ALBERT

Description: The aerodynamic development of an engine inlet requires a comprehensive program of both wind tunnel testing and Computational Fluid Dynamics (CFD) simulations. To save time and resources, much "testing" is done using CFD before any design ever enters a wind tunnel. The focus of my project this summer is on CFD analysis tool development. In particular, I am working to further develop the capabilities of the Planar Inlet Design and Analysis Process (PINDAP). "PINDAP" is a collection of computational tools that allow for efficient and accurate design and analysis of the aerodynamics about and through inlets that can make use of a planar (two-dimensional or axisymmetric) geometric and flow assumption. PINDAP utilizes the WIND CFD flow solver, which is capable of simulating the turbulent, compressible flow field. My project this summer is a continuation of work that I performed for two previous summers. Two years ago, I used basic features of the PINDAP to design a Mach 5 hypersonic scramjet engine inlet and to demonstrate the feasibility of the PINDAP. The following summer, I worked to develop its geometry and grid generation capabilities to include subsonic and supersonic inlets, complete bodies and cowls, conic leading and trailing edges, as well as airfoils. These additions allowed for much more design flexibility when using the program.

Description: We have designed and built an instrument to measure and monitor the "nightglow" of the Earth's atmosphere in the near ultraviolet (NUV). In this paper we describe the design of this instrument, called NIGHTGLOW. NIGHTGLOW is designed to be flown-from a high altitude research balloon, and circumnavigate the globe. NIGHTGLOW is a NASA, University of Utah, and New Mexico State University project. A test flight took place from Palestine, Texas on July 5, 2000, lasting about 8 hours. The instrument performed well and landed safely in Stiles, Texas with little damage. The resulting measurements of the NUV nightglow are consistent with previous measurements from sounding rockets and balloons. The results will be presented and discussed.

Description: A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

Description: High performance aircraft are, by their very nature, often required to undergo maneuvers involving high angles of attack. Under these conditions unsteady vortices emanating from the wing and the fuselage will impinge on the twin fins (required for directional stability) causing excessive buffet loads, in some circumstances, to be applied to the aircraft. These loads result in oscillatory stresses, which may cause significant amounts of fatigue damage. Active control is a possible solution to this important problem. A full-scale test was carried out on an F/A-18 fuselage and fins using piezoceramic actuators to control the vibrations. Buffet loads were simulated using very powerful electromagnetic shakers. The first phase of this test was concerned with the open loop system identification whereas the second stage involved implementing linear time invariant control laws. This paper looks at some of the problems encountered as well as the corresponding solutions and some results. It is expected that flight trials of a similar control system to alleviate buffet will occur as early as 2001.

Description: In this poster, we describe a web-based tool for verification and automatic generation of user interfaces. The verification component of the tool accepts as input a model of a machine and a model of its interface, and checks that the interface is adequate (correct). The generation component of the tool accepts a model of a given machine and the user's task, and then generates a correct and succinct interface. This write-up will demonstrate the usefulness of the tool by verifying the correctness of a user interface to a flight-control system. The poster will include two more examples of using the tool: verification of the interface to an espresso machine, and automatic generation of a succinct interface to a large hypothetical machine.

Description: To achieve NASA's ambitious mission objectives for the future, aircraft and spacecraft will need intelligence to take the correct action in a variety of circumstances. Vehicle intelligence can be defined as the ability to "do the right thing" when faced with a complex decision-making situation. It will be necessary to implement integrated autonomous operations and low-level adaptive flight control technologies to direct actions that enhance the safety and success of complex missions despite component failures, degraded performance, operator errors, and environment uncertainty. This paper will describe the array of technologies required to meet these complex objectives. This includes the integration of high-level reasoning and autonomous capabilities with multiple subsystem controllers for robust performance. Future intelligent systems will use models of the system, its environment, and other intelligent agents with which it interacts. They will also require planners, reasoning engines, and adaptive controllers that can recommend or execute commands enabling the system to respond intelligently. The presentation will also address the development of highly dependable software, which is a key component to ensure the reliability of intelligent systems.

Description: The placement of actuators on a wing determines the control effectiveness of the airplane. One approach to placement maximizes the moments about the pitch, roll, and yaw axes, while minimizing the coupling. For example, the desired actuators produce a pure roll moment without at the same time causing much pitch or yaw. For a typical wing, there is a large set of candidate locations for placing actuators, resulting in a substantially larger number of combinations to examine in order to find an optimum placement satisfying the mission requirements and mission constraints. A genetic algorithm has been developed for finding the best placement for four actuators to produce an uncoupled pitch moment. The genetic algorithm has been extended to find the minimum number of actuators required to provide uncoupled pitch, roll, and yaw control. A simplified, untapered, unswept wing is the model for each application.

Description: Flow separation in the low-pressure turbine (LPT) is a major barrier that limits further improvements of aerodynamic designs of turbine airfoils. The separation is responsible for performance degradation, and it prevents the design of highly loaded airfoils. The separation can be delayed, reduced, or eliminated completely if flow control techniques are used. Successful flow control technology will enable breakthrough improvements in gas turbine performance and design. The focus of this research project was the development and experimental demonstration of active separation control using glow discharge plasma (GDP) actuators in flow conditions simulating the LPT. The separation delay was shown to be successful, laying the foundation for further development of the technologies to practical application in the LPT. In a fluid mechanics context, the term "flow control" means a technology by which a very small input results in a very large effect on the flow. In this project, the interest is to eliminate or delay flow separation on LPT airfoils by using an active flow control approach, in which disturbances are dynamically inserted into the flow, they interact with the flow, and they delay separation. The disturbances can be inserted using a localized, externally powered, actuating device, examples are acoustic, pneumatic, or mechanical devices that generate vibrations, flow oscillations, or pulses. A variety of flow control devices have been demonstrated in recent years in the context of the external aerodynamics of aircraft wings and airframes, where the incoming flow is quiescent or of a very low turbulence level. However, the flow conditions in the LPT are significantly different because there are high levels of disturbances in the incoming flow that are characterized by high free-stream turbulence intensity. In addition, the Reynolds number, which characterizes the viscous forces in the flow and is related to the flow speed, is very low in the LPT passages.

Description: Researchers are constantly pursuing technologies that will increase the performance of gas turbine engines. The aspirated compressor concept discussed here would allow the compression system to perform its task with about one-half of the compressor blades. To accomplish this, the researchers applied boundary layer control to the blades, casing, and hub. This method of boundary layer control consisted of removing small amounts of air from the main flow path at critical areas of the compressor. This bleed air could be used by other systems such as engine cooling or could be re-injected into lower pressure areas that require air for enhanced performance. This effort was initiated by the Massachusetts Institute of Technology (MIT) in response to a solicitation from the Defense Advanced Research Projects Agency (DARPA) who sought to advance research in flow control technology. The NASA Glenn Research Center partnered with MIT (principal investigator), Honeywell Aircraft Engines (cycle analysis, structural analysis, and mechanical design), and Pratt & Whitney (cycle analysis and aero-analysis) to conceptualize, design, analyze, build, and test the aspirated fan stage. The aero-design and aero-analysis of this fan stage were jointly executed by MIT and Glenn to minimize the amount of bleed flow needed and to maintain the highest efficiency possible (ref. 1). Mechanical design issues were complicated by the need to have a shrouded rotor with hollow blades, with rotor stress levels beyond the capabilities of titanium. The high stress issues were addressed by designing a shroud that was filament wound with a carbon fiber/epoxy matrix, resulting in an assembly that was strong enough to handle the high stresses. Both the rotor (preceding photographs) and stator (following photograph) were fabricated in two halves and then bolted together at the hub and tip, permitting the bleed passages to be machined into each half before assembly.

Description: NASA Ames has a long tradition in leadership with the use of ballistic ranges and shock tubes for the purpose of studying the physics and phenomena associated with hypervelocity flight. Cutting-edge areas of research run the gamut from aerodynamics, to impact physics, to flow-field structure and chemistry. This legacy of testing began in the NACA era of the 1940's with the Supersonic Free Flight Tunnel, and evolved dramatically up through the late 1950s with the pioneering work in the Ames Hypersonic Ballistic Range. The tradition continued in the mid-60s with the commissioning of the three newest facilities: the Ames Vertical Gun Range (AVGR) in 1964, the Hypervelocity Free Flight Facility (HFFF) in 1965 and the Electric Arc Shock Tube (EAST) in 1966. Today the Range Complex continues to provide unique and critical testing in support of the Nation's programs for planetary geology and geophysics; exobiology; solar system origins; earth atmospheric entry, planetary entry, and aerobraking vehicles; and various configurations for supersonic and hypersonic aircraft.

Description: The benchmark solution for the cascade-gust interaction problem is computed using a linearized Euler code called LINFLUX. The inherently three-dimensional code is run in the thin-annulus limit to compute the two-dimensional cascade response. The calculations are carried out in the frequency-domain and the unsteady response at each of the gust s three frequency component is computed. The results are presented on modal basis for pressure perturbations (i.e., acoustic modes) as well as velocity perturbations (i.e., convected gust modes) at each frequency.

Description: An unsteady aerodynamic code, called GUST3D (ref. 3), has been developed to solve equation (8) for flows with periodic vortical disturbances. The code uses a frequency-domain approach with second-order central differences and a pressure radiation condition in the far field. GUST3D requires as input certain mean flow quantities which are calculated separately by a potential flow solver. The solver calculates the mean ow using a Gothert's Rule approximation (ref. 3). On the airfoil surface, it uses the solution calculated by the potential code FLO36 (ref. 4). Figures 1-2 show the mean pressure along the airfoil surface for the two airfoil geometries. In Figures 3 - 8, we present the RMS pressure on the airfoil surface. Each figure shows three GUST3D solutions (calculated on grids with different far-field boundary locations). Three solutions are shown to provide some indication of the numerical uncertainty in the results. Figures 9 - 13 present the acoustic intensity. We again show three solutions per case. Note that no results are presented for the k1 = k2 = 2.0 loaded airfoil case, as an acceptable solution could not be obtained. A few comments need to be made about the results shown. First, since the last Workshop, the GUST3D code has been substantially upgraded. This includes implementing a more accurate far-field boundary condition (ref. 5) and developing improved gridding capabilities. This is the reason for any differences that may exist between the present results and results from the last Workshop. Second, the intensity results on the circle R = 4C were obtained using a Kirchoff method (ref. 6). The Kirchoff surface was the circle R = 2C. Finally, the GUST3D code is most accurate for low reduced frequencies. A new domain decomposition approach (ref. 7) has been developed to improve accuracy. Both the single domain and domain decomposition approaches were used in generating the present results.

