ALBERT

Description: The transformation of engine control systems from centralized to distributed architecture is both necessary and enabling for future aeropropulsion applications. The continued growth of adaptive control applications and the trend to smaller, light weight cores is a counter influence on the weight and volume of control system hardware. A distributed engine control system using high temperature electronics and open systems communications will reverse the growing trend of control system weight ratio to total engine weight and also be a major factor in decreasing overall cost of ownership for aeropropulsion systems. The implementation of distributed engine control is not without significant challenges. There are the needs for high temperature electronics, development of simple, robust communications, and power supply for the on-board electronics.

Description: Advanced model-based control architecture overcomes the limitations state-of-the-art engine control and provides the potential of virtual sensors, for example for thrust and stall margin. "Tracking filters" are used to adapt the control parameters to actual conditions and to individual engines. For health monitoring standalone monitoring units will be used for on-board analysis to determine the general engine health and detect and isolate sudden faults. Adaptive models open up the possibility of adapting the control logic to maintain desired performance in the presence of engine degradation or to accommodate any faults. Improved and new sensors are required to allow sensing at stations within the engine gas path that are currently not instrumented due in part to the harsh conditions including high operating temperatures and to allow additional monitoring of vibration, mass flows and energy properties, exhaust gas composition, and gas path debris. The environmental and performance requirements for these sensors are summarized.

Description: Active Control can help to meet future engine requirements by an active improvement of the component characteristics. The concept is based on an intelligent control logic, which senses actual operating conditions and reacts with adequate actuator action. This approach can directly improve engine characteristics as performance, operability, durability and emissions on the one hand. On the other hand active control addresses the design constrains imposed by unsteady phenomena like inlet distortion, compressor surge, combustion instability, flow separations, vibration and noise, which only occur during exceptional operating conditions. The feasibility and effectiveness of active control technologies have been demonstrated in lab-scale tests. This chapter describes a broad range of promising applications for each engine component. Significant efforts in research and development remain to implement these technologies in engine rig and finally production engines and to demonstrate today s engine generation airworthiness, safety, reliability, and durability requirements. Active control applications are in particular limited by the gap between available and advanced sensors and actuators, which allow an operation in the harsh environment in an aero engine. The operating and performance requirements for actuators and sensors are outlined for each of the gas turbine sections from inlet to nozzle.

Description: This chapter provides a brief wrap-up of the task group report and focuses on the overall conclusions and recommendations for future work for the CAWAPI and VFE-2 facets beyond the task group. The overall conclusion is that the Technology Readiness Level (TRL) of CFD solvers has been improved in predicting the flow-physics of vortex-dominated flows during the work of the task group, by having flight and wind-tunnel data available for comparison. Moreover, like all good scientific studies, this task group has identified flight conditions on the F-16XL airplane or wind-tunnel test conditions for a specific leading-edge radius on the 65 delta-wing model where the TRL still needs to be increased.

Description: This chapter identifies the benefits that occurred to the AVT-113 task group members and the resulting progress made to two separate vortical flow proposals for task group status being combined into one. Both of these proposals dealt with multiple-vortices, and though they shared different focuses, the general topic, as well as the specific features of this flow, made it of great interest to each sub-task or facet member. The joint meetings increased our overall understanding of vortical flow and the synergistic benefits are summarized in terms of experimental and computational data, virtual laboratory usage, dissemination of results, and career development.

Description: Nine groups participating in the CAWAPI project have contributed steady and unsteady viscous simulations of a full-scale, semi-span model of the F-16XL aircraft. Three different categories of flight Reynolds/Mach number combinations were computed and compared with flight-test measurements for the purpose of code validation and improved understanding of the flight physics. Steady-state simulations are done with several turbulence models of different complexity with no topology information required and which overcome Boussinesq-assumption problems in vortical flows. Detached-eddy simulation (DES) and its successor delayed detached-eddy simulation (DDES) have been used to compute the time accurate flow development. Common structured and unstructured grids as well as individually-adapted unstructured grids were used. Although discrepancies are observed in the comparisons, overall reasonable agreement is demonstrated for surface pressure distribution, local skin friction and boundary velocity profiles at subsonic speeds. The physical modeling, be it steady or unsteady flow, and the grid resolution both contribute to the discrepancies observed in the comparisons with flight data, but at this time it cannot be determined how much each part contributes to the whole. Overall it can be said that the technology readiness of CFD-simulation technology for the study of vehicle performance has matured since 2001 such that it can be used today with a reasonable level of confidence for complex configurations.

Description: In support of the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) with its goal of improving the Technology Readiness Level of flow solvers by comparing results with measured F-16XL-1 flight data, NASA Langley employed the TetrUSS unstructured grid solver, USM3D, to obtain solutions for all seven flight conditions of interest. A newly available solver version that incorporates a number of turbulence models, including the two-equation linear and non-linear k- , was used in this study. As a first test, a choice was made to utilize only a single grid resolution with the solver for the simulation of the different flight conditions. Comparisons are presented with three turbulence models in USM3D, flight data for surface pressure, boundary-layer profiles, and skin-friction distribution, as well as limited predictions from other solvers. A result of these comparisons is that the USM3D solver can be used in an engineering environment to predict vortex-flow physics on a complex configuration at flight Reynolds numbers with a two-equation linear k- turbulence model.

Description: A review is presented of the initial experimental results and analysis that formed the basis the Vortex Flow Experiment 2 (VFE-2). The focus of this work was to distinguish the basic effects of Reynolds number, Mach number, angle of attack, and leading edge bluntness on separation-induced leading-edge vortex flows that are common to slender wings. Primary analysis is focused on detailed static surface pressure distributions, and the results demonstrate significant effects regarding the onset and progression of leading-edge vortex separation.

