-axis machines had a 125-foot-diameter downwind rotor mounted on a 100-foot rigid pinned truss tower. However, this more powerful first genera­tion of turbines had a rated power of 200 kilowatts at a wind speed of 18 miles per hour and made 40 revolutions per minute. The turbine had two aluminum blades that were each 59.9 feet long. The Westinghouse Electric Corporation was selected, by competitive bidding, as the contractor for building the Mod-0A, and Lockheed was selected to design and build the blades. NASA and Westinghouse personnel were involved in the installation, site tests, and checkout of the wind turbine systems.

The primary goal of the Mod-0A wind turbine was to gain expe­rience and obtain early operation performance data with horizontal – axis wind turbines in power utility environments, including resolving issues relating to power generation quality, and safety, and procedures for system startup, synchronization, and shutdown. This goal included demonstrating automatic operation of the turbine and assessing machine compatibility with utility power systems, as well as determining reliability and maintenance requirements. To accomplish this primary goal, small power utility companies or remote location sites were selected in order to study problems that might result from a significant percentage of power input into a power grid. NASA engineers also wanted to determine the reaction of the public and power utility companies to the operation of the turbines. The Mod-0A systems were online collectively for over 38,000 hours, generating over 3,600 megawatthours of electricity into power utility networks. NASA deter­mined that while some early reliability and rotor-blade life problems needed to be corrected, overall the Mod-0A wind turbine systems accomplished the engineering and research objectives of this phase of the program and made significant contributions to second – and third-generation machines that were to follow the Mod-0A and Mod-1 projects. Interface of the Mod-0A with the power utili­ties demonstrated satisfactory operating results during their ini­tial tests from November 1977 to March 1978. The wind turbine was successfully synchronized to the utility network in an unattended mode. Also, dynamic blade loads during the initial operating period were in good agreement with the calculation using the MOSTAB computer code. Finally, successful testing on the Mod-0 provided the database that led the way for private development of a wide

range of small wind turbines that were placed in use during the late 1980s.[1503]

Closely related to the Mod-0A turbine was the Mod-1 project, for which planning started in 1976, with installation of the machine taking place in May 1979. In addition to noise level and television interference testing (see below), the primary objective of the Mod-1 program was to demonstrate the feasibility of remote utility wind turbine control. Three technical assessments were planned to evaluate machine performance, interface with the power utility, and examine the effects on the environ­ment. This system was a one-of-a-kind prototype that was much larger than the Mod-0A, with a rated power of 2,000 kilowatts (later reduced to 1,350) and a blade swept diameter of 200 feet. The Mod-1 was the largest wind turbine constructed up to that time. Considerable testing was done on the Mod-1 because the last experience with megawatt-size wind turbines was nearly 40 years earlier with the Smith-Putnam 1.25- megawatt machine, a very different design. Full-span blade pitch was used to control the rotor speed at a constant 35 revolutions per minute (later reduced to 23 rpm). The machine was mounted on a steel tubular truss tower that was 12 feet square at the top and 48 feet square at the bottom. General Electric was the prime contractor for designing, fabri­cating, and installing the Mod-1. The two steel blades were manufactured by the Boeing Engineering and Construction Company. There was also a set of composite rotor blades manufactured by the Kaman Aerospace Corporation that was fully compatible for testing on the Mod-1 system. The wind turbine, which was in Boone, NC, was tested with the Blue Ridge Electrical Membership Corporation from July 1979 to January 1981. The machine, operating in fully automatic synchronized mode, fed into the power network within utility standards.[1504]

One of the testing objectives of this first-generation prototype was to determine noise levels and any potential electromagnetic inter­ference with microwave relay, radio, and television associated with
mountainous terrain. These potential problems were among those identified by an initial study undertaken by NASA Lewis, General Electric, and the Solar Energy Research Institute. An analytical model developed at NASA Lewis of acoustic emissions from the rotor recommended that the rotor speed be reduced from 35 to 23 revolu­tions per minute, and the 2,000-kilowatt generator was replaced with a 1,350-kilowatt, 1,200-rpm generator. This change to the power train made a significant reduction in measured rotor noise. During the noise testing, however, the Mod-1, like the Mod-0A, experienced a failure in the low-speed shaft of the drive train and, because NASA engineers determined that both machines had accomplished their purposes, they were removed from the utility sites. Lessons learned from the engineer­ing studies and testing of the first-generation wind turbine systems indi­cated the need for technological improvements to make the machines more acceptable for large utility applications. These lessons proved valu­able in the design, construction, and operation of the next generation of DOE-NASA wind turbines. Other contributions from the Mod-1 pro­gram included low-cost wind turbine design concepts and metal and composite blade design and fabrication. Also, computer codes were verified for dynamic and loads analysis.

Although the Mod-1 was a one-of-kind prototype, there was a con­ceptual design that was designated as the Mod-1A. The conceptual design incorporated improvements identified during the Mod-1 project but, because of schedule and budget constraints, were not able to be used in fabrication of the Mod-1 machine. One of the improvements involved ideas to lessen the weight of the wind turbine. Also, one of the proposed configurations made use of a teetered hub and upwind blades with par­tial span control. Although the Mod-1A was not built, many of the ideas were incorporated into the second – and third-generation DOE-NASA wind turbines.