Description: The TURBO-AE aeroelastic code has been used to verify the flutter stability of the Efficient Low Noise Fan (ELNF), which is also referred to as the trailing-edge blowing fan. The ELNF is a unique technology demonstrator being designed and fabricated at the NASA Glenn Research Center for testing in Glenn's 9-by-15-Foot Low-Speed Wind Tunnel. In the ELNF, air can be blown out of slots near the trailing edges of the fan blades to fill in the wakes downstream of the rotating blades. This filling of the wakes leads to a reduction of the rotor-stator interaction (tone) noise that results from the interaction of wakes with the downstream stators. The ELNF will demonstrate a 1.6-EPNdB1 reduction in tone noise through wake filling, without increasing the broadband noise. Furthermore, the reduced blade row interaction will decrease the possibility of forced response and enable closer spacing of blade rows, thus reducing engine length and weight. During the design of the ELNF, the rotor blades were checked for flutter stability using the detailed aeroelastic analysis capability of the three-dimensional Navier-Stokes TURBOAE code. The aeroelastic calculations were preceded by steady calculations in which the blades were not allowed to vibrate. For each rotational speed, as the back-pressure was increased, the mass flow rate decreased, and the operating point moved along the constant speed characteristic (speed-line) from choke to stall as shown on the fan map. The TURBO-AE aeroelastic analyses were performed separately for the first two vibration modes (bending and torsion) and covered the complete range of interblade phase angles or nodal diameters at which flutter can occur. The results indicated that the ELNF blades would not encounter flutter at takeoff conditions. The calculations were then repeated for a part-speed condition (70-percent rotational speed), and the results again showed no flutter in the operating region. On the fan map (shown), the predicted flutter point at part speed condition was located beyond the stall line, which means that the ELNF will not encounter flutter since it will never operate beyond the stall line. All the calculations done so far have been for the nonblowing case, and selected calculations will be repeated with air blowing from the trailing edge of the fan.

Description: Recently, there has been an increase in the development of intelligent engine technology with advanced active component control. The computer engine models used in these control studies are component-level models (CLM), models that link individual component models of state space and nonlinear algebraic equations, written in a computer language such as Fortran. The difficulty faced in performing control studies on Fortran-based models is that Fortran is not supported with control design and analysis tools, so there is no means for implementing real-time control. It is desirable to have a simulation environment that is straightforward, has modular graphical components, and allows easy access to health, control, and engine parameters through a graphical user interface. Such a tool should also provide the ability to convert a control design into real-time code, helping to make it an extremely powerful tool in control and diagnostic system development. Simulation time management is shown: Mach number versus time, power level angle versus time, altitude versus time, ambient temperature change versus time, afterburner fuel flow versus time, controller and actuator dynamics, collect initial conditions, CAD output, and component-level model: CLM sensor, CAD input, and model output. The Controls and Dynamics Technologies Branch at the NASA Glenn Research Center has developed and demonstrated a flexible, generic turbofan engine simulation platform that can meet these objectives, known as the Modular Aero-Propulsion System Simulation (MAPSS). MAPSS is a Simulink-based implementation of a Fortran-based, modern high pressure ratio, dual-spool, low-bypass, military-type variable-cycle engine with a digital controller. Simulink (The Mathworks, Natick, MA) is a computer-aided control design and simulation package allows the graphical representation of dynamic systems in a block diagram form. MAPSS is a nonlinear, non-real-time system composed of controller and actuator dynamics (CAD) and component-level model (CLM) modules. The controller in the CAD module emulates the functionality of a digital controller, which has a typical update rate of 50 Hz. The CLM module simulates the dynamics of the engine components and uses an update rate of 2500 Hz, which is needed to iterate to balance mass and energy among system components. The actuators in the CAD module use the same sampling rate as those in the CLM. Two graphs of normalized spool speed versus time in seconds and one graph of normalized average metal temperature versus time in seconds is shown. MAPSS was validated via open-loop and closed-loop comparisons with the Fortran simulation. The preceding plots show the normalized results of a closed-loop comparison looking at three states of the model: low-pressure spool speed, high-pressure spool speed, and the average metal temperature measured from the combustor to the high-pressure turbine. In steady state, the error between the simulations is less than 1 percent. During a transient, the difference between the simulations is due to a correction in MAPSS that prevents the gas flow in the bypass duct inlet from flowing forward instead of toward the aft end, which occurs in the Fortran simulation. A comparison between MAPSS and the Fortran model of the bypass duct inlet flow for power lever angles greater than 35 degrees is shown.

Description: The difference in delamination onset predictions based on the type and location of the assumed initial damage are compared in a specimen consisting of a tapered flange laminate bonded to a skin laminate. From previous experimental work, the damage was identified to consist of a matrix crack in the top skin layer followed by a delamination between the top and second skin layer (+45 deg./-45 deg. interface). Two-dimensional finite elements analyses were performed for three different assumed flaws and the results show a considerable reduction in critical load if an initial delamination is assumed to be present, both under tension and bending loads. For a crack length corresponding to the peak in the strain energy release rate, the delamination onset load for an assumed initial flaw in the bondline is slightly higher than the critical load for delamination onset from an assumed skin matrix crack, both under tension and bending loads. As a result, assuming an initial flaw in the bondline is simpler while providing a critical load relatively close to the real case. For the configuration studied, a small delamination might form at a lower tension load than the critical load calculated for a 12.7 mm (0.5") delamination, but it would grow in a stable manner. For the bending case, assuming an initial flaw of 12.7 mm (0.5") is conservative, the crack would grow unstably.

Description: The intent of this talk is to present the stability and control (S and C) priorities as seen by the Langley team. No roadmaps or 5 year plans will be presented. We are actively soliciting your feedback, your ideas, and your help in building and executing this program. The outline of this viewgraph presentation includes: 1) Background; 2) NASA Constraints and Priorities; 3) Potential Program Content (high priority issues, approach); 4) Prepared Critiques; 5) Comments by Attendees; 6) Closing Comments.

Description: The Compressor Branch vision is to be recognized as world-class leaders in research for fluid mechanics of compressors. Its mission is to conduct research and develop technology to advance the state of the art of compressors and transfer new technology to U.S. industries. Maintain partnerships with U.S. industries, universities, and other government organizations. Maintain a balance between customers focused and long range research. Flow control comprises enabling technologies to meet compression system performance requirements driven by emissions and fuel reduction goals (e.g., in UEET), missions (e.g., access-to-space), aerodynamically aggressive vehicle configurations (e.g., UAV and future blended wing body configurations with highly distorted inlets), and cost goals (e.g., in VAATE). The compression system requirements include increased efficiency, power-to-weight, and adaptability (i.e., robustness in terms of wide operability, distortion tolerance, and engine system health and reliability). The compressor flow control task comprises efforts to develop, demonstrate, and transfer adaptive flow control technology to industry to increase aerodynamic loading at current blade row loss levels, to enable adaptive1 y wide operability, and to develop plant models for adaptive compression systems. In this context, flow control is the controlled modification of a flow field by a deliberate means beyond the natural (uncontrolled) shaping of the solid surfaces that define the principal flow path. The objective of the compressor flow control task is to develop and apply techniques that control circulation, aerodynamic blockage, and entropy production in order to enhance the performance and operability of compression systems for advanced aero-propulsion applications. This summer I would be working with a curved-diffuser because it simulates what happens with flow in the stator blades in the compressor. With this experiment I will be doing some data analysis and parametric study of the injector slot geometries to get the best aerodynamic performance of it. This includes some data reduction, redesign and fast prototyping of the injector nozzle.

Description: Flow structure of momentum-dominated pure helium jets discharged vertically into ambient air was investigated using high-speed rainbow schlieren deflectometry (RSD) technique. Effects of the operating parameters, i.e., Reynolds number (Re) and Richardson number (Ri), on the oscillatory behavior of the flow were examined over a range of experimental conditions. To seek the individual effect of these parameters, one of them was fixed and the other was varied with certain constraints. Measurements revealed highly periodic oscillations in the laminar region as well as high regularity in transition and turbulent regions. Maximum spectral power profiles at different axial locations indicated the oscillation amplitude increasing until the breakdown of the jet in the turbulent regime. The transition from the laminar to turbulent flow was also investigated. Fast Fourier transform analysis performed in the transition regime showed that the flow oscillates at a unique frequency, which was the same in the upstream laminar flow region. Measured deflection angle data were used in Abel inversion algorithm to construct the helium concentration fields. Instantaneous helium concentration contours revealed changes in the flow structure and evolution of vortical structures during an oscillation cycle. Temporal evolution plots of helium concentration at different axial location showed repeatable oscillations at all axial and radial locations up to the turbulent regime. A cross-correlation technique, applied to find the spatial displacements of the vortical structures, provided correlation coefficient peaks between consecutive schlieren images. Results show that the vortical structure convected and accelerated only in the axial direction.

Description: The NASA-Ames ballistic range provides a unique capability for aerothermodynamic testing of configurations in hypersonic, real-gas, free-flight environments. The facility can closely simulate conditions at any point along practically any trajectory of interest experienced by a spacecraft entering an atmosphere. Sub-scale models of blunt atmospheric entry vehicles are accelerated by a two-stage light-gas gun to speeds as high as 20 times the speed of sound to fly ballistic trajectories through an 24 m long vacuum-rated test section. The test-section pressure (effective altitude), the launch velocity of the model (flight Mach number), and the test-section working gas (planetary atmosphere) are independently variable. The model travels at hypersonic speeds through a quiescent test gas, creating a strong bow-shock wave and real-gas effects that closely match conditions achieved during actual atmospheric entry. The challenge with ballistic range experiments is to obtain quantitative surface measurements from a model traveling at hypersonic speeds. The models are relatively small (less than 3.8 cm in diameter), which limits the spatial resolution possible with surface mounted sensors. Furthermore, since the model is in flight, surface-mounted sensors require some form of on-board telemetry, which must survive the massive acceleration loads experienced during launch (up to 500,000 gravities). Finally, the model and any on-board instrumentation will be destroyed at the terminal wall of the range. For these reasons, optical measurement techniques are the most practical means of acquiring data. High-speed thermal imaging has been employed in the Ames ballistic range to measure global surface temperature distributions and to visualize the onset of transition to turbulent-flow on the forward regions of hypersonic blunt bodies. Both visible wavelength and infrared high-speed cameras are in use. The visible wavelength cameras are intensified CCD imagers capable of integration times as short as 2 ns. The infrared camera uses an Indium Antimonide (InSb) sensor in the 3 to 5 micron band and is capable of integration times as short as 500 ns. The projectiles are imaged nearly head-on using expendable mirrors offset slightly from the flight path. The proposed paper will discuss the application of high-speed digital imaging systems in the NASA-Ames hypersonic ballistic range, and the challenges encountered when applying these systems. Example images of the thermal radiation from the blunt nose of projectiles flying at nearly 14 times the speed of sound will be given.

Description: Single crystal superalloy turbine blades used in high pressure turbomachinery are subject to conditions of high temperature, triaxial steady and alternating stresses, fretting stresses in the blade attachment and damper contact locations, and exposure to high-pressure hydrogen. The blades are also subjected to extreme variations in temperature during start-up and shutdown transients. The most prevalent HCF failure modes observed in these blades during operation include crystallographic crack initiation/propagation on octahedral planes, and noncrystallographic initiation with crystallographic growth. Numerous cases of crack initiation and crack propagation at the blade leading edge tip, blade attachment regions, and damper contact locations have been documented. Understanding crack initiation/propagation under mixed-mode loading conditions is critical for establishing a systematic procedure for evaluating HCF life of single crystal turbine blades. This paper presents analytical and numerical techniques for evaluating two and three dimensional subsurface stress fields in anisotropic contacts. The subsurface stress results are required for evaluating contact fatigue life at damper contacts and dovetail attachment regions in single crystal nickel-base superalloy turbine blades. An analytical procedure is , presented, for evaluating the subsurface stresses in the elastic half-space, using a complex potential method outlined by Lekhnitskii. Numerical results are presented for cylindrical and spherical anisotropic contacts, using finite element analysis. Effects of crystal orientation on stress response and fatigue life are examined.