Description: In this chapter numerical simulations of the flow around F-16XL are performed as a contribution to the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) using the PAB3D CFD code. Two turbulence models are used in the calculations: a standard k-epsilon model, and the Shih-Zhu-Lumley (SZL) algebraic stress model. Seven flight conditions are simulated for the flow around the F-16XL where the free stream Mach number varies from 0.242 to 0.97. The range of angles of attack varies from 0 deg to 20 deg. Computational results, surface static pressure, boundary layer velocity profiles, and skin friction are presented and compared with flight data. Numerical results are generally in good agreement with flight data, considering that only one grid resolution is utilized for the different flight conditions simulated in this study. The Algebraic Stress Model (ASM) results are closer to the flight data than the k-epsilon model results. The ASM predicted a stronger primary vortex, however, the origin of the vortex and footprint is approximately the same as in the k-epsilon predictions.

Description: The objective of the Cranked-Arrow Wing Aerodynamics Project International (CAWAPI) was to allow a comprehensive validation of Computational Fluid Dynamics methods against the CAWAP flight database. A major part of this work involved the generation of high-quality computational grids. Prior to the grid generation an IGES file containing the air-tight geometry of the F-16XL aircraft was generated by a cooperation of some of the CAWAPI partners. Based on this geometry description both structured and unstructured grids have been generated. The baseline structured (multi-block) grid (and a family of derived grids) has been generated by the National Aerospace Laboratory (NLR). The baseline all-tetrahedral and hybrid unstructured grids were generated at the NASA Langley Research Center and the U.S. Air Force Academy, respectively. To provide more geometrical resolution, additional unstructured grids were generated at EADS-MAS, the UTSimCenter, and Boeing Phantom Works. All the grids generated within the framework of CAWAPI will be discussed.

Description: The RTO Task Group AVT-113 "Understanding and Modeling Vortical Flows to Improve the Technology Readiness Level for Military Aircraft" was established in April 2003. Two facets of the group, "Cranked Arrow Wing Aerodynamic Project International (CAWAPI)" and "Vortex Flow Experiment-2 (VFE-2)", worked closely together. However, because of the different requirements of each part, the CAWAPI facet concluded its work earlier (December 2006) than the VFE-2 facet (December 2007). In this first chapter of the Final Report of the Task Group an overview on its work is given, and the objectives for the Task Group are described.

Description: Flight surface flow data of various types for the F-16XL-1 aircraft, employed in the Cranked Arrow Wing Aerodynamics Project (CAWAP), are available.

Description: The Virtual Laboratory (VL) was to be an integral part of the database service that NASA provided to the international community, and for a brief period the VL was fully operational in the CAWAPI facet of the AVT-113 task group. This chapter details how one can construct a VL and also some of the lessons learned along the way that required changes to be made. The VL was to support both the CAWAPI and VFE-2 facets but due to the lack of funding and sufficient Information Technology (IT) support people with the right skills, the VFE-2 facet only reached the advanced planning stage with little software in place. However, both efforts point out the value of a VL in a task group like AVT-113 and illustrate that there needs to be a budgeted item for the IT effort to bring the VL to full operational status in each application.

Description: In the present paper the main results of the new experiments from VFE-2 are summarized. These include some force and moment results, surface and off-body measurements, as well as steady and fluctuating quantities. Some critical remarks are added, and an outlook for future investigations is given.

Description: This paper provides a brief history of the F-16XL-1 aircraft, its role in the High Speed Research (HSR) program and how it was morphed into the Cranked Arrow Wing Aerodynamics Project (CAWAP). Various flight, wind-tunnel and Computational Fluid Dynamics (CFD) data sets were generated during the CAWAP. These unique and open flight datasets for surface pressures, boundary-layer profiles and skin-friction distributions, along with surface flow data, are described and sample data comparisons given. This is followed by a description of how the project became internationalized to be known as Cranked Arrow Wing Aerodynamics Project International (CAWAPI) and is concluded by an introduction to the results of a 5-year CFD predictive study of data.

Description: In this Appendix, sample data are provided in support of Chapter 18. Links and references are also provided.

Description: A computational fluid dynamics (CFD) method has been employed to compute vortical flows around slender wing/body configurations. The emphasis of the paper is on the effectiveness of an adaptive grid procedure in "capturing" concentrated vortices generated at sharp edges or flow separation lines of lifting surfaces flying at high angles of attack. The method is based on a tetrahedral unstructured grid technology developed at the NASA Langley Research Center. Two steady-state, subsonic, inviscid and Navier-Stokes flow test cases are presented to demonstrate the applicability of the method for solving vortical flow problems. The first test case concerns vortex flow over a simple 65 delta wing with different values of leading-edge radius. Although the geometry is quite simple, it poses a challenging problem for computing vortices originating from blunt leading edges. The second case is that of a more complex fighter configuration. The superiority of the adapted solutions in capturing the vortex flow structure over the conventional unadapted results is demonstrated by comparisons with the wind-tunnel experimental data. The study shows that numerical prediction of vortical flows is highly sensitive to the local grid resolution and that the implementation of grid adaptation is essential when applying CFD methods to such complicated flow problems.

Description: An experimental investigation for the flow about a 65 deg. delta wing has been conducted in the NASA Langley National Transonic Facility (NTF). The tests were conducted at Reynolds numbers, based on the mean aerodynamic chord, ranging from 6 million to 120 million and at Mach numbers ranging from 0.4 to 0.9. The model incorporated four different leading-edge bluntness values. The data include detailed static surfacepressure distributions as well as normal-force and pitching-moment coefficients. The test program was designed to quantify the effects of Mach number, Reynolds number, and leading-edge bluntness on the onset and progression of leading-edge vortex separation.

