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An important issue for Aerospace and Defence Systems providers is how to reduce the risks associated with installing a new Fault Detection Tool (FDT) on a system. It is highly desirable that some degree of assessment, selection and validation is carried out before the FDT is integrated with the system. This paper describes the initial phases of a project to investigate the processes behind the assessment and validation using an Experimental Aircraft Fuel Rig (referred to as the Advanced Diagnostic Test-bed ADT). This paper also presents results from preliminary verification and validation work that has been used on a mathematical model of the ADT, and also some results from some initial diagnostic technique assessment that has been performed using real experimental data from the ADT and simulated data from mathematical models.

This paper demonstrates the application of a recently developed anti-windup technique for systems with rate-saturated actuators, to a realistic flight control example. The paper makes two contributions: firstly a new approach to tuning the anti-windup compensator is devised, allowing a transparent trade-off between performance and the size of an estimate of the region of attraction. Secondly, the anti-windup algorithm is applied to a nonlinear simulation model of the longitudinal and lateral dynamics of an experimental aircraft. The results show the potential of anti-windup to lessen an aircraft's susceptibility to pilot-induced-oscillations. This promise has since been confirmed in flight tests.

An accurate radio altimeter with a dual-channel (quadrature) homodyne receiver is presented in this article. This new design improves the accuracy while operating in the standard 4.3 GHz radio-altimeter allocation. The new radio altimeter is intended as a landing aid for small general-aviation aircraft, experimental aircraft and UAVs and is equipped with a simple synthesized-voice human interface

The U.S. Army Redstone Technical Test Center (RTTC) has been supporting captive flight testing of missile sensors and seekers since the 1980's. Successful integration and test of sensors in an airborne environment requires attention to a broad range of disciplines. Data collection requirements drive instrumentation and flight profile configurations, which along with cost and airframe performance factors influence the choice of test aircraft. Installation methods used for instrumentation must take into consideration environmental and airworthiness factors. In addition, integration of test equipment into the aircraft will require an airworthiness release; procedures vary between the government for military aircraft, and the Federal Aviation Administration (FAA) for the use of private, commercial, or experimental aircraft. Sensor mounting methods will depend on the type of sensor being used, both for sensor performance and crew safety concerns. Pilots will require navigation input to permit the execution of accurate and repeatable flight profiles. Some tests may require profiles that are not supported by standard navigation displays, requiring the use of custom hardware/software. Test locations must also be considered in their effect on successful data collection. Restricted airspace may also be required, depending on sensor emissions and flight profiles.

Online flight trajectory optimization algorithms are developed in order to guide and control aircraft for emergency landing. Further, the applicability of generated flight trajectories to real flights is evaluated using a flight simulator and an experimental aircraft. By combining the real-time A* path search algorithm with the R-TABU optimization method and an inverse dynamic method, a near-optimal four-dimensional flight trajectory is computed in real time. The proposed method is applied to design a flight trajectory for emergency landing. The generated flight trajectory is shown to pilots as a tunnel-in-the-sky image in order to track the trajectory by manual operation. Both the simulation and flight experiments refine the algorithms and demonstrate the applicability of the proposed system.

Eugene Ely’s demonstrations of take-off from and landing on ships was followed closely by Glenn Curtiss demonstrating seaplane take-off from and landing on water, and the beginning of catapult development for surface warships. However, U.S. Naval Aviation until after World War I centered on flying boats operating from coastal bases. By the 1920’s, catapults for surface ships and an experimental aircraft carrier received considerable attention, and a role for Naval Aviation in fleet operations was solidified. For surface ships the compressed air catapult and the powder catapult remained in use through World War II. Meanwhile, compressed air catapults on Langley were followed by fly wheel catapults on Lexington and Saratoga, and these in turn gave way to hydraulic catapults in the mid 1930s. During World War II the necessity of using catapults on Escort Carriers (CVEs), as well as the need for larger deck parks and night operations on the fleet carriers (CVs) made the catapult an integral part of carrier operations. With the introduction of jet aircraft to carriers, and the experience of the war in Korea, the hydraulic catapult was reaching its limits. The UK Royal Navy developed the steam catapult in the early 1950s which is still in use today. A discussion of the Flexdeck program in the mid 1950s is also included.