Description: Recent research has shown that adaptive neural based control systems are very effective in restoring stability and control of an aircraft in the presence of damage or failures. The application of an adaptive neural network with a flight critical control system requires a thorough and proven process to ensure safe and proper flight operation. Unique testing tools have been developed as part of a process to perform verification and validation (V&V) of real time adaptive neural networks used in recent adaptive flight control system, to evaluate the performance of the on line trained neural networks. The tools will help in certification from FAA and will help in the successful deployment of neural network based adaptive controllers in safety-critical applications. The process to perform verification and validation is evaluated against a typical neural adaptive controller and the results are discussed.

Description: Traditional control has proven to be ineffective to deal with catastrophic changes or slow degradation of complex, highly nonlinear systems like aircraft or spacecraft, robotics, or flexible manufacturing systems. Control systems which can adapt toward changes in the plant have been proposed as they offer many advantages (e.g., better performance, controllability of aircraft despite of a damaged wing). In the last few years, use of neural networks in adaptive controllers (neuro-adaptive control) has been studied actively. Neural networks of various architectures have been used successfully for online learning adaptive controllers. In such a typical control architecture, the neural network receives as an input the current deviation between desired and actual plant behavior and, by on-line training, tries to minimize this discrepancy (e.g.; by producing a control augmentation signal). Even though neuro-adaptive controllers offer many advantages, they have not been used in mission- or safety-critical applications, because performance and safety guarantees cannot b e provided at development time-a major prerequisite for safety certification (e.g., by the FAA or NASA). Verification and Validation (V&V) of an adaptive controller requires the development of new analysis techniques which can demonstrate that the control system behaves safely under all operating conditions. Because of the requirement to adapt toward unforeseen changes during operation, i.e., in real time, design-time V&V is not sufficient.

Description: The Flight Dynamics Analysis Branch (FDAB) of NASNGSFC is tasked with determining the orbital lifetime of several developmental and operational satellites, which include the Hubble Space Telescope. A DOS based program developed by the FDAB many years ago, called PC Lifetime, is used to determine a satellite s lifetime and could soon be in need of a replacement. STK s Lifetime Object Tool is a possible candidate. Due to the reduced support of the PC Lifetime program, and the growing incompatibility of older programs with new operating systems, a comparative analysis was done to determine if STWLifetime could meet the stringent requirements that were laid before it. The use of highly accurate numerical propagators such as STK s High Precision Orbit Propagator ( OP) and the Goddard Trajectory Determination System (GTDS) provided a basis on which to compare STWLifetime s results. Several test cases were run, but the main four test cases would determine whether or not STWLifetime could be PC- Lifetime s replacement. These four cases include a geotransfer orbit, two circular LEOS, and a Poiar LEO. Following rigorous testmg procedures, a conclusion will be determined. STK has proved to be a versatile program on many satellite missions and the FDAB has high hopes that it can pass FDAB s requirements for orbital lifetime prediction.

Description: This paper describes the development of a planned approach for Autonomous operation of an Unmanned Aerial Vehicle (UAV). A Hybrid approach will seek to provide Knowledge Generation thru the application of Artificial Intelligence (AI) and Intelligent Agents (IA) for UAV control. The application of many different types of AI techniques for flight will be explored during this research effort. The research concentration will be directed to the application of different AI methods within the UAV arena. By evaluating AI approaches, which will include Expert Systems, Neural Networks, Intelligent Agents, Fuzzy Logic, and Complex Adaptive Systems, a new insight may be gained into the benefits of AI techniques applied to achieving true autonomous operation of these systems thus providing new intellectual merit to this research field. The major area of discussion will be limited to the UAV. The systems of interest include small aircraft, insects, and miniature aircraft. Although flight systems will be explored, the benefits should apply to many Unmanned Vehicles such as: Rovers, Ocean Explorers, Robots, and autonomous operation systems. The flight system will be broken down into control agents that will represent the intelligent agent approach used in AI. After the completion of a successful approach, a framework of applying a Security Overseer will be added in an attempt to address errors, emergencies, failures, damage, or over dynamic environment. The chosen control problem was the landing phase of UAV operation. The initial results from simulation in FlightGear are presented.

Description: Effective delay of turbulent boundary layer separation could be achieved via closed-loop control. Constructing such a system requires that sensor data be processed, real-time, and fed into the controller to determine the output. Current methods for detection of turbulent boundary layer separation are lacking the capability of localized, fast and reliable identification of the boundary layer state. A method is proposed for short-time FFT processing of time series, measured by hot-film sensors, with the purpose of identifying the alternation of the balance between small and large scales as the boundary layer separates, favoring the large scales. The method has been validated by comparison to other criteria of separation detection and over a range of baseline and controlled flow conditions on a simplified high-lift system, incorporating active flow control.

Description: Results of an experimental investigation on isolated synthetic jets in crossflow from rectangular orifices of different aspect ratio and orientation are presented. Three aspect ratios, AR = 4, 8, and 16, with pitch a = 90deg., were investigated, Additionally, the AR = 8 case was pitched at 20 deg.. The yaw angle, , was varied through 0 deg., 10 deg., 45 deg. and 90 deg.. All orifices had same exit area and the data were compared with synthetic as well as steady jet from a circular orifice of same area. Hotwire measurements were performed to obtain all three components of mean velocity and turbulent stresses. Data were acquired for momentum-flux ratio up to J = 50. Distributions of time- and phase-averaged data were obtained on the cross sectional plane at x/D = 0.5, 5 and 10, as well as on the axial plane of the symmetry. Qualitative flowfield similarity between synthetic and steady jets is observed. However, high-momentum 'cap' above the low-momentum 'dome', characteristic of steady jets, does not necessarily appear in the synthetic jet. The position and shape of the high-momentum region depend on the distance from the orifice, pitch, yaw as well as momentum-flux ratio. Consequently, the location of the minimum velocity in the 'dome' measured at the plane of symmetry, y(submin), is adopted as a reference for penetration estimate and trajectory comparison. For AR = 16, the dome is the largest in area with maximum velocity deficit. However, the penetration is somewhat higher for AR = 4. Increase in yaw reduces the spatial extent of the dome and the penetration height but augments the velocity deficit. At low J the dome is connected to the boundary layer and traces of the cap of high momentum fluid are visible above it. Increase in J lifts the dome and reorganizes the high-momentum fluid around its perimeter, eventually bringing it underneath. Phase-averaged data document dynamic topological changes within the cycle. Phase-averaged streamwise velocity contours on the cross-sectional plane exhibit behavior commensurate with that seen in time-averaged data at various J.

Description: Analytical methods for stability analysis of large amplitude aircraft motion have been slow to develop because many nonlinear system stability assessment methods are restricted to a state-space dimension of less than three. The proffered approach is to create regional cell-to-cell maps for strategically located two-dimensional subspaces within the higher-dimensional model statespace. These regional solutions capture nonlinear behavior better than linearized point solutions. They also avoid the computational difficulties that emerge when attempting to create a cell map for the entire state-space. Example stability results are presented for a general aviation aircraft and a micro-aerial vehicle configuration. The analytical results are consistent with characteristics that were discovered during previous flight-testing.

Description: The NASA F-15 Intelligent Flight Control System project team has developed a series of flight control concepts designed to demonstrate the benefits of a neural network-based adaptive controller. The objective of the team is to develop and flight-test control systems that use neural network technology to optimize the performance of the aircraft under nominal conditions as well as stabilize the aircraft under failure conditions. Failure conditions include locked or failed control surfaces as well as unforeseen damage that might occur to the aircraft in flight. This report presents flight-test results for an adaptive controller using stability and control derivative values from an online learning neural network. A dynamic cell structure neural network is used in conjunction with a real-time parameter identification algorithm to estimate aerodynamic stability and control derivative increments to the baseline aerodynamic derivatives in flight. This set of open-loop flight tests was performed in preparation for a future phase of flights in which the learning neural network and parameter identification algorithm output would provide the flight controller with aerodynamic stability and control derivative updates in near real time. Two flight maneuvers are analyzed a pitch frequency sweep and an automated flight-test maneuver designed to optimally excite the parameter identification algorithm in all axes. Frequency responses generated from flight data are compared to those obtained from nonlinear simulation runs. An examination of flight data shows that addition of the flight-identified aerodynamic derivative increments into the simulation improved the pitch handling qualities of the aircraft.

Description: This paper presents the results of a computational fluid dynamics validation exercise using the General Aerodynamic Simulation Program (GASP) code for vehicles under consideration for the X-38 program. The scope of the exercise involved simulation and comparison with wind tunnel aeroheating and aerodynamic data in the hypersonic regime. In addition, simulations were performed and results are presented for the full-scale flight vehicle along an expected re- entry trajectory.

Description: A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.

Description: This final report summarizes the research accomplished under NASA-Ames Cooperative Agreement NCC2-5517 during the period which extended from August 1, 2002 to July 31,2004.

Description: A variety of novel control effector concepts have recently emerged that may enable new approaches to flight control. In particular, the potential exists to shift the composition of the typical aircraft control effector suite from a small number of high authority, specialized devices (rudder, aileron, elevator, flaps), toward larger numbers of smaller, less specialized, distributed device arrays. The concept envisions effector and sensor networks composed of relatively small high-bandwidth devices able to simultaneously perform a variety of control functions using feedback from disparate data sources. To investigate this concept, a remotely piloted flight vehicle has been equipped with an array of 24 trailing edge shape-change effectors and associated pressure measurements. The vehicle, called the Multifunctional Effector and Sensor Array (MESA) testbed, was recently tested in NASA Langley's 12-ft Low Speed wind tunnel to characterize its stability properties, control authorities, and distributed pressure sensitivities for use in a dynamic simulation prior to flight testing. Another objective was to implement and evaluate a scheme for actively controlling the spanwise pressure distribution using the shape-change array. This report describes the MESA testbed, design of the pressure distribution controller, and results of the wind tunnel test.

Description: Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in realtime, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.

Description: The motivation for Computational Aeroelasticity (CA) and the elements of one type of the analysis or simulation process are briefly reviewed. The need for streamlining and improving the overall process to reduce elapsed time and improve overall accuracy is discussed. Further effort is needed to establish the credibility of the methodology, obtain experience, and to incorporate the experience base to simplify the method for future use. Experience with the application of a variety of Computational Aeroelasticity programs is summarized for the transonic flutter of two wings, the AGARD 445.6 wing and a typical business jet wing. There is a compelling need for a broad range of additional flutter test cases for further comparisons. Some existing data sets that may offer CA challenges are presented.