Description: Turbine vane heat transfer predictions are given for smooth and rough vanes where the experimental data show transition moving forward on the vane as the surface roughness physical height increases. Consistent with smooth vane heat transfer, the transition moves forward for a fixed roughness height as the Reynolds number increases. Comparisons are presented with published experimental data. Some of the data are for a regular roughness geometry with a range of roughness heights, Reynolds numbers, and inlet turbulence intensities. The approach taken in this analysis is to treat the roughness in a statistical sense, consistent with what would be obtained from blades measured after exposure to actual engine environments. An approach is given to determine the equivalent sand grain roughness from the statistics of the regular geometry. This approach is guided by the experimental data. A roughness transition criterion is developed, and comparisons are made with experimental data over the entire range of experimental test conditions. Additional comparisons are made with experimental heat transfer data, where the roughness geometries are both regular and statistical. Using the developed analysis, heat transfer calculations are presented for the second stage vane of a high pressure turbine at hypothetical engine conditions.

Description: Computational fluid dynamics (CFD) was used to evaluate the flow field and thrust performance of a promising concept for reducing the noise at take-off of dual-stream turbofan nozzles. The concept, offset stream technology, reduces the jet noise observed on the ground by diverting (offsetting) a portion of the fan flow below the core flow, thickening and lengthening this layer between the high-velocity core flow and the ground observers. In this study a wedge placed in the internal fan stream is used as the diverter. Wind, a Reynolds averaged Navier-Stokes (RANS) code, was used to analyze the flow field of the exhaust plume and to calculate nozzle performance. Results showed that the wedge diverts all of the fan flow to the lower side of the nozzle, and the turbulent kinetic energy on the observer side of the nozzle is reduced. This reduction in turbulent kinetic energy should correspond to a reduction in noise. However, because all of the fan flow is diverted, the upper portion of the core flow is exposed to the freestream, and the turbulent kinetic energy on the upper side of the nozzle is increased, creating an unintended noise source. The blockage due to the wedge reduces the fan mass flow proportional to its blockage, and the overall thrust is consequently reduced. The CFD predictions are in very good agreement with experimental flow field data, demonstrating that RANS CFD can accurately predict the velocity and turbulent kinetic energy fields. While this initial design of a large scale wedge nozzle did not meet noise reduction or thrust goals, this study identified areas for improvement and demonstrated that RANS CFD can be used to improve the concept.

Description: Damaged aircraft have occasionally had to rely solely on thrust to maneuver as a consequence of losing hydraulic power needed to operate flight control surfaces. The lack of successful landings in these cases inspired research into more effective methods of utilizing propulsion-only control. That research demonstrated that one of the major contributors to the difficulty in landing is the slow response of the engines as compared to using traditional flight control. To address this, research is being conducted into ways of making the engine more responsive under emergency conditions. This can be achieved by relaxing controller limits, adjusting schedules, and/or redesigning the regulators to increase bandwidth. Any of these methods can enable faster response at the potential expense of engine life and increased likelihood of stall. However, an example sensitivity analysis revealed a complex interaction of the limits and the difficulty in predicting the way to achieve the fastest response. The sensitivity analysis was performed on a realistic engine model, and demonstrated that significantly faster engine response can be achieved compared to standard Bill of Material control. However, the example indicates the need for an intelligent approach to controller limit adjustment in order for the potential to be fulfilled.

Description: This slide presentation reviews the work of the Experimental Capabilities Supersonic project, that is being reorganized into Flight Research and Validation. The work of Experimental Capabilities Project in FY '09 is reviewed, and the specific centers that is assigned to do the work is given. The portfolio of the newly formed Flight Research and Validation (FRV) group is also reviewed. The various projects for FY '10 for the FRV are detailed. These projects include: Eagle Probe, Channeled Centerbody Inlet Experiment (CCIE), Supersonic Boundary layer Transition test (SBLT), Aero-elastic Test Wing-2 (ATW-2), G-V External Vision Systems (G5 XVS), Air-to-Air Schlieren (A2A), In Flight Background Oriented Schlieren (BOS), Dynamic Inertia Measurement Technique (DIM), and Advanced In-Flight IR Thermography (AIR-T).

Description: Recently proposed flight research at NASA Dryden Flight Research Center (DFRC) has prompted study into the aerodynamic effects of modifications made to the surfaces of laminar airfoils. The research is focused on the high-aspect ratio, laminar-flow type wings commonly found on UAVs and other aircraft with a high endurance requirement. A broad range of instrumentation possibilities, such as structural, pressure, and temperature sensing devices may require the alteration of the airfoil outer mold line as part of the installation process. This study attempts to characterize the effect of installing this additiona1 instrumentation on key airfoil performance factors, such as transition location, lift and drag curves, and stall point. In particular, the general case of an airfoil that is channeled in the spanwise direction is considered, and the impact on key performance characteristics is assessed. Particular attention is focused on exploring the limits of channel depth and low-Reynolds number on performance and stall characteristics. To quantify the effect of increased skin friction due to premature transition caused by protruding or recessed instrumentation, two simplified, conservative scenarios are used to consider two potential sources of diaturbance: A) that leading edge alterations would cause linearly expanding areas (triangles) of turbulent flow on both surfaces of the wing upstream of the natural transition point, and B) that a channel or bump on the upper surface would trip turbulent flow across the whole upper surface upstream of the natural transition point. A potentially more important consideration than the skin friction drag increment is the change in overall airfoil performance due to the installation of instrumentation along most of the wingspan. To quantify this effect, 2D CFD simulations of the flow over a representative mid-span airfoil section were conducted in order to assess the change in lift and drag curves for the airfoil in the presence of disturbed flow due to the installed instrumentation. A discussion as to the impact on high-altitude and low-speed operation of this and similar aircraft is provided.