This paper considers the problem of robust static output H/sub 2/ controller design for linear parameter-varying (LPV) systems using parameter-dependent Lyapunov functions. The proposed method uses an iterative algorithm that has monotonic convergence to an optimal solution. The proposed design method is also applied to a previously proposed model-following controller design to overcome its drawbacks. We design several model-following controllers for lateral/directional motions of an experimental aircraft and demonstrate controller performance by experiments.

Transonic flutter wind tunnel tests with a flexible scaled model of the Jet-powered Supersonic Experimental Airplane were conducted at Japan Aerospace Exploration Agency (JAXA). The test results would be used for establishing and verifying analysis tools to evaluate transonic aeroelasticity. The test article was designed as 11% scaled model of the experimental airplane which length and width are 11.5m and 4.93m respectively. The wind tunnel was operated by sweeping either total pressure or by sweeping Mach number, or both proportionally. Typical dip shape boundary of transonic flutter was obtained and limit cycle oscillation was also revealed below the flutter boundary. In this paper, the summary of the flutter tests is reported. I. Introduction eroelastic flutter evaluation plays a significant role in the structural design of aircraft since flutter phenomenon can cause serious damage or loss of aircraft. Structural requirements for less weight and less cost make the role more important on high-speed aircraft, especially a supersonic transport (SST) due to the complexities in aerodynamic and structural aspects in transonic region. Transonic Flows are characterized by the presence of adjacent regions of subsonic and supersonic flow, usually accompanied by shock waves. Nonlinear behavior of the shock waves causes unexpected unstable vibrations such as transonic flutter (transonic-dip) and limit cycle oscillation (LCO) which are inconvenient for performance of high speed aircraft. In the past, there have been many activities, not only numerically but also experimentally, in investigation of this kind of nonlinear aeroelasticity. Since 1997, the National Aerospace Laboratory (NAL), presently the Japan Aerospace Exploration Agency (JAXA), has been conducting the National Experimental Supersonic Transport (NEXST) program to prepare the nation's technical base for coming international development of the next generation SST (Ref. 1). Major objectives of the NEXST program are as follows; a) to develop system integration technology for aircraft b) to establish CFD-based aerodynamic design technology with inverse and optimization method c) to progress component technologies in aerodynamic and composite structures design In order to demonstrate these technologies, the developments of two types of scaled supersonic experimental aircraft are promoted; one is non-powered experimental airplane and the other is jet-powered one. The non-powered airplane (NEXST-1) was developed to demonstrate the inverse aerodynamic design technology to realize a low drag configuration. The cranked arrow wing, area rule and warp concept are integrated with the natural laminar wing technology. During the development of the NEXST-1, the aeroelastic characteristics of a clean arrow wing are investigated elaborately (Ref.2). Following the NEXST-1, aeroelastic investigation of the jet-powered experimental airplane (NEXST-2) was conducted in the basic design phase. The NEXST-2 has a role to demonstrate the optimized-inverse design system and some component technologies including engine-airframe integration, high performance stable air-intake and composite structure for the wing. The NEXST-2 is characterized as unique

A global model approach developed on data from the experimental aircraft program X-31 EFM was recently improved during the X-31 VECTOR program. Both programs are briefly presented, focusing on extremely precise ESTOL landings following a slow, thrust-vectored approach at high angle of attack. High-accuracy navigation and inertial sensor systems enable onboard calculation of the height above runway with sufficient accuracy which is also required to identify ground effect. System identification procedure utilizes a global model to cover the entire flight regime including high-lift configuration during power approach and landing. Analysis of specially designed flight test led to aerodynamic increment tables for supplementary update of the original database. The wind-tunnel and computational fluid dynamics (CFD) predicted ground effect was incorporated into the higly accurate identified global model, which was then used for conventional landings resimulation to support the initial extremely short takeoff and landing (ESTOL) to the ground flight clearance process. Recently an incremental ground effect model was implemented to supplement the original data set. First identification results from conventional and ESTOL landings show some improvements compared to the predictions at very low height above ground.