Description: This paper presents novel analytical results for eigenvalues and eigenvectors produced using discrete time aerodynamic and aeroelastic models. An unsteady, incompressible vortex lattice aerodynamic model is formulated in discrete time; the importance of several modeling parameters is examined. A detailed study is made of the behavior of the aerodynamic eigenvalues both in discrete and continuous time. The aerodynamic model is then incorporated into aeroelastic equations of motion. Eigenanalyses of the coupled equations produce stability results and modal characteristics which are valid for critical and non-critical velocities. Insight into the modeling and physics associated with aeroelastic system behavior is gained by examining both the eigenvalues and the eigenvectors. Potential pitfalls in discrete time model construction and analysis are examined.

Description: A Momentum Source Model has been revised and implemented in the flow solver OVERFLOW-D. In this approach, the fan forces are evaluated from two-dimensional airfoil tables as a function of local Mach number and angle-of-attack and applied as source terms in the discretized Navier-Stokes equations. The model revisions include a new model for forces in the tip region and axial distribution of the source terms. The model revisions improve the results significantly. The Momentum Source Model agrees well with a discrete blade model for all computed collective pitch angles. The two models agree well with experimental data for thrust vs. torque. The Momentum Source Model is a good complement to Discrete Blade Models for ducted fan computations. The lower computational and labor costs make parametric studies, optimization studies and interactional aerodynamics studies feasible for cases beyond what is practical with a Discrete Blade Model today.

Description: The effect of nozzle trailing edge thickness on broadband acoustic radiation and the production of tones is investigated for coannular nozzles. Experiments were performed for a core nozzle trailing edge thickness between 0.38 mm and 3.17 mm. The on-set of discrete tones was found to be predominantly affected by the velocity ratio, the ratio of the fan velocity to the core velocity, although some dependency on trailing edge thickness was also noted. For a core nozzle trailing edge thickness greater than or equal to 0.89 mm, tones were produced for velocity ratios between 0.91 and 1.61. For a constant nozzle trailing edge thickness, the frequency varied almost linearly with the core velocity. The Strouhal number based on the core velocity changed with nozzle trailing edge thickness and varied between 0.16 and 0.2 for the core nozzles used in the experiments. Increases in broadband noise with increasing trailing edge thickness were observed for tone producing and non-tone producing conditions. A variable thickness trailing edge (crenellated) nozzle resulted in no tonal production and a reduction of the broadband trailing edge noise relative to that of the corresponding constant thickness trailing edge.

Description: Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter. Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parametrics included angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.3 x 10(exp 6) to 3.0 x 10(exp 6) per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), deformation of the wing windward surface, and main landing gear and/or door deployment. The measured aerodynamic increments for the damage scenarios examined were generally small in magnitude, as were the flight-derived values during most of the entry prior to loss of communication. A progressive damage scenario is presented that qualitatively matches the flight observations for the STS-107 entry.

Description: The main objective of this work is to predict the Aerothermal Performance Constraint (APC) for a class of Crew Transfer Vehicles (CTV) with shap noses and wing leading edges made out of UHTC which is a family of Ultra High Temperature Ceramics materials developed at NASA Ames. The APC is based on the theoretical temperature limit of the material which is usually encountered at the CTV nose or wing leading edge. The APC places a lower limit on the trajectory of the CTV in the altitude velocity space. The APC is used as one of the constraints in developing reentry and abort trajectories for the CTV. The trajectories are then used to generate transient thermal response of the nosecaps and wing leading edges which are represented as either a one piece of UHTC or two piece (UHTC + RCC) with perfect axial contact. The final paper will include more details about the analysis procedure and will also include results for reentry and abort design trajectories.

Description: A collection of computer programs has been developed that implements a procedure for predicting multiple-pure-tone (MPT) noise generated by fan blades of an aircraft engine (e.g., a turbofan engine). MPT noise arises when the fan is operating with supersonic relative tip Mach No. Under this flow condition, there is a strong upstream running shock. The strength and position of this shock are very sensitive to blade geometry variations. For a fan where all the blades are identical, the primary tone observed upstream of the fan will be the blade passing frequency. If there are small variations in geometry between blades, then tones below the blade passing frequency arise MPTs. Stagger angle differences as small as 0.1 can give rise to significant MPT. It is also noted that MPT noise is more pronounced when the fan is operating in an unstarted mode. Computational results using a three-dimensional flow solver to compute the complete annulus flow with non-uniform fans indicate that MPT noise can be estimated in a relatively simple way. Hence, once the effect of a typical geometry variation of one blade in an otherwise uniform blade row is known, the effect of all the blades being different can be quickly computed via superposition. Two computer programs that were developed as part of this work are used in conjunction with a user s computational fluid dynamics (CFD) code to predict MPT spectra for a fan with a specified set of geometric variations: (1) The first program ROTBLD reads the users CFD solution files for a single blade passage via an API (Application Program Interface). There are options to replicate and perturb the geometry with typical variations stagger, camber, thickness, and pitch. The multi-passage CFD solution files are then written in the user s file format using the API. (2) The second program SUPERPOSE requires two input files: the first is the circumferential upstream pressure distribution extracted from the CFD solution on the multi-passage mesh, the second file defines the geometry variations of each blade in a complete fan. Superposition is used to predict the spectra resulting from the geometric variations.

Description: Boundary Layer separation causes the airfoil to stall and therefore imposes dramatic performance degradation on the airfoil. In recent years, flow separation control has been one of the active research areas in the field of aerodynamics due to its promising performance improvements on the lifting device. These active flow separation control techniques include steady and unsteady air injection as well as suction on the airfoil surface etc. This paper will be focusing on the steady and unsteady air injection on the airfoil. Although wind tunnel experiments revealed that the performance improvements on the airfoil using injection techniques, the details of how the key variables such as air injection slot geometry and air injection angle etc impact the effectiveness of flow separation control via air injection has not been studied. A parametric study of both steady and unsteady air injection active flow control will be the main objective for this summer. For steady injection, the key variables include the slot geometry, orientation, spacing, air injection velocity as well as the injection angle. For unsteady injection, the injection frequency will also be investigated. Key metrics such as lift coefficient, drag coefficient, total pressure loss and total injection mass will be used to measure the effectiveness of the control technique. A design of experiments using the Box-Behnken Design is set up in order to determine how each of the variables affects each of the key metrics. Design of experiment is used so that the number of experimental runs will be at minimum and still be able to predict which variables are the key contributors to the responses. The experiments will then be conducted in the 1ft by 1ft wind tunnel according to the design of experiment settings. The data obtained from the experiments will be imported into JMP, statistical software, to generate sets of response surface equations which represent the statistical empirical model for each of the metrics as a function of the key variables. Next, the variables such as the slot geometry can be optimized using the build-in optimizer within JMP. Finally, a wind tunnel testing will be conducted using the optimized slot geometry and other key variables to verify the empirical statistical model. The long term goal for this effort is to assess the impacts of active flow control using air injection at system level as one of the task plan included in the NASAs URETI program with Georgia Institute of Technology.

Description: A performance evaluation of the Hybrid Predictor corrector Aerocapture Scheme (HYPAS) guidance algorithm for aerocapture at Neptune is presented in this paper for a Mission to Neptune and the Neptune moon Triton'. This mission has several challenges not experienced in previous aerocapture guidance assessments. These challengers are a very high Neptune arrival speed, atmospheric exit into a high energy orbit about Neptune, and a very high ballistic coefficient that results in a low altitude acceleration capability when combined with the aeroshell LD. The evaluation includes a definition of the entry corridor, a comparison to the theoretical optimum performance, and guidance responses to variations in atmospheric density, aerodynamic coefficients and flight path angle for various vehicle configurations (ballistic numbers). The benefits of utilizing angle-of-attack modulation in addition to bank angle modulation to improve flight performance is also discussed. The results show that despite large sensitivities in apoapsis targeting, the algorithm performs within the allocated AV budget for the Neptune mission bank angle only modulation. The addition of angle-of-attack modulation with as little as 5 degrees of amplitude significantly improves the scatter in final orbit apoapsis. Although the angle-of-attack modulation complicates the vehicle design, the performance enhancement reduces aerocapture risk and reduces the propellant consumption needed to reach the high energy target orbit for a conventional propulsion system.

Description: NASA's Strategic Plan for the Aerospace Technology Enterprise includes ambitious objectives focused on affordable air travel, reduced emissions, and expanded aviation-system capacity. NASA Dryden Flight Research Center, in cooperation with NASA Ames Research Center, the Boeing Company, and the University of California, Los Angeles, has embarked on an autonomous-formation-flight project that promises to make significant strides towards these goals. For millions of years, birds have taken advantage of the aerodynamic benefit of flying in formation. The traditional "V" formation flown by many species of birds (including gulls, pelicans, and geese) enables each of the trailing birds to fly in the upwash flow field that exists just outboard of the bird immediately ahead in the formation. The result for each trailing bird is a decrease in induced drag and thus a reduction in the energy needed to maintain a given speed. Hence, for migratory birds, formation flight extends the range of the system of birds over the range of birds flying solo. The Autonomous Formation Flight (AFF) Project is seeking to extend this symbiotic relationship to aircraft.

Description: Active flow control devices including mass injection systems and zero-net-mass flux actuators (synthetic jets) have been employed to delay flow separation. These devices are capable of interacting with low-speed, subsonic flows, but situations exist where a stronger crossflow interaction is needed. Small actuators that utilize detonation of premixed fuel and oxidizer should be capable of producing supersonic exit jet velocities. An actuator producing exit velocities of this magnitude should provide a more significant interaction with transonic and supersonic crossflows. This concept would be applicable to airfoils on high-speed aircraft as well as inlet and diffuser flow control. The present work consists of the development of a detonation actuator capable of producing a detonation in a single shot (one cycle). Multiple actuator configurations, initial fill pressures, oxidizers, equivalence ratios, ignition energies, and the addition of a turbulence generating device were considered experimentally and computationally. It was found that increased initial fill pressures and the addition of a turbulence generator aided in the detonation process. The actuators successfully produced Chapman-Jouguet detonations and wave speeds on the order of 3000 m/s.

Description: This final report will document the accomplishments of the work of this project. 1) The incremental-iterative (II) form of the reverse-mode (adjoint) method for computing first-order (FO) aerodynamic sensitivity derivatives (SDs) has been successfully implemented and tested in a 2D CFD code (called ANSERS) using the reverse-mode capability of ADIFOR 3.0. These preceding results compared very well with similar SDS computed via a black-box (BB) application of the reverse-mode capability of ADIFOR 3.0, and also with similar SDs calculated via the method of finite differences. 2) Second-order (SO) SDs have been implemented in the 2D ASNWERS code using the very efficient strategy that was originally proposed (but not previously tested) of Reference 3, Appendix A. Furthermore, these SO SOs have been validated for accuracy and computational efficiency. 3) Studies were conducted in Quasi-1D and 2D concerning the smoothness (or lack of smoothness) of the FO and SO SD's for flows with shock waves. The phenomenon is documented in the publications of this study (listed subsequently), however, the specific numerical mechanism which is responsible for this unsmoothness phenomenon was not discovered. 4) The FO and SO derivatives for Quasi-1D and 2D flows were applied to predict aerodynamic design uncertainties, and were also applied in robust design optimization studies.