Description: The research in Supersonic Cruise Efficiency Propulsion (SCE-P) Technical Challenge area of NASA's Supersonics project is discussed. The research in SCE-P is being performed to enable efficient supersonic flight over land. Research elements in this area include: Advance Inlet Concepts, High Performance/Wider Operability Fan and Compressor, Advanced Nozzle Concepts, and Intelligent Sensors/Actuators. The research under each of these elements is briefly discussed.

Description: Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

Description: Unsteady 3-D RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as to experiment. A low Reynolds number k-epsilon turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the tangential direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this work is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.

Description: Multiple sonic boom wind tunnel models were tested in the NASA Ames Research Center 9-by 7-Foot Supersonic Wind Tunnel to reestablish related test techniques in this facility. The goal of the testing was to acquire higher fidelity sonic boom signatures with instrumentation that is significantly more sensitive than that used during previous wind tunnel entries and to compare old and new data from established models. Another objective was to perform tunnel-to-tunnel comparisons of data from a Gulfstream sonic boom model tested at the NASA Langley Research Center 4-foot by 4-foot Unitary Plan Wind Tunnel.

Description: In order to develop the capability to evaluate control system technologies, NASA Ames Research Center (Ames) began a test program to build a Hover Test Vehicle (HTV) - a ground-based simulated flight vehicle. The HTV would integrate simulated propulsion, avionics, and sensors into a simulated flight structure, and fly that test vehicle in terrestrial conditions intended to simulate a flight environment, in particular for attitude control. The ultimate purpose of the effort at Ames is to determine whether the low-cost hardware and flight software techniques are viable for future low cost missions. To enable these engineering goals, the project sought to develop a team, processes and procedures capable of developing, building and operating a fully functioning vehicle including propulsion, GN&C, structure, power and diagnostic sub-systems, through the development of the simulated vehicle.

Description: State of the art boundary layer stability modeling capabilities are increasingly seeing application to entry flight vehicles. With the advent of user friendly and robust implementations of two-dimensional chemical nonequilibrium stability modeling with the STABL/PSE-CHEM software, the need for flight data to calibrate such analyses capabilities becomes more critical. Recent efforts to perform entry flight testing with the Orbiter geometry related to entry aerothermodynamics and boundary layer transition is allowing for a heightened focus on the Orbiter configuration. A significant advancement in the state of the art can likely be achieved by establishing a basis of understanding for the occurrence of boundary layer transition on the Orbiter due to discrete protruding gap fillers and the nominal distributed roughness of the actual thermal protection system. Recent success in demonstrating centerline two-dimensional stability modeling on the centerline of the Orbiter at flight entry conditions provides a starting point for additional investigations. The more detailed paper will include smooth Orbiter configuration boundary layer stability results for several typical orbiter entry conditions. In addition, the numerical modeling approach for establishing the mean laminar flow will be reviewed and the method for determining boundary layer disturbance growth will be overviewed. In addition, if actual Orbiter TPS surface data obtained via digital surface scans become available, it may be possible to investigate the effects of an as-flown flight configuration on boundary layer transition compared to a smooth CAD reference.

Description: Hypersonic chemical nonequilibrium simulations of low earth orbit entry flow fields are becoming increasingly commonplace as software and computational capabilities become more capable. However, development of robust and accurate software to model these environments will always encounter a significant barrier in developing a suite of high quality calibration cases. The US3D hypersonic nonequilibrium Navier Stokes analysis capability has been favorably compared to a number of wind tunnel test cases. Extension of the calibration basis for this software to Orbiter flight conditions will provide an incremental increase in confidence. As part of the Orbiter Boundary Layer Transition Flight Experiment and the Hypersonic Thermodynamic Infrared Measurements project, NASA is performing entry flight testing on the Orbiter to provide valuable aerothermodynamic heating data. An increase in interest related to orbiter entry environments is resulting from this activity. With the advent of this new data, comparisons of the US3D software to the new flight testing data is warranted. This paper will provide information regarding the framework of analyses that will be applied with the US3D analysis tool. In addition, comparisons will be made to entry flight testing data provided by the Orbiter BLT Flight Experiment and HYTHIRM projects. If data from digital scans of the Orbiter windward surface become available, simulations will also be performed to characterize the difference in surface heating between the CAD reference OML and the digitized surface provided by the surface scans.

Description: A series of arc-jet tests in support of the Shuttle Orbiter Boundary Layer Transition flight experiment was conducted in the Channel Nozzle of the NASA Johnson Space Center Atmospheric Reentry Materials and Structures Facility. The boundary layer trip was a protrusion of a certain height and geometry fabricated as part of a 6"x6" tile insert, a special test article made of the Boeing Rigid Insulation tile material and coated with the Reaction Cured Glass used for the bottom fuselage tiles of the Space Shuttle Orbiter. A total of five such tile inserts were manufactured: four with the 0.25-in. trip height, and one with the 0.35-in. trip height. The tile inserts were interchangeably installed in the center of the 24"x24" variable configuration tile array mounted in the 24"x24" test section of the channel nozzle. The objectives of the test series were to demonstrate that the boundary layer trip can safely withstand the Space Shuttle Orbiter flight-like re-entry environments and provide temperature data on the protrusion surface, surfaces of the nearby tiles upstream and downstream of the trip, as well as the bond line between the tiles and the structure. The targeted test environments were defined for the tip of the protrusion, away from the nominal surface of the tile array. The arc jet test conditions were approximated in order to produce the levels of the free stream total enthalpy at the protrusion height similar to those expected in flight. The test articles were instrumented with surface, sidewall and bond line thermocouples. Additionally, Tempilaq temperature-indicating paint was applied to the nominal tiles of the tile array in locations not interfering with the protrusion trip. Five different grades of paint were used that disintegrate at different temperatures between 1500 and 2000 deg F. The intent of using the paint was to gauge the RCG-coated tile surface temperature, as well as determine its usefulness for a flight experiment. This paper provides an overview of the channel nozzle arc jet, test articles and test conditions, as well as the results of the arc-jet tests including the measured temperature response of the test articles, their pre- and post-test surface scans, condition of the thermal paint, and continents on the protrusion tip heating achieved in tests compared to the computational fluid dynamics predictions.