Multi-objective design optimization of subsonic diffuser for supersonic aircraft has been performed by using Divided Range Multi-objective Genetic Algorithm (DRMOGA) to obtain higher efficiency and stability for the engine performance based on compressible viscous flow simulation. To model the diffuser geometry, the biquadratic equation is used for the centerline design and definition of the distribution of the area and the aspect ratio. The cross sectional shape and area of the diffuser is changed from a rectangular into a circle gradually by using simple polynomials. The present design was applied not only to a simplified duct model but also to an actual supersonic experimental aircraft.

On March 14, 2003 an experimental aircraft fitted with surface plasmon resonance (SPR) biosensors connected to an air sampling system performed a 90-minute flight over Renton, Washington, demonstrating the first-ever use of SPR sensors for airborne biodetection. In this paper, we describe the instrumentation constructed for this purpose, the experiment conducted, and the results obtained. Instrumentation was based on Texas Instruments’ Spreeta SPR sensors combined with sample collection and fluidic apparatus designed for airborne sensing. Detection targets were two innocuous proteins ovalbumin and horseradish peroxidase. We describe future enhancements necessary to apply this technology on an unmanned airborne vehicle.

This paper presents a summary of the plenary talk discussing the emerging need for unmanned aircraft, recent advances in nonlinear control methods for aircraft, and recent flight test programs involving the demonstration of advanced nonlinear control technologies. Topics presented include: neural adaptive control and flight testing on the X-36; adaptive formation flight control for close formation flight: and flight testing of the X-45A Joint-Unmanned Combat Air System.

Marques P. (2013). Emerging technologies in UAV aerodynamics. International Journal of Unmanned Systems Engineering. 1(S1): 3-4. Novel unmanned aerial vehicle (UAV) technologies address innovative airframe concepts for reasons of fuel efficiency, aircraft noise signature attenuation and reduced maintenance costs. New aerodynamic technologies prioritise flow control and flapless designs, in place of traditional hinged control surfaces, for enhanced manoeuvrability, reduced wing structural weight, and improved stealth. This technical note presents contemporary advances in Blended Wing Body (BWB) airframe design and flow control mechanisms for flapless flight control. Based upon the Coanda effect, engine thrust vectoring provides pitch control and Circulation Control (CC) replaces conventional ailerons for roll control. The integration of flapless aerodynamic technology in blended wing-body aircraft is envisaged to play a crucial role in future experimental flight research programmes for the development of next generation UAV concepts. © Marques Aviation Ltd – Press. 1. BLENDED WING BODY DESIGN The Blended Wing Body concept offers promising advantages in structural, aerodynamic and operational efficiency over the conventional fuselage-and-wing designs (Fig. 1). The unmanned Boeing X-48 Blended Wing Body experimental aircraft represents the futuristic philosophy supported by Boeing

For a research aircraft, "conventional" control laws (CLs) are implemented on a "baseline" flight computer (FC) while research CLs are typically housed on a dedicated research computer. Therefore, for an experimental aircraft used to test specific fault tolerant flight control systems, a safety logic scheme is needed to ensure a safe transition from conventional to research CLs (while at nominal conditions) as well as from research CLs at nominal conditions to conditions with "simulated" failures on specific control surfaces. This paper describes the design of such a safety scheme for the NASA Intelligent Flight Control System (IFCS) F-15 Program. The goals of the IFCS F-15 program are to investigate the performance of a set of fault tolerant CLs based on the use of dynamic inversion with neural augmentation. The different transitions are monitored using information relative to flight conditions and controller-related performance criteria. The testing of the scheme is performed with a Simulink-based flight simulation code and interface developed at West Virginia University for the NASA IFCS F-15 aircraft.