Description: A study to determine a limiting distance to span ratio for the extrapolation of near-field pressure signatures is described and discussed. This study was to be done in two wind-tunnel facilities with two wind-tunnel models. At this time, only the first half had been completed, so the scope of this report is limited to the design of the models, and to an analysis of the first set of measured pressure signatures. The results from this analysis showed that the pressure signatures measured at separation distances of 2 to 5 span lengths did not show the desired low-boom shapes. However, there were indications that the pressure signature shapes were becoming 'flat-topped'. This trend toward a 'flat-top' pressure signatures shape was seen to be a gradual one at the distance ratios employed in this first series of wind-tunnel tests.

Description: Closed-form direct and inverse kinematics of a new three degree-of-freedom (DOF) parallel manipulator with inextensible limbs and base-mounted actuators are presented. The manipulator has higher resolution and precision than the existing three DOF mechanisms with extensible limbs. Since all of the manipulator actuators are base-mounted; higher payload capacity, smaller actuator sizes, and lower power dissipation can be obtained. The manipulator is suitable for alignment applications where only tip, tilt, and piston motions are significant. The direct kinematics of the manipulator is reduced to solving an eighth-degree polynomial in the square of tangent of half-angle between one of the limbs and the base plane. Hence, there are at most 16 assembly configurations for the manipulator. In addition, it is shown that the 16 solutions are eight pairs of reflected configurations with respect to the base plane. Numerical examples for the direct and inverse kinematics of the manipulator are also presented.

Description: A detailed boundary layer investigation of an NACA 0012 airfoil was made by applying the WIND code at the NASA Glenn Research Center. Structured C-grids of selected configurations were fitted around the airfoil. Grids with dense spacing to capture the boundary layer provided for obtaining accurate results in this region. Concurrent with the boundary layer investigation, contour plots of overall performance of the airfoil were also obtained. A Mach number of 0.80, a Reynolds number of 4,000,000, and a zero angle of attack were set to correspond with numerical and experimental investigations reported in the literature, for comparison with the overall performance. The results were favorable for both the boundary layer investigations and the performance. As WIND evolves, it is intended to run WIND using unstructured grids to compare such results with the current structured grid results.

Description: Langley Stability and Transition Analysis Code (LASTRAC) is a general-purpose, physics-based transition prediction code released by NASA for laminar flow control studies and transition research. This paper describes the LASTRAC extension to general three-dimensional (3D) boundary layers such as finite swept wings, cones, or bodies at an angle of attack. The stability problem is formulated by using a body-fitted nonorthogonal curvilinear coordinate system constructed on the body surface. The nonorthogonal coordinate system offers a variety of marching paths and spanwise waveforms. In the extreme case of an infinite swept wing boundary layer, marching with a nonorthogonal coordinate produces identical solutions to those obtained with an orthogonal coordinate system using the earlier release of LASTRAC. Several methods to formulate the 3D parabolized stability equations (PSE) are discussed. A surface-marching procedure akin to that for 3D boundary layer equations may be used to solve the 3D parabolized disturbance equations. On the other hand, the local line-marching PSE method, formulated as an easy extension from its 2D counterpart and capable of handling the spanwise mean flow and disturbance variation, offers an alternative. A linear stability theory or parabolized stability equations based N-factor analysis carried out along the streamline direction with a fixed wavelength and downstream-varying spanwise direction constitutes an efficient engineering approach to study instability wave evolution in a 3D boundary layer. The surface-marching PSE method enables a consistent treatment of the disturbance evolution along both streamwise and spanwise directions but requires more stringent initial conditions. Both PSE methods and the traditional LST approach are implemented in the LASTRAC.3d code. Several test cases for tapered or finite swept wings and cones at an angle of attack are discussed.

Description: The Linear Autoland Simulink model was created to be a modular test environment for testing of control system components in commercial aircraft. The input variables, physical laws, and referenced frames used are summarized. The state space theory underlying the model is surveyed and the location of the control actuators described. The equations used to realize the Dryden gust model to simulate winds and gusts are derived. A description of the pseudo-random number generation method used in the wind gust model is included. The longitudinal autopilot, lateral autopilot, automatic throttle autopilot, engine model and automatic trim devices are considered as subsystems. The experience in converting the Airlabs FORTRAN aircraft control system simulation to a graphical simulation tool (Matlab/Simulink) is described.

Description: A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Satisfactory global calibrations of the PSP were obtained at =0.70, 0.90, and 1.20, angles of attack from 10 degrees to 20 degrees, and angles of sideslip of 0 and 2.5 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at 57 discrete locations on the model. Both techniques clearly revealed the significant influence on the surface pressure distributions of the vortices shed from the sharp, chine-like leading edges. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M infinity =0.70 and 2.6 percent at M infinity =0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

Description: A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Global PSP calibrations were obtained using an in-situ method featuring the simultaneous electronically-scanned pressures (ESP) measurements. Both techniques revealed the significant influence leading-edge vortices on the surface pressure distributions. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M(sub infinity)=0.70 and 2.6 percent at M(sub infinity)=0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

Description: Time scales in turbulent and transitional flow provide a link between experimental data and modeling, both in terms of physical content and for quantitative assessment. The problem of interest here is the definition of time scales in an unsteady flow. Using representative samples of data from GEAE low pressure turbine experiment in low speed research turbine facility with wake-induced transition, we document several methods to extract dominant frequencies, and compare the results. We show that conventional methods of time scale evaluation (based on autocorrelation functions and on Fourier spectra) and wavelet-based methods provide similar information when applied to stationary signals. We also show the greater flexibility of the wavelet-based methods when dealing with intermittent or strongly modulated data, as are encountered in transitioning boundary layers and in flows with unsteady forcing associated with wake passing. We define phase-averaged dominant frequencies that characterize the turbulence associated with freestream conditions and with the passing wakes downstream of a rotor. The relevance of these results for modeling is discussed in the paper.

Description: Autonomous Flight Safety System (AFSS) is an independent flight safety system designed for small to medium sized expendable launch vehicles launching from or needing range safety protection while overlying relatively remote locations. AFSS replaces the need for a man-in-the-loop to make decisions for flight termination. AFSS could also serve as the prototype for an autonomous manned flight crew escape advisory system. AFSS utilizes onboard sensors and processors to emulate the human decision-making process using rule-based software logic and can dramatically reduce safety response time during critical launch phases. The Range Safety flight path nominal trajectory, its deviation allowances, limit zones and other flight safety rules are stored in the onboard computers. Position, velocity and attitude data obtained from onboard global positioning system (GPS) and inertial navigation system (INS) sensors are compared with these rules to determine the appropriate action to ensure that people and property are not jeopardized. The final system will be fully redundant and independent with multiple processors, sensors, and dead man switches to prevent inadvertent flight termination. AFSS is currently in Phase III which includes updated algorithms, integrated GPS/INS sensors, large scale simulation testing and initial aircraft flight testing.

Description: A computational investigation was performed to assess the effectiveness of Detached-Eddy Simulation (DES) as a tool for predicting icing effects. The AVUS code, a public domain flow solver, was employed to compute solutions for an iced wing configuration using DES and steady Reynolds Averaged Navier-Stokes (RANS) equation methodologies. The configuration was an extruded GLC305/944-ice shape section with a rectangular planform. The model was mounted between two walls so no tip effects were considered. The numerical results were validated by comparison with experimental data for the same configuration. The time-averaged DES computations showed some improvement in lift and drag results near stall when compared to steady RANS results. However, comparisons of the flow field details did not show the level of agreement suggested by the integrated quantities. Based on our results, we believe that DES may prove useful in a limited sense to provide analysis of iced wing configurations when there is significant flow separation, e.g., near stall, where steady RANS computations are demonstrably ineffective. However, more validation is needed to determine what role DES can play as part of an overall icing effects prediction strategy. We conclude the report with an assessment of existing computational tools for application to the iced wing problem and a discussion of issues that merit further study.

Description: The unprecedented advances being made in computational fluid dynamic (CFD) technology have demonstrated the powerful capabilities of codes in applications to civil and military aircraft. Used in conjunction with wind-tunnel and flight investigations, many codes are now routinely used by designers in diverse applications such as aerodynamic performance predictions and propulsion integration. Typically, these codes are most reliable for attached, steady, and predominantly turbulent flows. As a result of increasing reliability and confidence in CFD, wind-tunnel testing for some new configurations has been substantially reduced in key areas, such as wing trade studies for mission performance guarantees. Interest is now growing in the application of computational methods to other critical design challenges. One of the most important disciplinary elements for civil and military aircraft is prediction of stability and control characteristics. CFD offers the potential for significantly increasing the basic understanding, prediction, and control of flow phenomena associated with requirements for satisfactory aircraft handling characteristics.

Description: The Numerical Propulsion System Simulation (NPSS), an advanced engineering simulation environment used to design and analyze aircraft engines, has been enhanced by integrating control development tools into it. One of these tools is a generic controller interface that allows NPSS to communicate with control development software environments such as MATLAB and EASY5. The other tool is a linear model generator (LMG) that gives NPSS the ability to generate linear, time-invariant state-space models. Integrating these tools into NPSS enables it to be used for control system development. This paper will discuss the development and integration of these tools into NPSS. In addition, it will show a comparison of transient model results of a generic, dual-spool, military-type engine model that has been implemented in NPSS and Simulink. It will also show the linear model generator s ability to approximate the dynamics of a nonlinear NPSS engine model.

Description: This report presents results of the work completed on the preliminary design of Fan 3 of NASA s 22-inch Fan Low Noise Research project. Fan 3 was intended to build on the experience gained from Fans 1 and 2 by demonstrating noise reduction technology that surpasses 1992 levels by 6 dB. The work was performed as part of NASA s Advanced Subsonic Technology (AST) program. Work on this task was conducted in the areas of CFD code validation, acoustic prediction and validation, rotor parametric studies, and fan exit guide vane (FEGV) studies up to the time when a NASA decision was made to cancel the design, fabrication and testing phases of the work. The scope of the program changed accordingly to concentrate on two subtasks: (1) Rig data analysis and CFD code validation and (2) Fan and FEGV optimization studies. The results of the CFD code validation work showed that this tool predicts 3D flowfield features well from the blade trailing edge to about a chord downstream. The CFD tool loses accuracy as the distance from the trailing edge increases beyond a blade chord. The comparisons of noise predictions to rig test data showed that both the tone noise tool and the broadband noise tool demonstrated reasonable agreement with the data to the degree that these tools can reliably be used for design work. The section on rig airflow and inlet separation analysis describes the method used to determine total fan airflow, shows the good agreement of predicted boundary layer profiles to measured profiles, and shows separation angles of attack ranging from 29.5 to 27deg for the range of airflows tested. The results of the rotor parametric studies were significant in leading to the decision not to pursue a new rotor design for Fan 3 and resulted in recommendations to concentrate efforts on FEGV stator designs. The ensuing parametric study on FEGV designs showed the potential for 8 to 10 EPNdB noise reduction relative to the baseline.