Description: The Rake Airflow Gage Experiment was flown on the Propulsion Flight Test Fixture at NASA Dryden Flight Research Center using one of Dryden s F-15B research testbed aircraft. Propulsion Flight Test Fixture is a modular, pylon-based platform for flight testing propulsion system components, such as the Channeled Centerbody Inlet Experiment, an innovative, variable-geometry, mixed compression supersonic inlet under development at NASA Dryden. The objective of this flight test was to ascertain the flowfield angularity and local Mach number profile of the aerodynamic interface plane that is defined by the planned location of the tip of the inlet centerbody. Knowledge of the flowfield characteristics at this location underneath will be essential to computational modeling of the new inlet as well as future propulsion systems flight testing using the test fixture. This paper describes the preparation for and execution of the flight test, as well as results and validation of the algorithm used to calculate local Mach number and angularity from the rake's pressure measurements.

Description: The Aerostructures Test Wing (ATW) was developed to test unique concepts for flutter prediction and control synthesis. A follow-on to the successful ATW, denoted ATW2, was fabricated as a test bed to validate a variety of instrumentation in flight and to collect data for development of advanced signal processing algorithms for flutter prediction and aviation safety. As a means to estimate flutter speed, a ground vibration test (GVT) was performed. The results of a GVT are typically utilized to update structural dynamics finite element (FE) models used for flutter analysis. In this study, two GVT methodologies were explored to determine which nodes provide the best sensor locations: (i) effective independence and (ii) kinetic energy sorting algorithms. For measurement, ten and twenty sensors were used for three and 10 target test modes. A total of six accelerometer configurations measured frequencies and mode shapes. This included locations used in the original ATW GVT. Moreover, an optical measurement system was used to acquire data without mass effects added by conventional sensors. A considerable frequency shift was observed in comparing the data from the accelerometers to the optical data. The optical data provided robust data for use of the ATW2 finite element model update.

Description: This DVD has several short videos showing some of the work that Dryden is involved in with experimental aircraft. These are: shots showing the Active AeroElastic Wing (AAW) loads calibration tests, AAW roll maneuvers, AAW flight control surface inputs, Helios flight, and takeoff, and Pathfinder takeoff, flight and landing.

Description: A systematic approach for selecting an optimal suite of sensors for on-board aircraft gas turbine engine health estimation is presented. The methodology optimally chooses the engine sensor suite and the model tuning parameter vector to minimize the Kalman filter mean squared estimation error in the engine s health parameters or other unmeasured engine outputs. This technique specifically addresses the underdetermined estimation problem where there are more unknown system health parameters representing degradation than available sensor measurements. This paper presents the theoretical estimation error equations, and describes the optimization approach that is applied to select the sensors and model tuning parameters to minimize these errors. Two different model tuning parameter vector selection approaches are evaluated: the conventional approach of selecting a subset of health parameters to serve as the tuning parameters, and an alternative approach that selects tuning parameters as a linear combination of all health parameters. Results from the application of the technique to an aircraft engine simulation are presented, and compared to those from an alternative sensor selection strategy.

Description: One of the operational modes of the Terminal Area Simulation System (TASS) model simulates the three-dimensional interaction of wake vortices within turbulent domains in the presence of thermal stratification. The model allows the investigation of turbulence and stratification on vortex transport and decay. The model simulations for this work all assumed fully-periodic boundary conditions to remove the effects from any surface interaction. During the Base Period of this contract, NWRA completed generation of these datasets but only presented analysis for the neutral stratification runs of that set (Task 3.4.1). Phase 1 work began with the analysis of the remaining stratification datasets, and in the analysis we discovered discrepancies with the vortex time to link predictions. This finding necessitated investigating the source of the anomaly, and we found a problem with the background turbulence. Using the most up to date version TASS with some important defect fixes, we regenerated a larger turbulence domain, and verified the vortex time to link with a few cases before proceeding to regenerate the entire 25 case set (Task 3.4.2). The effort of Phase 2 (Task 3.4.3) concentrated on analysis of several scenarios investigating the effects of closely spaced aircraft. The objective was to quantify the minimum aircraft separations necessary to avoid vortex interactions between neighboring aircraft. The results consist of spreadsheets of wake data and presentation figures prepared for NASA technical exchanges. For these formation cases, NASA carried out the actual TASS simulations and NWRA performed the analysis of the results by making animations, line plots, and other presentation figures. This report contains the description of the work performed during this final phase of the contract, the analysis procedures adopted, and sample plots of the results from the analysis performed.

Description: A corona discharge device generates an ionic wind and thrust, when a high voltage corona discharge is struck between sharply pointed electrodes and larger radius ground electrodes. The objective of this study was to examine whether this thrust could be scaled to values of interest for aircraft propulsion. An initial experiment showed that the thrust observed did equal the thrust of the ionic wind. Different types of high voltage electrodes were tried, including wires, knife-edges, and arrays of pins. A pin array was found to be optimum. Parametric experiments, and theory, showed that the thrust per unit power could be raised from early values of 5 N/kW to values approaching 50 N/kW, but only by lowering the thrust produced, and raising the voltage applied. In addition to using DC voltage, pulsed excitation, with and without a DC bias, was examined. The results were inconclusive as to whether this was advantageous. It was concluded that the use of a corona discharge for aircraft propulsion did not seem very practical.