To explore the feasibility of utilizing an IR imaging system to support flow visualization studies, an initial series of tests were conducted using an AN/AAS-38, NITE Hawk targeting pod. The targeting pod, installed on the left side of an F/A-18 aircraft provides a stabilized infrared imaging capability in the 8-12 micron spectral band. Initial data acquired with system indicated that IR thermography was a very promising tool for flow visualization. For the next phase of the investigation, an advanced version of the NITE Hawk targeting pod equipped with a staring 3-5 micron sensor was utilized. Experimental results obtained with this sensor indicated improved sensitivity and resolution. This method was limited to position the experiment and chase aircraft sufficiently close to each other and with the sightline angle required to acquire the region of interest. For the current phase of the investigation, the proven 3-5 micron staring sensor was deployed in an externally mounted podlet, located on the experimental aircraft with a fixed line of sight, centered on the region of interest. Based on initial data collection efforts, this approach appears to provide consistent high quality data for a wide range of flight conditions. To minimize the size of the podlet and resultant drag, the sensor was oriented parallel to the air flow. This also placed the line of sight parallel to the experiment. A fold mirror was incorporated in the design to fold the line of sight inboard and down to center on the region of interest. The experimental results obtained during the current test phase have provided consistently high quality images clearly mapping regions of laminar and turbulent flow. Several examples of these images and further details of the experimental approach are presented.

An Attitude Heading Reference System (AHRS) is presented which has been developed for use in general aviation aircraft. The AHRS is contained in a 4.5"/spl times/4.0"/spl times/4.5" aluminum box, weighing 2.2 lb., with integral mounting flanges. Two connectors are provided, one for power and data out and another for a remote flux gate compass input. The AHRS is being marketed as part of an Electronic Flight Instrument System (EFIS) for use in experimental aircraft. Work is underway to certify the AHRS to enable its use in any aircraft as part of an Electronic Horizontal Situation Indicator (EHSI) and Electronic Air Data Indicator (EADI). The advantages of this system are low power usage, ease of interfacing the device to other aircraft systems and the durability and longevity of the solid state components. The internal architecture of the AHRS is presented along with the considerations used in the design process. The use of relatively inexpensive solid state piezo-electric rate gyros to determine instantaneous rotations about the three axes of the plane present unique problems due to the inherent drift characteristics of the rate gyros. Solid state G-cell accelerometers, with integral self-testing, are used as long term inertial reference devices to stabilize the gyros.

Multiple model adaptive estimation (MMAE) is a method of estimating unknown system parameters by modeling different discrete parameter values in several filters that are run in parallel and compared. The parameters for this research are failure status conditions associated with flight control actuators and sensors on the LAMBDA unmanned research vehicle, an experimental aircraft. Six actuator failures and eight sensor failures are modeled, along with the fully functional aircraft, in fifteen elemental Kalman filters. These filters propagate and update their own aircraft state estimates in real time. A probability computation representing the likelihood of each elemental filter's match to the true condition of the aircraft is used to generate relative probabilities for each filter's hypothesis. The MMAE algorithm is extended for the identification of dual failures through the use of a hierarchical structure of filter banks. Aircraft state excitation is required for effective MMAE performance. Sinusoidal dither signals are applied to the command inputs of a flight control system which controls pitch rate, roll rate, and sideslip angle. >

The growth of aviation since the first flight of a heavier-than-air powered manned vehicle in 1903 has been somewhat remarkable. Some of the events that have influenced this growth are reviewed in this paper. This review will include some events prior to World War I; the influence of the war itself; the events during the post-war years including the establishment of aeronautical research laboratories; and the influence of World War II which, among other things, introduced new technologies that included rocket and jet propulsion and supersonic aerodynamics. The subsequent era of aeronautical research and the attendant growth in aviation over the past half century will be reviewed from the view point of the author who, since 1944, has been involved in the NACA/NASA aeronautical research effort at what is now the Langley Research Center in Hampton, Virginia. The review will discuss some of the research programs related to the development of some experimental aircraft, the Century series of fighter aircraft, multi-mission aircraft, advanced military aircraft and missiles, advanced civil aircraft, supersonic transports, spacecraft and others.