Description: Present simulation of turbulent flows involving shock wave/boundary layer interaction invariably overestimates heat flux by almost a factor of two. One possible reason for such a performance is a result of the fact that the turbulence models employed make use of Morkovin's hypothesis. This hypothesis is valid for non-hypersonic Mach numbers and moderate rates of heat transfer. At hypersonic Mach numbers, high rates of heat transfer exist in regions where shock wave/boundary layer interactions are important. As a result, one should not expect traditional turbulence models to yield accurate results. The goal of this investigation is to explore the role of a variable Prandtl number formulation in predicting heat flux in flows dominated by strong shock wave/boundary layer interactions. The intended applications involve external flows in the absence of combustion such as those encountered in supersonic inlets. This can be achieved by adding equations for the temperature variance and its dissipation rate. Such equations can be derived from the exact Navier-Stokes equations. Traditionally, modeled equations are based on the low speed energy equation where the pressure gradient term and the term responsible for energy dissipation are ignored. It is clear that such assumptions are not valid for hypersonic flows. The approach used here is based on the procedure used in deriving the k-zeta model, in which the exact equations that governed k, the variance of velocity, and zeta, the variance of vorticity, were derived and modeled. For the variable turbulent Prandtl number, the exact equations that govern the temperature variance and its dissipation rate are derived and modeled term by term. The resulting set of equations are free of damping and wall functions and are coordinate-system independent. Moreover, modeled correlations are tensorially consistent and invariant under Galilean transformation. The final set of equations will be given in the paper.

Description: A mode choice model that generates on-demand air travel forecasts at a set of GA airports based on changes in economic characteristics, vehicle performance characteristics such as speed and cost, and demographic trends has been integrated with a model to generate itinerate aircraft operations by airplane category at a set of 3227 airports. Numerous intermediate outputs can be generated, such as the number of additional trips diverted from automobiles and schedule air by the improved performance and cost of on-demand air vehicles. The total number of transported passenger miles that are diverted is also available. From these results the number of new aircraft to service the increased demand can be calculated. Output from the models discussed is in the format to generate the origin and destination traffic flow between the 3227 airports based on solutions to a gravity model.

Description: Work on stability and control included the following reports:Introductory Remarks; Introduction to Computational Methods for Stability and Control (COMSAC); Stability & Control Challenges for COMSAC: a NASA Langley Perspective; Emerging CFD Capabilities and Outlook A NASA Langley Perspective; The Role for Computational Fluid Dynamics for Stability and Control:Is it Time?; Northrop Grumman Perspective on COMSAC; Boeing Integrated Defense Systems Perspective on COMSAC; Computational Methods in Stability and Control:WPAFB Perspective; Perspective: Raytheon Aircraft Company; A Greybeard's View of the State of Aerodynamic Prediction; Computational Methods for Stability and Control: A Perspective; Boeing TacAir Stability and Control Issues for Computational Fluid Dynamics; NAVAIR S&C Issues for CFD; An S&C Perspective on CFD; Issues, Challenges & Payoffs: A Boeing User s Perspective on CFD for S&C; and Stability and Control in Computational Simulations for Conceptual and Preliminary Design: the Past, Today, and Future?

Description: This paper describes the avian wing geometry (Seagull, Merganser, Teal and Owl) extracted from non-contact surface measurements using a three-dimensional laser scanner. The geometric quantities, including the camber line and thickness distribution of airfoil, wing planform, chord distribution, and twist distribution, are given in convenient analytical expressions. Thus, the avian wing surfaces can be generated and the wing kinematics can be simulated. The aerodynamic characteristics of avian airfoils in steady inviscid flows are briefly discussed. The avian wing kinematics is recovered from videos of three level-flying birds (Crane, Seagull and Goose) based on a two-jointed arm model. A flapping seagull wing in the 3D physical space is re-constructed from the extracted wing geometry and kinematics.

Description: This report describes a guidance system for agile vehicles based on a hybrid closed-loop model of the vehicle dynamics. The hybrid model represents the vehicle dynamics through a combination of linear-time-invariant control modes and pre-programmed, finite-duration maneuvers. This particular hybrid structure can be realized through a control system that combines trim controllers and a maneuvering control logic. The former enable precise trajectory tracking, and the latter enables trajectories at the edge of the vehicle capabilities. The closed-loop model is much simpler than the full vehicle equations of motion, yet it can capture a broad range of dynamic behaviors. It also supports a consistent link between the physical layer and the decision-making layer. The trajectory generation was formulated as an optimization problem using mixed-integer-linear-programming. The optimization is solved in a receding horizon fashion. Several techniques to improve the computational tractability were investigate. Simulation experiments using NASA Ames 'R-50 model show that this approach fully exploits the vehicle's agility.

Description: Several flight-critical aerodynamic problems such as buffet, flutter, stall, and wing rock are strongly affected or caused by abrupt changes in unsteady aerodynamic loads and moments. Advanced sensing and flow diagnostic techniques have made possible simultaneous identification and tracking, in real-time, of the critical surface, viscosity-related aerodynamic phenomena under both steady and unsteady flight conditions. The wind tunnel study reported here correlates surface hot-film measurements of leading edge stagnation point and separation point, with unsteady aerodynamic loads on a NACA 0015 airfoil. Lift predicted from the correlation model matches lift obtained from pressure sensors for an airfoil undergoing harmonic pitchup and pitchdown motions. An analytical model was developed that demonstrates expected stall trends for pitchup and pitchdown motions. This report demonstrates an ability to obtain unsteady aerodynamic loads in real-time, which could lead to advances in air vehicle safety, performance, ride-quality, control, and health management.

Description: A genetic algorithm approach suitable for solving multi-objective optimization problems is described and evaluated using a series of aerodynamic shape optimization problems. Several new features including two variations of a binning selection algorithm and a gene-space transformation procedure are included. The genetic algorithm is suitable for finding pareto optimal solutions in search spaces that are defined by any number of genes and that contain any number of local extrema. A new masking array capability is included allowing any gene or gene subset to be eliminated as decision variables from the design space. This allows determination of the effect of a single gene or gene subset on the pareto optimal solution. Results indicate that the genetic algorithm optimization approach is flexible in application and reliable. The binning selection algorithms generally provide pareto front quality enhancements and moderate convergence efficiency improvements for most of the problems solved.

Description: The aerodynamic drag of an aerial refueling assembly was calculated during the Automated Aerial Refueling project at the NASA Dryden Flight Research Center. An F/A-18A airplane was specially instrumented to obtain accurate fuel flow measurements and to determine engine thrust. A standard Navy air refueling store with a retractable refueling hose and paradrogue was mounted to the centerline pylon of the F/A-18A airplane. As the paradrogue assembly was deployed and stowed, changes in the calculated thrust of the airplane occurred and were equated to changes in vehicle drag. These drag changes were attributable to the drag of the paradrogue assembly. The drag of the paradrogue assembly was determined to range from 200 to 450 lbf at airspeeds from 170 to 250 KIAS. Analysis of the drag data resulted in a single drag coefficient of 0.0056 for the paradrogue assembly that adequately matched the calculated drag for all flight conditions. The drag relief provided to the tanker airplane when a receiver airplane engaged the paradrogue is also documented from 35 to 270 lbf at the various flight conditions tested. The results support the development of accurate aerodynamic models to be used in refueling simulations and control laws for fully autonomous refueling.

Description: Force and moment measurements from an F-16XL during forced pitch oscillation tests result in dynamic stability derivatives, which are measured in combinations. Initial computational simulations of the motions and combined derivatives are attempted via a low-order, time-dependent panel method computational fluid dynamics code. The code dynamics are shown to be highly questionable for this application and the chosen configuration. However, three methods to computationally separate such combined dynamic stability derivatives are proposed. One of the separation techniques is demonstrated on the measured forced pitch oscillation data. Extensions of the separation techniques to yawing and rolling motions are discussed. In addition, the possibility of considering the angles of attack and sideslip state vector elements as distributed quantities, rather than point quantities, is introduced.

Description: In this paper, an approach for in-flight fault detection and isolation (FDI) of aircraft engine sensors based on a bank of Kalman filters is developed. This approach utilizes multiple Kalman filters, each of which is designed based on a specific fault hypothesis. When the propulsion system experiences a fault, only one Kalman filter with the correct hypothesis is able to maintain the nominal estimation performance. Based on this knowledge, the isolation of faults is achieved. Since the propulsion system may experience component and actuator faults as well, a sensor FDI system must be robust in terms of avoiding misclassifications of any anomalies. The proposed approach utilizes a bank of (m+1) Kalman filters where m is the number of sensors being monitored. One Kalman filter is used for the detection of component and actuator faults while each of the other m filters detects a fault in a specific sensor. With this setup, the overall robustness of the sensor FDI system to anomalies is enhanced. Moreover, numerous component fault events can be accounted for by the FDI system. The sensor FDI system is applied to a commercial aircraft engine simulation, and its performance is evaluated at multiple power settings at a cruise operating point using various fault scenarios.

Description: Development of improved modeling methods to provide increased fidelity of flight predictions for aircraft motions during flight in flow regimes with large nonlinearities requires improvements in test techniques for measuring and characterizing wind tunnel data. This paper presents a method for providing a measure of data integrity for static and forced oscillation test techniques. Data integrity is particularly important when attempting to accurately model and predict flight of today s high performance aircraft which are operating in expanded flight envelopes, often maneuvering at high angular rates at high angles-of-attack, even above maximum lift. Current aerodynamic models are inadequate in predicting flight characteristics in the expanded envelope, such as rapid aircraft departures and other unusual motions. Present wind tunnel test methods do not factor changes of flow physics into data acquisition schemes, so in many cases data are obtained over more iterations than required, or insufficient data may be obtained to determine a valid estimate with statistical significance. Additionally, forced oscillation test techniques, one of the primary tools used to develop dynamic models, do not currently provide estimates of the uncertainty of the results during an oscillation cycle. A method to optimize the required number of forced oscillation cycles based on decay of uncertainty gradients and balance tolerances is also presented.

Description: An overview of basic aeroheating design issues for Reusable Launch Vehicles (RLV), which addresses the application of hypersonic ground-based testing, and computational fluid dynamic (CFD) and engineering codes, is presented. Challenges inherent to the prediction of aeroheating environments required for the successful design of the RLV Thermal Protection System (TPS) are discussed in conjunction with the importance of employing appropriate experimental/computational tools. The impact of the information garnered by using these tools in the resulting analyses, ultimately enhancing the RLV TPS design is illustrated. A wide range of topics is presented in this overview; e.g. the impact of flow physics issues such as boundary-layer transition, including effects of distributed and discrete roughness, shockshock interactions, and flow separation/reattachment. Also, the benefit of integrating experimental and computational studies to gain an improved understanding of flow phenomena is illustrated. From computational studies, the effect of low-density conditions and of uncertainties in material surface properties on the computed heating rates are highlighted as well as the significant role of CFD in improving the Outer Mold Line (OML) definition to reduce aeroheating while maintaining aerodynamic performance. Appropriate selection of the TPS design trajectories and trajectory shaping to mitigate aeroheating levels and loads are discussed. Lastly, an illustration of an aeroheating design process is presented whereby data from hypersonic wind-tunnel tests are integrated with predictions from CFD codes and engineering methods to provide heating environments along an entry trajectory as required for TPS design.

Description: This paper is a description of the analysis of blockage corrections for bodies of revolution for the slotted-wall configuration of the National Transonic Facility (NTF) at the NASA Langley Research Center (LaRC). A wall correction method based on the measured wall signature is used. Test data from three different-sized blockage bodies and four wall ventilation settings were analyzed at various Mach numbers and unit Reynolds numbers. The results indicate that with the proper selection of the boundary condition parameters, the wall correction method can predict blockage corrections consistent with the wall measurements for Mach numbers as high as 0.95.