Description: A particular failing of Reynolds-averaged Navier-Stokes separated turbulent flow computations is addressed within the context of a kappa-omega two-equation turbulence model. The failing is the tendency for turbulence models to under-predict turbulent shear stress in the shear layers of some separation bubbles, yielding late boundary layer reattachment and recovery. Inspired by unpublished work of Volker, Langtry, and Menter, the author undertook an independent investigation in an attempt to improve the ability of the Menter shear stress transport (SST) model to predict flowfield characteristics in and downstream of separation bubbles. The fix is an ad hoc term that is a function of the local ratio of turbulent production to dissipation; it is used to multiply the omega-destruction term, increasing eddy viscosity in separated regions. With this fix, several flowfields are investigated. Results show that, although the "separation fix" can provide dramatic improvement in some cases, it is not consistently good for all flows. Thus, although it may prove helpful in many situations in its current form, this model may benefit from further refinements, including better sensitization to the energetics of turbulence in the separated region.

Description: The Ko displacement theory, formulated for weak nonuniform (slowly changing cross sections) cantilever beams, was applied to the deformed shape analysis of the doubly-tapered wings of the Ikhana unmanned aircraft. The two-line strain-sensing system (along the wingspan) was used for sensing the bending strains needed for the wing-deformed shapes (deflections and cross-sectional twist) analysis. The deflection equation for each strain-sensing line was expressed in terms of the bending strains evaluated at multiple numbers of strain-sensing stations equally spaced along the strain-sensing line. For the preflight shape analysis of the Ikhana wing, the strain data needed for input to the displacement equations for the shape analysis were obtained from the nodal-stress output of the finite-element analysis. The wing deflections and cross-sectional twist angles calculated from the displacement equations were then compared with those computed from the finite-element computer program. The Ko displacement theory formulated for weak nonlinear cantilever beams was found to be highly accurate in the deformed shape predictions of the doubly-tapered Ikhana wing.

Description: The turbo-shaft engine is an important propulsion system used to power vehicles on land, sea, and in the air. As the power plant for many high performance helicopters, the characteristics of the engine and control are critical to proper vehicle operation as well as being the main determinant to overall vehicle performance. When applied to vertical flight, important distinctions exist in the turbo-shaft engine control system due to the high degree of dynamic coupling between the engine and airframe and the affect on vehicle handling characteristics. In this study, the impact of engine control system architecture is explored relative to engine performance, weight, reliability, safety, and overall cost. Comparison of the impact of architecture on these metrics is investigated as the control system is modified from a legacy centralized structure to a more distributed configuration. A composite strawman system which is typical of turbo-shaft engines in the 1000 to 2000 hp class is described and used for comparison. The overall benefits of these changes to control system architecture are assessed. The availability of supporting technologies to achieve this evolution is also discussed.

Description: The effects of hot corrosion pits on low cycle fatigue life and failure modes of the disk superalloy ME3 were investigated. Low cycle fatigue specimens were subjected to hot corrosion exposures producing pits, then tested at low and high temperatures. Fatigue lives and failure initiation points were compared to those of specimens without corrosion pits. Several tests were interrupted to estimate the fraction of fatigue life that fatigue cracks initiated at pits. Corrosion pits significantly reduced fatigue life by 60 to 98 percent. Fatigue cracks initiated at a very small fraction of life for high temperature tests, but initiated at higher fractions in tests at low temperature. Critical pit sizes required to promote fatigue cracking were estimated, based on measurements of pits initiating cracks on fracture surfaces.

Description: The static aeroelastic equilibrium equations for slender, straight wings are modified to incorporate the effects of aerodynamically-coupled formation flight. A system of equations is developed by applying trim constraints and is solved for component lift distribution, trim angle-of-attack, and trim aileron deflection. The trim values are then used to calculate the elastic twist distribution of the wing box. This system of equations is applied to a formation of two gliders in trimmed flight. Structural and aerodynamic properties are assumed for the gliders, and solutions are calculated for flexible and rigid wings in solo and formation flight. It is shown for a sample application of two gliders in formation flight, that formation disturbances produce greater twist in the wingtip immersed in the vortex than for either the opposing wingtip or the wings of a similar airplane in solo flight. Changes in the lift distribution, resulting from wing twist, increase the performance benefits of formation flight. A flexible wing in formation flight will require greater aileron deflection to achieve roll trim than a rigid wing.

Description: The principal objective of the Supersonics Project is to develop and validate multidisciplinary physics-based predictive design, analysis and optimization capabilities for supersonic vehicles. For aircraft, the focus will be on eliminating the efficiency, environmental and performance barriers to practical supersonic flight. Previous flight projects found that a shaped sonic boom could propagate all the way to the ground (F-5 SSBD experiment) and validated design tools for forebody shape modifications (F-5 SSBD and Quiet Spike experiments). The current project, Lift and Nozzle Change Effects on Tail Shock (LaNCETS) seeks to obtain flight data to develop and validate design tools for low-boom tail shock modifications. Attempts will be made to alter the shock structure of NASA's NF-15B TN/837 by changing the lift distribution by biasing the canard positions, changing the plume shape by under- and over-expanding the nozzles, and changing the plume shape using thrust vectoring. Additional efforts will measure resulting shocks with a probing aircraft (F-15B TN/836) and use the results to validate and update predictive tools. Preliminary flight results are presented and are available to provide truth data for developing and validating the CFD tools required to design low-boom supersonic aircraft.

Description: A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion system and compared the results to those of the baseline single thruster system.