This paper introduces a newly developed ultrasonic anemometer (UA) that can accurately measure three dimensional velocity components on the airplane. Calibration tests were initially done in a wind tunnel. Improvements were made it possible to measure an increased velocity range compared to the conventional UA. As a result, the new UA can be used in nearly all the flight speed ranges of an experimental aircraft (B-65) operated in National Aerospace Laboratory (NAL). To substantiate the wind speed data measured by the B-65, they were compared with data from the MU radar ground facility, observing the middle and upper atmosphere by emitting VHP radar. Results confirmed that the wind velocity measured by the UA has a difference of 0.1 m/s in the vertical component, and 1m/s in the horizontal component compared to the data by the MU radar. As one of the applications of the UA demonstrated, it was found that there is a slow upward flow in the water vapor freezing process at an altitude of between 2 and 3 km. Atmospheric turbulence trends, (the well known -5/3 power law) were also observed. Thus the newly developed UA offers a useful tool for atmospheric observation by an aircraft.

NOMENCLATURE In recent years, an extensive effort supported by the National Aeronautics and Space Administration (NASA) has been undertaken to rejuvenate the declining general aviation (GA) industry in the United States. One area of emphasis is in the development of advanced flight systems capable of simplifying the operation of GA aircraft in all weather conditions, opening new markets to the GA industry by making viable the use of private aircraft as transportation. Presented herein is a description of the design and plan for proof of concept of a state estimation system utilizing a low-cost distributed sensor array. This system will be used as the primary source of information for a decoupled flight control system on a V4 scale remotely piloted model aircraft named Aladdin, an entry in a student design competition sponsored by NASA, the Federal Aviation Administration (FAA), the Experimental Aircraft Association (EAA), and the Air Force. The state estimation system includes the incorporation of emerging technologies in the area of sensors and data processing, with redundancy achieved through dissimilar distributed sensor arrays. f Associate Professor * Professor and Head 5 Graduate Student Symbol Description Units a Acceleration vector body axes m/s a' Acceleration measurement m/s ~' vector from i" accelerometer package body axes g Gravitational acceleration vector m/s body axes d) or a Aircraft angular acceleration rad/s vector body axes to Aircraft angular velocity vector rad/s body axes Li Position vector locating i" m accelerometer package w/r aircraft CG body axes P, Q> R Components of the aircraft rad/s angular velocity vector, ea 0,0, y Euler angles defining the aircraft rad attitude £ Vector containing the four quaternion parameters used to define the aircraft attitude Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc.

In this paper the concept of metrological modeling was considered. The follow definition offered for a metrological modeling: experimental and/or analytically study of phenomena, processes and products, based on their similarity and common factors. The similarity criteria was founded for models and studied objects. We justified the use of metrological modeling method in weighing system calibration and aircraft balancing, as currently these systems calibrate with metrological characteristics definition for weight measurement and there is no accuracy estimation for the founded position of center of gravity. Because of difficulties to have a real scale object with precisely know position of center of gravity, it is appropriate to use metrological modeling with scaled model. Additionally, the realistic scaled experimental aircraft model for weighing system calibration was developed. High accuracy of the calibration of weighing systems and balance of aircraft was provided, in the result of model simplification and its ability to accurately determine the parameters.

Most existing subspace identification algorithms assume that a single input to output data set is available. Motivated by a real life problem on the F18-SRA experimental aircraft, we show how these algorithms are readily adapted to handle multiple data sets. We show by means of an example the relevance of such an improvement.