Description: A multigrid solver is defined as having textbook multigrid efficiency (TME) if the solutions to the governing system of equations are attained in a computational work which is a small (less than 10) multiple of the operation count in evaluating the discrete residuals. TME in solving the incompressible inviscid fluid equations is demonstrated for leading- edge stagnation flows. The contributions of this paper include (1) a special formulation of the boundary conditions near stagnation allowing convergence of the Newton iterations on coarse grids, (2) the boundary relaxation technique to facilitate relaxation and residual restriction near the boundaries, (3) a modified relaxation scheme to prevent initial error amplification, and (4) new general analysis techniques for multigrid solvers. Convergence of algebraic errors below the level of discretization errors is attained by a full multigrid (FMG) solver with one full approximation scheme (F.4S) cycle per grid. Asymptotic convergence rates of the F.4S cycles for the full system of flow equations are very fast, approaching those for scalar elliptic equations.

Description: The objective of this paper is to report the results from the research being conducted in reconfigurable fight controls at NASA Ames. A study was conducted with three NASA Dryden test pilots to evaluate two approaches of reconfiguring an aircraft's control system when failures occur in the control surfaces and engine. NASA Ames is investigating both a Neural Generalized Predictive Control scheme and a Neural Network based Dynamic Inverse controller. This paper highlights the Predictive Control scheme where a simple augmentation to reduce zero steady-state error led to the neural network predictor model becoming redundant for the task. Instead of using a neural network predictor model, a nominal single point linear model was used and then augmented with an error corrector. This paper shows that the Generalized Predictive Controller and the Dynamic Inverse Neural Network controller perform equally well at reconfiguration, but with less rate requirements from the actuators. Also presented are the pilot ratings for each controller for various failure scenarios and two samples of the required control actuation during reconfiguration. Finally, the paper concludes by stepping through the Generalized Predictive Control's reconfiguration process for an elevator failure.

Description: A structured-grid quality measure is proposed, combining three traditional measurements: intersection angles, stretching, and curvature. Quality assesses whether the grid generated provides the best possible tradeoffs in grid stretching and skewness that enable accurate flow predictions, whereas the grid density is assumed to be a constraint imposed by the available computational resources and the desired resolution of the flow field. The usefulness of this quality measure is assessed by comparing heat transfer predictions from grid convergence studies for grids of varying quality in the range of [0.6-0.8] on an 8'half-angle sphere-cone, at laminar, perfect gas, Mach 10 wind tunnel conditions.

Description: Development of a general aerodynamic model that is adequate for predicting the forces and moments in the nonlinear and unsteady portions of the flight envelope has not been accomplished to a satisfactory degree. Predicting aerodynamic response during arbitrary motion of an aircraft over the complete flight envelope requires further development of the mathematical model and the associated methods for ground-based testing in order to allow identification of the model. In this study, a general nonlinear unsteady aerodynamic model is presented, followed by a summary of a linear modeling methodology that includes test and identification methods, and then a progressive series of steps suggesting a roadmap to develop a general nonlinear methodology that defines modeling, testing, and identification methods. Initial steps of the general methodology were applied to static and oscillatory test data to identify rolling-moment coefficient. Static measurements uncovered complicated dependencies of the aerodynamic coefficient on angle of attack and sideslip in the stall region making it difficult to find a simple analytical expression for the measurement data. In order to assess the effect of sideslip on the damping and unsteady terms, oscillatory tests in roll were conducted at different values of an initial offset in sideslip. Candidate runs for analyses were selected where higher order harmonics were required for the model and where in-phase and out-of-phase components varied with frequency. From these results it was found that only data in the angle-of-attack range of 35 degrees to 37.5 degrees met these requirements. From the limited results it was observed that the identified models fit the data well and both the damping-in-roll and the unsteady term gain are decreasing with increasing sideslip and motion amplitude. Limited similarity between parameter values in the nonlinear model and the linear model suggest that identifiability of parameters in both terms may be a problem. However, the proposed methodology can still be used with careful experiment design and carefully selected values of angle of attack, sideslip, amplitude, and frequency of the oscillatory data.

Description: During the past decade, piezoelectric actuators as the active element in synthetic jets demonstrated that they could significantly enhance the overall lift on an airfoil. However, durability, system weight, size, and power have limited their use outside a laboratory. These problems are not trivial, since piezoelectric actuators are physically brittle and display limited displacement. The objective of this study is to characterize the relevant properties for the design of a synthetic jet utilizing three types of piezoelectric actuators as mechanical diaphragms, Radial Field Diaphragms, Thunders, and Bimorphs so that the shape cavity volume does not exceed 147.5 cubic centimeters on a 7centimeter x 7centimeter aerial coverage. These piezoelectric elements were selected because of their geometry, and overall free-displacement. Each actuator was affixed about its perimeter in a cavity, and relevant parameters such as clamped displacement variations with voltage and frequency, air velocities produced through an aperture, and sound pressure levels produced by the piezoelectric diaphragms were measured.

Description: The ability of the Laura flow solver to predict local heating augmentation factors for shallow cavities is assessed. This assessment is part of a larger e ort within the Space Shuttle return-to-flight program to develop technologies to support on-orbit tile repair decisions. The comparison is made against global phosphor thermography images taken in the Langley Aerothermodynamic Laboratory 20-Inch Mach 6 Air Tunnel. The cavities are rectangular in shape, with lengths L/H of 14 20 and depths H/ of 1.1 5.2. The fully laminar results, for Re = 300, show good agreement between the data sets. For Re = 503, the wind tunnel data indicates boundary layer transition with turbulent flow both within and downstream of the cavity. The turbulent flow structures are significantly di erent from the laminar predictions, with order of magnitude increases in the heating augmentations. Because of the di erent flow structures, no simple bump factor can be used to correct the laminar calculations to account for the turbulent heating levels. A fine gradation in wind tunnel cases will be required to clearly delineate the laminar-to-turbulent transition point, and hence the limits of applicability of the laminar numerical approach.

Description: A simple optical technique for the structural characterization of lightweight space columns is presented. The technique is useful for determining the coefficient of thermal expansion during cool down as well as the induced strain during tension and compression testing. The technique is based upon object-to-image plane scaling and does not require any photogrammetric calibrations or computations. Examples of the measurement of the coefficient of thermal expansion are presented for several lightweight space columns. Examples of strain measured during tension and compression testing are presented along with comparisons to results obtained with Linear Variable Differential Transformer (LVDT) position transducers.

Description: The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolutions, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Description: A series of active flow control experiments were recently conducted on a simplified high-lift system. The purpose of the experiments was to explore the prospects of eliminating all but simply hinged leading and trailing edge flaps, while controlling separation on the supercritical airfoil using multiple periodic excitation slots. Excitation was provided by three. independently controlled, self-contained, piezoelectric actuators. Low frequency excitation was generated through amplitude modulation of the high frequency carrier wave, the actuators' resonant frequencies. It was demonstrated, for the first time, that pulsed modulated signal from two neighboring slots interact favorably to increase lift. Phase sensitivity at the low frequency was measured, even though the excitation was synthesized from the high-frequency carrier wave. The measurements were performed at low Reynolds numbers and included mean and unsteady surface pressures, surface hot-films, wake pressures and particle image velocimetry. A modest (6%) increase in maximum lift (compared to the optimal baseline) was obtained due t o the activation of two of the three actuators.

Description: The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolution, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Description: In recent decades, NASA's interest in spacecraft rendezvous and proximity operations has grown. Additional instrumentation is needed to improve manned docking operations' safety, as well as to enable telerobotic operation of spacecraft or completely autonomous rendezvous and docking. To address this need, Advanced Optical Systems, Inc., Orbital Sciences Corporation, and Marshall Space Flight Center have developed the Advanced Video Guidance Sensor (AVGS) under the auspices of the Demonstration of Autonomous Rendezvous Technology (DART) program. Given a cooperative target comprising several retro-reflectors, AVGS provides six-degree-of-freedom information at ranges of up to 300 meters for the DART target. It does so by imaging the target, then performing pattern recognition on the resulting image. Longer range operation is possible through different target geometries. Now that AVGS is being readied for its test flight in 2004, the question is: what next? Modifications can be made to AVGS, including different pattern recognition algorithms and changes to the retro-reflector targets, to make it more robust and accurate. AVGS could be coupled with other space-qualified sensors, such as a laser range-and-bearing finder, that would operate at longer ranges. Different target configurations, including the use of active targets, could result in significant miniaturization over the current AVGS package. We will discuss these and other possibilities for a next-generation docking sensor or sensor suite that involve AVGS.

Description: Differential Evolution (DE) is a simple and robust evolutionary strategy that has been provEn effective in determining the global optimum for several difficult optimization problems. Although DE offers several advantages over traditional optimization approaches, its use in applications such as aerodynamic shape optimization where the objective function evaluations are computationally expensive is limited by the large number of function evaluations often required. In this paper various approaches for improving the efficiency of DE are reviewed and discussed. Several approaches that have proven effective for other evolutionary algorithms are modified and implemented in a DE-based aerodynamic shape optimization method that uses a Navier-Stokes solver for the objective function evaluations. Parallelization techniques on distributed computers are used to reduce turnaround times. Results are presented for standard test optimization problems and for the inverse design of a turbine airfoil. The efficiency improvements achieved by the different approaches are evaluated and compared.

Description: The Videogrammetric Model Deformation (VMD) technique, developed at NASA Langley Research Center, was recently used to measure displacements and local surface angle changes on a static aeroelastic wind-tunnel model. The results were assessed for consistency, accuracy and usefulness. Vertical displacement measurements and surface angular deflections (derived from vertical displacements) taken at no-wind/no-load conditions were analyzed. For accuracy assessment, angular measurements were compared to those from a highly accurate accelerometer. Shewhart's Variables Control Charts were used in the assessment of consistency and uncertainty. Some bad data points were discovered, and it is shown that the measurement results at certain targets were more consistent than at other targets. Physical explanations for this lack of consistency have not been determined. However, overall the measurements were sufficiently accurate to be very useful in monitoring wind-tunnel model aeroelastic deformation and determining flexible stability and control derivatives. After a structural model component failed during a highly loaded condition, analysis of VMD data clearly indicated progressive structural deterioration as the wind-tunnel condition where failure occurred was approached. As a result, subsequent testing successfully incorporated near- real-time monitoring of VMD data in order to ensure structural integrity. The potential for higher levels of consistency and accuracy through the use of statistical quality control practices are discussed and recommended for future applications.