Description: Preliminary design trades are presented for liquid hydrogen fuel systems for remotely-operated, high-altitude aircraft that accommodate three different propulsion options: internal combustion engines, and electric motors powered by either polymer electrolyte membrane fuel cells or solid oxide fuel cells. Mission goal is sustained cruise at 60,000 ft altitude, with duration-aloft a key parameter. The subject aircraft specifies an engine power of 143 to 148 hp, gross liftoff weight of 9270 to 9450 lb, payload of 440 lb, and a hydrogen fuel capacity of 2650 to 2755 lb stored in two spherical tanks (8.5 ft inside diameter), each with a dry mass goal of 316 lb. Hydrogen schematics for all three propulsion options are provided. Each employs vacuum-jacketed tanks with multilayer insulation, augmented with a helium pressurant system, and using electric motor driven hydrogen pumps. The most significant schematic differences involve the heat exchangers and hydrogen reclamation equipment. Heat balances indicate that mission durations of 10 to 16 days appear achievable. The dry mass for the hydrogen system is estimated to be 1900 lb, including 645 lb for each tank. This tank mass is roughly twice that of the advanced tanks assumed in the initial conceptual vehicle. Control strategies are not addressed, nor are procedures for filling and draining the tanks.

Description: NASA Dryden Flight Research Center is a world-class flight research facility located at Edwards AFB, CA. With access to a 44 sq. mile dry lakebed and 350 testable days per year, it is the ideal location for flight research. DFRC has been undertaking aircraft research for approximately six decades including the famous X-aircraft (X-1 through X-48) and many science and exploration platforms. As part of this impressive heritage, DFRC has garnered more hours of full-sized electric aircraft testing than any other facility in the US, and possibly the world. Throughout the 80 s and 90 s Dryden was the home of the Pathfinder, Pathfinder Plus, and Helios prototype solar-electric aircraft. As part of the ERAST program, these electric aircraft achieved a world record 97,000 feet altitude for propeller-driven aircraft. As a result of these programs, Dryden s staff has collected thousands of man-hours of electric aircraft research and testing. In order to better answer the needs of the US in providing aircraft technologies with lower fuel consumption, lower toxic emissions (NOx, CO, VOCs, etc.), lower greenhouse gas (GHG) emissions, and lower noise emissions, NASA has engaged in cross-discipline research under the Aeronautics Research Mission Directorate (ARMD). As a part of this overall effort, Mark Moore of LaRC has initiated a cross-NASA-center electric propulsion working group (EPWG) to focus on electric propulsion technologies as applied to aircraft. Electric propulsion technologies are ideally suited to overcome all of the obstacles mentioned above, and are at a sufficiently advanced state of development component-wise to warrant serious R&D and testing (TRL 3+). The EPWG includes participation from NASA Langley Research Center (LaRC), Glenn Research Center (GRC), Ames Research Center (ARC), and Dryden Flight Research Center (DFRC). Each of the center participants provides their own unique expertise to support the overall goal of advancing the state-of-the-art in aircraft electric propulsion technologies. DFRC will leverage its vast experience in flight test to assist in the integration and flight test phases of any electric propulsion program. DFRC s core competencies, that have particular relevance to the goals of the EPWG, include flight research planning and execution and providing aircraft test beds for researching and testing electric propulsion concepts and equipment. There are three flight regimes that the EPWG is focusing on: subsonic small GA and UAV, subsonic transport class, and supersonic. DFRC proposes two classes of test bed aircraft, to answer the early- and mid-phase testing requirements of all flight regimes the EPWG is concerned with. First, a highly efficient PIK motor glider will be used to test concepts and equipment associated with the subsonic GA and UAV aircraft regime (N+1). Second, a small fleet of subscale remotely-piloted aircraft test beds, similar to the X48B Blended Wing Body aircraft tested at Dryden, will be developed to answer the unique testing requirements of the subsonic GA and UAV, subsonic transport and possibly the supersonic class of aircraft (N+2, N+3). These aircraft can be tested in either serial stages or concurrent stages, depending on the actual test requirements and program schedules. Both classes of test bed aircraft are described below.

Description: The National Aeronautics and Space Administration's (NASA's) major wind tunnel (WT), propulsion test (PT), and simulation facilities exist to serve NASA's and the nation's aeronautics needs. RAND Corporation researchers conducted a prior study of these facilities from 2002 to 2003, identifying (1) NASA's continuing ability to serve national needs, (2) which facilities appear strategically important from an engineering perspective given the vehicle classes the nation investigates and produces, and (3) management challenges and issues. This documented briefing (DB) is the final report from a new, one-year study (conducted from September 2006 through January 2008), partially updating the prior assessment. The study focuses on updating the list of facilities in the prior study that were deemed to be strategically important (again, from an engineering perspective) in serving those needs. This update also adds a new assessment of national needs for six major aeronautics simulators at NASA and lists those deemed strategically important.

Description: Method and system for monitoring and analyzing, in real time, variation with time of an aircraft flight parameter. A time-dependent recovery band, defined by first and second recovery band boundaries that are spaced apart at at least one time point, is constructed for a selected flight parameter and for a selected time recovery time interval length .DELTA.t(FP;rec). A flight parameter, having a value FP(t=t.sub.p) at a time t=t.sub.p, is likely to be able to recover to a reference flight parameter value FP(t';ref), lying in a band of reference flight parameter values FP(t';ref;CB), within a time interval given by t.sub.p.ltoreq.t'.ltoreq.t.sub.p.DELTA.t(FP;rec), if (or only if) the flight parameter value lies between the first and second recovery band boundary traces.