Parameter identification of the X-31A experimental aircraft was conducted throughout the envelope expansion using mainly pilot input maneuvers. The quality of the identification results, especially in the high-angle-of-attack regime, suffered from high correlations between the aircraft controls and states as well as from insufficient sideslip excitation. These problems are caused by the high feedback gains used in the flight control system to control the basically unstable airplane. Therefore, in 1994-1995 a dedicated flight test program using single surface excitation for parameter identification at high angles of attack was conducted. The results prove that identification of highly augmented aircraft benefits considerably from this technique. This article presents the realization of the single-surface excitation using the X-31A flutter test box. Also, the flight test program, the applied evaluation methods, and the identification models are discussed. Selected identification results from the single-surface excitation tests are presented and compared to those obtained from the pilot input maneuvers.

Parameter identification of the X-31A experimental aircraft was conducted throughout the envelope expansion using mainly pilot input maneuvers. The quality of the identification results, especially in the high-angle-of-attack regime, suffered from high correlations between the aircraft controls and states as well as from insufficient sideslip excitation. These problems are caused by the high feedback gains used in the flight control system to control the basically unstable airplane. Therefore, in 1994-1995 a dedicated flight test program using single surface excitation for parameter identification at high angles of attack was conducted. The results prove that identification of highly augmented aircraft benefits considerably from this technique. This article presents the realization of the single-surface excitation using the X-31A flutter test box. Also, the flight test program, the applied evaluation methods, and the identification models are discussed. Selected identification results from the single-surface excitation tests are presented and compared to those obtained from the pilot input maneuvers.

The X-31A experimental aircraft, a joint venture between Rockwell International, USA and Deutsche Aerospace, Germany, was designed and constructed to demonstrated enhanced fighter maneuverability including operation in the post stall regime. The DLR Institute of Flight Mechanics is involved in the flight testing phase by contributing its system identification capabilities for the extraction of mathematical models and aerodynamical parameters. It is a major challenge as the hasic aircraft is highly unstable and is only flown with the control system engaged. Therefore, the paper concentrates on two associated problem areas: the identification of unstable systems and the identification of open-loop aircraft dynamics from closed-loop flight test data. Both, timeand frequencydomain identification methods were applied to determine the coefhcients for the longitudinal motion of the X-31A. The techniques are characterized and representative results are to demonstrate and discuss the suitability and constraints of the individual approaches in detail.

A performance comparison and tradeoff study was conducted between eight unique aircraft configurations for high performance light aircraft and remotely piloted vehicles. These configurations included conventional tractor, conventional pusher, canard, tandem-wing, joined-wing, and 3-surface designs, which were analyzed though the use of microcomputer-based performance and lattice-vortex programs. Actual experimental aircraft were utilized as models which were scaled to a useful load of 600 pounds and given a common powerplant of 115 horsepower. The joined-wing, tandem-wing and conventional pusher were found to exhibit sufficient improvement over a conventional tractor configuration to warrant serious consideration for design selection. The performance and stability programs provided reasonably accurate predictions of aircraft performance when given actual aircraft dimensions and power available data and warrant use for preliminary aircraft design. http://archive.org/details/aircraftconfigur1094537539 Lieutenant Commander, United States Navy Approved for public release; distribution is unlimited.

controls of the Bell XS-1 experimental aircraft, became the first person to fly faster than the speed of sound. Aerodynamic design of this straight-wing monoplane was based on wind-tunnel tests of a model plus empirical knowledge, still embryonic, of airflow near sonic speed. Engineering theory, by contrast, took another five to six years to "break the sound barrier" by calculations for even a simple aerodynamic shape. The theoretical problem thus had its initial solution after the practical one, a common situation in engineering. The story of the XS-1 has been told many times. This is the story of the theoretical breakthrough. The making of engineering theory receives little detailed attention in the historical literature. Events and key players in this form of engineering science have been written about, but few fine-grained case studies exist of how a theory came into being.' This is in a sense curious, since history of technology has been focused largely on artifacts, and an engineering theory is itself an artifact-more specifically, a tool-to be employed in designing other artifacts. Theoretical