Description: NASA s Synthetic Vision Systems (SVS) project is developing technologies with practical applications that will eliminate low visibility conditions as a causal factor to civil aircraft accidents while replicating the operational benefits of clear day flight operations, regardless of the actual outside visibility condition. A major thrust of the SVS project involves the development/demonstration of affordable, certifiable display configurations that provide intuitive out-the-window terrain and obstacle information with advanced pathway guidance for transport aircraft. This experiment evaluated the influence of different tunnel and guidance concepts upon pilot situation awareness (SA), mental workload, and flight path tracking performance for Synthetic Vision display concepts using a Head-Up Display (HUD). Two tunnel formats (dynamic, minimal) were evaluated against a baseline condition (no tunnel) during simulated IMC approaches to Reno-Tahoe International airport. Two guidance cues (tadpole, follow-me aircraft) were also evaluated to assess their influence on the tunnel formats. Results indicated that the presence of a tunnel on an SVS HUD had no effect on flight path performance but that it did have significant effects on pilot SA and mental workload. The dynamic tunnel concept with the follow-me aircraft guidance symbol produced the lowest workload and provided the highest SA among the tunnel concepts evaluated.

Description: A full-engine simulation of the three-dimensional flow in the GE90 94B high-bypass ratio turbofan engine has been achieved. It would take less than 11 hr of wall clock time if starting from scratch through the exploitation of parallel processing. The simulation of the compressor components, the cooled high-pressure turbine, and the low-pressure turbine was performed using the APNASA turbomachinery flow code. The combustor flow and chemistry were simulated using the National Combustor Code (NCC). The engine simulation matches the engine thermodynamic cycle for a sea-level takeoff condition. The simulation is started at the inlet of the fan and progresses downstream. Comparisons with the cycle point are presented. A detailed look at the blockage in the turbomachinery is presented as one measure to assess and view the solution and the multistage interaction effects.

Description: Initial steps in the application o f a low-order panel method computational fluid dynamic (CFD) code to the calculation of aircraft dynamic stability and control (S&C) derivatives are documented. Several capabilities, unique to CFD but not unique to this particular demonstration, are identified and demonstrated in this paper. These unique capabilities complement conventional S&C techniques and they include the ability to: 1) perform maneuvers without the flow-kinematic restrictions and support interference commonly associated with experimental S&C facilities, 2) easily simulate advanced S&C testing techniques, 3) compute exact S&C derivatives with uncertainty propagation bounds, and 4) alter the flow physics associated with a particular testing technique from those observed in a wind or water tunnel test in order to isolate effects. Also presented are discussions about some computational issues associated with the simulation of S&C tests and selected results from numerous surface grid resolution studies performed during the course of the study.

Description: The NASA F-15 Intelligent Flight Control System project team developed a series of flight control concepts designed to demonstrate neural network-based adaptive controller benefits, with the objective to develop and flight-test control systems using neural network technology to optimize aircraft performance under nominal conditions and stabilize the aircraft under failure conditions. This report presents flight-test results for an adaptive controller using stability and control derivative values from an online learning neural network. A dynamic cell structure neural network is used in conjunction with a real-time parameter identification algorithm to estimate aerodynamic stability and control derivative increments to baseline aerodynamic derivatives in flight. This open-loop flight test set was performed in preparation for a future phase in which the learning neural network and parameter identification algorithm output would provide the flight controller with aerodynamic stability and control derivative updates in near real time. Two flight maneuvers are analyzed - pitch frequency sweep and automated flight-test maneuver designed to optimally excite the parameter identification algorithm in all axes. Frequency responses generated from flight data are compared to those obtained from nonlinear simulation runs. Flight data examination shows that addition of flight-identified aerodynamic derivative increments into the simulation improved aircraft pitch handling qualities.

Description: Aerodynamic design is an iterative process involving geometry manipulation and complex computational analysis subject to physical constraints and aerodynamic objectives. A design cycle consists of first establishing the performance of a baseline design, which is usually created with low-fidelity engineering tools, and then progressively optimizing the design to maximize its performance. Optimization techniques have evolved from relying exclusively on designer intuition and insight in traditional trial and error methods, to sophisticated local and global search methods. Recent attempts at automating the search through a large design space with formal optimization methods include both database driven and direct evaluation schemes. Databases are being used in conjunction with surrogate and neural network models as a basis on which to run optimization algorithms. Optimization algorithms are also being driven by the direct evaluation of objectives and constraints using high-fidelity simulations. Surrogate methods use data points obtained from simulations, and possibly gradients evaluated at the data points, to create mathematical approximations of a database. Neural network models work in a similar fashion, using a number of high-fidelity database calculations as training iterations to create a database model. Optimal designs are obtained by coupling an optimization algorithm to the database model. Evaluation of the current best design then gives either a new local optima and/or increases the fidelity of the approximation model for the next iteration. Surrogate methods have also been developed that iterate on the selection of data points to decrease the uncertainty of the approximation model prior to searching for an optimal design. The database approximation models for each of these cases, however, become computationally expensive with increase in dimensionality. Thus the method of using optimization algorithms to search a database model becomes problematic as the number of design variables is increased.

Description: A free-to-roll study of the low-speed lateral characteristics of the pre-production F/A-18E was conducted in the NASA Langley 12-Foot Low-Speed Tunnel. In developmental flight tests the F/A-18E unexpectedly experienced uncommanded lateral motions in the power approach configuration. The objective of this study was to determine the feasibility of using the free-to-roll technique for the detection of uncommanded lateral motions for the preproduction F/A-18E in the power approach configuration. The data revealed that this technique in conjunction with static data revealed insight into the cause of the lateral motions. The free-to-roll technique identified uncommanded lateral motions at the same angle-of-attack range as experienced in flight tests. The cause of the uncommanded lateral motions was unsteady asymmetric wing stall. The paper also shows that free-to-roll data or static force and moment data alone are not enough to accurately capture the potential for an aircraft to experience uncommanded lateral motion.

Description: The achievement of flight with manned, powered, heavier-than-air aircraft in 1903 marked the beginning of a new era in the means of transportation. A special advantage for aircraft was in speed. However, when an aircraft penetrates the air at very high speeds, the disturbed air is compressed and there are changes in the density, pressure and temperature of the air. These compressibility effects change the aerodynamic characteristics of an aircraft and introduce problems in drag, stability and control. Many aircraft designed in the post-World War II era were plagued with the effects of compressibility. Accordingly, the study of the aerodynamic behavior of aircraft, spacecraft and missiles at high-speed became a major part of the research activity of the NACA/NASA. The intent of the research was to determine the causes and provide some solutions for the aerodynamic problems resulting from the effects of compressibility. The purpose of this paper is to review some of the high-speed aerodynamic research work conducted at the Langley Research Center from the viewpoint of the author who has been active in much of the effort.

Description: The procedure for investigating the effect of deviation from flush mounting of pressure transducers on the exterior of Tu-144LL in flight is described. Experimental data in the mach-number range 0.58 - 2.0 are presented for distortion of the measured wall-pressure fluctuation spectra of the turbulent boundary layer by recessed and protruding transducers. The results of flight experiments are compared with data of wind tunnel experiments. The distortion of measured turbulent boundary layer wall pressure fluctuations caused by transducer-surface deviation from the surrounding surface as a function of dimensionless parameters is predicted and presented on the basis of dimensional analysis.

Description: The stability and control of rotors at high advance ratio are considered. Teetering, articulated, gimbaled, and rigid hub types are considered for a compound helicopter (rotor and fixed wing). Stability predictions obtained using an analytical rigid flapping blade analysis, a rigid blade CAMRAD II model, and an elastic blade CAMRAD II model are compared. For the flapping blade analysis, the teetering rotor is the most stable, 5howing no instabilities up to an advance ratio of 3 and a Lock number of 18. With an elastic blade model, the teetering rotor is unstable at an advance ratio of 1.5. Analysis of the trim controls and blade flapping shows that for small positive collective pitch, trim can be maintained without excessive control input or flapping angles.

Description: The problems of transport delay in a flight simulator, such as its sources and effects, are reviewed. Then their effects on a pilot-in-the-loop control system are investigated with simulations. Three current prominent delay compensators the lead/lag filter, McFarland filter, and the Sobiski/Cardullo filter were analyzed and compared. This paper introduces two novel delay compensation techniques an adaptive predictor using the Kalman estimator and a state space predictive filter using a reference aerodynamic model. Applications of these two new compensators on recorded data from the NASA Langley Research Center Visual Motion Simulator show that they achieve better compensation over the current ones.

Description: This paper describes an approach based on using both bias momentum and multiple control moment gyros for controlling the attitude of statically unstable thrust-levitated vehicles in hover or slow translation. The stabilization approach described in this paper uses these internal angular momentum transfer devices for stability, augmented by thrust vectoring for trim and other outer loop control functions, including CMG stabilization/ desaturation under persistent external disturbances. Simulation results show the feasibility of (1) improved vehicle performance beyond bias momentum assisted vector thrusting control, and (2) using control moment gyros to significantly reduce the external torque required from the vector thrusting machinery.

Description: Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale and BO-105 model (HART-I). The measured data from flight and wind tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free wake models: rolled-up and multiple-trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma, and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple-trailer with consolidation wake model. In the high speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter chord location) of the Puma has an important influence on the phase correlation.

Description: The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.

Description: When aircraft operate in stationary or low speed conditions, airflow into the engine accelerates around the inlet lip and pockets of turbulence that cause noise and vibration can be ingested. This problem has been encountered with engines equipped with the scarf inlet, both in full scale and in model tests, where the noise produced during the static test makes it difficult to assess the noise reduction performance of the scarf inlet. NASA Langley researchers have implemented boundary layer control in an attempt to reduce the influence of the flow nonuniformity in a 12-in. diameter model of a high bypass fan engine mounted in an anechoic chamber. Static pressures and boundary layer profiles were measured in the inlet and far field acoustic measurements were made to assess the effectiveness of the blowing treatment. The blowing system was found to lack the authority to overcome the inlet distortions. Methods to improve the implementation of boundary layer control to reduce inlet distortion are discussed.

Description: This paper deals with the computational analysis of buoyancy effects in the near field of an isothermal helium jet injected into quiescent ambient air environment. The transport equations of helium mass fraction coupled with the conservation equations of mixture mass and momentum were solved using a staggered grid finite volume method. Laminar, axisymmetric, unsteady flow conditions were considered for the analysis. An orthogonal system with non-uniform grids was used to capture the instability phenomena. Computations were performed for Earth gravity and during transition from Earth to different gravitational levels. The flow physics was described by simultaneous visualizations of velocity and concentration fields at Earth and microgravity conditions. Computed results were validated by comparing with experimental data substantiating that buoyancy induced global flow oscillations present in Earth gravity are absent in microgravity. The dependence of oscillation frequency and amplitude on gravitational forcing was presented to further quantify the buoyancy effects.

Description: The AVOSS project demonstrated the feasibility of applying aircraft wake vortex sensing and prediction technologies to safe aircraft spacing for single runway arrivals. On average, AVOSS provided spacing recommendations that were less than the current FAA prescribed spacing rules, resulting in a potential airport efficiency gain. Subsequent efforts have included quantifying the operational specifications for future Wake Vortex Advisory Systems (WakeVAS). In support of these efforts, each of the candidate subsystems for a WakeVAS must be specified. The specifications represent a consensus between the high-level requirements and the capabilities of the candidate technologies. This report documents the beginnings of an effort to quantify the capabilities of the AVOSS Prediction Algorithm (APA). Specifically, the APA horizontal position and circulation strength output sensitivity to the resolution of its wind and turbulence inputs is examined. The results of this analysis have implications for the requirements of the meteorological sensing and prediction systems comprising a WakeVAS implementation.