Presentation of the Daniel and Florence Guggenheim International Memorial Lecture is both a high honour and a challenge. The challenge is implicit in the final two words—to present to you both a lecture and a fitting memorial to the remarkable Guggenheim family. I believe that the subject which I have chosen is particularly appropriate for this task. As many of you know, Daniel Guggenheim, in the years from 1925 to his death in 1930, sponsored and endowed the now famous Guggenheim Aeronautical Laboratories in major universities across the USA. It is clear from his writing that these were not simply gifts to the universities—conventional endowments—but that they reflected Daniel Guggenheim's vision of a future in which the aircraft would link cities and countries in a safe, efficient civil transport system surpassing all other modes of travel.

I T is challenging to have the opportunity to present views regarding fighter aircraft at this time, in the second half of the 1970's, since the western aviation world has just witnessed and is still witnessing, after a pause of some 15 years, a wave of new fighter designs being developed and launched into production and service. Some of these new airplanes will dominate the scene for the next 15 years or more, that is, well into the 1990's. Nevertheless, are we already entering the period where concepts for yet another generation of fighters are being developed? The evolution of the fighter aircraft, and there have been a number of revolutionary steps, commenced with powered flight as pioneered by the Wright Brothers. A brief review of this history may provide a better perspective for our views, even if the author of this paper is not a researcher of aircraft history and therefore may not be accurate in detail. The initial generation of fighter aircraft (in W.W.I) had to solve the problem of developing an effective armament, the art of maneuvering flight having been provided by the Wright Brothers only a few years earlier. The unarmed early airplanes were nevertheless providing effective reconnaissance and were as such already "fighters." The Wright Brothers thus delivered the world's first actual fighter to the U.S. Army Signal Corps. The "armed fighter" was, in effect, only a reaction to an earlier airborne threat to the land and naval forces. The initial fighter armament of handheld guns was soon overtaken by the aircraft-mounted machine gun, but it was difficult for the pilot to control the airplane with one hand and to point the gun with the other, especially if a propeller in the front was in the way of the natural line of sight. The obvious solution of taking a weapons operator or "gunner" along was detrimental to the fighter's rate of climb and speed. The other solution of reverting to a "pusher installation" of the engine also resulted in a heavier airplane. The truly ingenious solution was that of firing through the tractor-propeller with a rigidly mounted gun and accurately pointing and aiming by controlling the direction of flight and attitude of the aircraft. The propeller was protected first by local armor and later by a synchronizing system. This system was, I believe, invented by R. Garros of France in 1915, met instant success, and set a pattern which is still valid. I am recounting this well-known history because I believe that it really started the fighters as a special breed of airplane (see Fig. 1). After its birth, the fighter aircraft during the next 20 years evolved apparently with the main aim of maximum l-g SEP or rate of climb. This was done by increasing engine power in a number of steps from 80 hp by a factor of 10 while improving specific weights such that gross weight increased from 1100 Ib by a factor of 3.5 only. (Power loading by 2.85.) Airframe size changed little in this period; it was built around the pilot with a wingspan of 30 ft. Aerodynamics must have been subordinate to the structural designers who retained the light, simple, and well-proven externally braced and fabric-covered designs. Materials remained generally steel tube and wood, even though aluminum had already been used for the Zeppelin airships. This type of construction also had the advantage of being easy to repair. By the early 1930's, however, the designers of airliners and bombers really started to apply aerodynamics, including retractable landing gears, and combined this with higher wing loadings and stress-skin aluminum structures. They were outspeeding the contemporary fighters, which did only about 230 mph with an engine of 600 hp. The fighter community had to react since it could not justify its existence for long by pointing out how excellent they remained in fighting their own

The X-29 is an experimental aircraft built by the Grumman Aerospace Corp. for the US Defense Research Projects Agency. Relaxed static stability, one of the aircraft's key design features, means that the pilot must depend exclusively on computers to fly the plane. The X-29's computers monitor and adjust the positions of the plane's control surfaces every 25 milliseconds to follow the pilot's directions. Because it is designed to be unstable, the X-29 has its center of lift ahead of the center of gravity rather than behind it, unlike most stable aircraft. The way that the aircraft's canards and flaperons contribute to maneuverability is explained, and an account is given of how a computer-aided design and manufacturing system was used to design the composite wing. The aircraft has three digital computers backed up by three analog computers. The three digital computers are linked to one another by data buses in a ring-like configuration.

The X-29 forward-swept wing aircraft required a power takeoff shaft to transmit power from the starter to the engine during engine starting and to transmit power from the engine to the remote gearbox during ground and flight operations of aircraft accessories. Shaft design requirements of high speed and great length could not be met with off-the-shelf designs. Meeting adequate critical speed margin posed a significant design problem. Various shaft concepts were studied. The final selection was a graphite/epoxy center shaft utilizing metal flexible end assemblies. Extensive testing at unit, subsystem, and system levels proved the adequacy of the design.

A joint Dutch/German aircraft parameter identification program has been conducted to investigate various factors affecting identification results. Different design methods led to five test signals which were implemented via a hydraulic control system in a flight test program with a De Havilland DHO-2 Beaver experimental aircraft. Aircraft parameters were calculated using two different parameter identification algorithms, classical Maximum Likelihood method and flight path reconstruction followed by linear regression analysis. Each test signal was repeated nine resp. ten times allowing a statistical evaluation based on Cramer-Rao lower bounds and sample variances.

This article represents a first attempt at numerical simulation of the flow field around a complete lifting body configuration utilizing the Reynolds-averaged Navier-Stokes equations. The numerical solution generated for the experimental aircraft concept X24C-10D, at a Mach number of 5.95, exhibited accurate prediction of detailed flow properties and integrated areodynamic coefficients as well. In addition, the present analysis demonstrated that a vectorized processor with an external memory device is an efficient and viable means for processing the Navier-Stokes equations.

An analytical study was conducted to assess the impact of the use of broad specification fuels with reduced hydrogen content on the design, performance, durability, emissions, and operational characteristics of combustors for commercial aircraft gas turbine engines. The study was directed at defining necessary design revisions to combustors designed for use of Jet A when such are operated on ERBS (Experimental Referee Broad Specification) fuel which has a nominal hydrogen content of 12.8% in current Jet A. The results indicate that improvements in combustor liner cooling, and/or materials, and methods of fuel atomization will be required if the chemical and physical fuel properties approach those of ERBS.

The AD-1 manned flight test program was conducted to evaluate the stability, control and handling characteristics of oblique wing aircraft. The results of the aeroelastic stability analysis are presented for both the wing alone and the wing with ailerons. A comparison was made between the results obtained using the traditional k-method of flutter analysis and the results using the PK or British method of flutter analysis. Studies were performed using the latest version of the NASTRAN computer code as well as the PASS/FLUT program.

Reliable methods for predicting external aircraft flowfields are required for airframe/engine and aircraft/store integrations, as well as for predicting store trajectories. Several analytical techniques exist that can predict the flowfield about complete aircraft configurations. Three of these - the PAN AIR panel method, the Boppe transonic small disturbance code, and the influence function method (IFM) technique - are examined in this paper. All three methods have certain advantages and disadvantages. The PAN AIR method has the greatest geometric flexibility, but only gives consistently good results at subsonic speeds and lacks an integrated boundary layer calculation. The Boppe code gives a better prediction of wing-induced flowfield effects at all Mach numbers but cannot properly account for body-induced effects. The IFM method has the best ability to predict aircraft flowfield effects but requires store grid force and moment data in a horizontal traverse to make the flowfield predictions.