Description: Computational tools are being developed for the design and analysis of supersonic inlets. The objective is to update existing tools and provide design and low-order aerodynamic analysis capability for advanced inlet concepts. The Inlet Tools effort includes aspects of creating an electronic database of inlet design information, a document describing inlet design and analysis methods, a geometry model for describing the shape of inlets, and computer tools that implement the geometry model and methods. The geometry model has a set of basic inlet shapes that include pitot, two-dimensional, axisymmetric, and stream-traced inlet shapes. The inlet model divides the inlet flow field into parts that facilitate the design and analysis methods. The inlet geometry model constructs the inlet surfaces through the generation and transformation of planar entities based on key inlet design factors. Future efforts will focus on developing the inlet geometry model, the inlet design and analysis methods, a Fortran 95 code to implement the model and methods. Other computational platforms, such as Java, will also be explored.

Description: A method of controlling a shear layer for a fluid dynamic body introduces first periodic disturbances into the fluid medium at a first flow separation location. Simultaneously, second periodic disturbances are introduced into the fluid medium at a second flow separation location. A phase difference between the first and second periodic disturbances is adjusted to control flow separation of the shear layer as the fluid medium moves over the fluid dynamic body.

Description: This slide presentation reviews the development and construction of the wireless acoustic instruments surrounding the space shuttle's main engines in preparation for STS-129. The presentation also includes information on end-of-life processing and the mounting procedure for the devices.

Description: No abstract available

Description: Measurement tools for high speed air flow are sought both in industry and academia. Particular interest is shown in air flows that exhibit aerodynamic shocks. Shocks are accompanied by sudden changes in density, pressure, and temperature. Optical detection and characterization of such shocks can be difficult because the medium is normally transparent air. A variety of techniques to analyze these flows are available, but they often require large windows and optical components as in the case of Schlieren measurements and/or large operating powers which precludes their use for in-flight monitoring and applications. The one-dimensional scanning approach in this work is a compact low power technique that can be used to non-intrusively detect shocks. The shock is detected by analyzing the optical pattern generated by a small diameter laser beam as it passes through the shock. The optical properties of a shock result in diffraction and spreading of the beam as well as interference fringes. To investigate the feasibility of this technique a shock is simulated by a 426 m diameter optical fiber. Analysis of results revealed a direct correlation between the optical fiber or shock location and the beam s diffraction pattern. A plot of the width of the diffraction pattern vs. optical fiber location reveals that the width of the diffraction pattern was maximized when the laser beam is directed at the center of the optical fiber. This work indicates that the one-dimensional scanning approach may be able to determine the location of an actual shock. Near and far field effects associated with a small diameter laser beam striking an optical fiber used as a simulated shock are investigated allowing a proper one-dimensional scanning beam technique.

Description: Worldwide concerns of air quality and climate change have made environmental protection one of the most critical issues in aviation today. NASA's current Fundamental Aeronautics research program is directed at three generations of aircraft in the near, mid and far term, with initial operating capability around 2015, 2020, and 2030, respectively. Each generation has associated goals for fuel burn, NOx, noise, and field-length reductions relative to today's aircrafts. The research for the 2020 generation is directed at enabling a hybrid wing body (HWB) aircraft to meet NASA's aggressive technology goals. This paper presents the conceptual cycle and mechanical designs of the two engine concepts, podded and embedded systems, which were proposed for a HWB cargo freighter. They are expected to offer significant benefits in noise reductions without compromising the fuel burn.

Description: Current collaborative research with General Electric Aviation on Open Rotor propulsion as part of the Subsonic Fixed Wing Project Ultra High Bypass Engine Partnership Element is discussed. The Subsonic Fixed Wing Project goals are reviewed, as well as their relative technology level compared to previous NASA noise program goals. The current Open Rotor propulsion research activity at NASA and GE are discussed including the contributions each entity bring toward the research project, and technical plans and objectives.

Description: One goal of the NASA Fundamental Aeronautics Program is the assessment of computational fluid dynamic (CFD) codes used for the design and analysis of many aerospace systems. This paper describes the assessment of the SWIFT turbomachinery analysis code for two similar transonic compressors, NASA rotor 37 and stage 35. The two rotors have identical blade profiles on the front, transonic half of the blade but rotor 37 has more camber aft of the shock. Thus the two rotors have the same shock structure and choking flow but rotor 37 produces a higher pressure ratio. The two compressors and experimental data are described here briefly. Rotor 37 was also used for test cases organized by ASME, IGTI, and AGARD in 1994-1998. Most of the participating codes over predicted pressure and temperature ratios, and failed to predict certain features of the downstream flowfield. Since then the AUSM+ upwind scheme and the k- turbulence model have been added to SWIFT. In this work the new capabilities were assessed for the two compressors. Comparisons were made with overall performance maps and spanwise profiles of several aerodynamic parameters. The results for rotor 37 were in much better agreement with the experimental data than the original blind test case results although there were still some discrepancies. The results for stage 35 were in very good agreement with the data. The results for rotor 37 were very sensitive to turbulence model parameters but the results for stage 35 were not. Comparison of the rotor solutions showed that the main difference between the two rotors was not blade camber as expected, but shock/boundary layer interaction on the casing.

Description: A series of low-speed wind tunnel tests of a Blended-Wing-Body tri-jet configuration to evaluate the low-speed static and dynamic stability and control characteristics over the full envelope of angle of attack and sideslip are summarized. These data were collected for use in simulation studies of the edge-of-the-envelope and potential out-of-control flight characteristics. Some selected results with lessons learned are presented.

Description: A current NASA Research Announcement (NRA) project being conducted by Georgia Tech Research Institute (GTRI) personnel and NASA collaborators includes the development of Circulation Control (CC) blown airfoils to improve subsonic aircraft high-lift and cruise performance. The emphasis of this program is the development of CC active flow control concepts for both high-lift augmentation, drag control, and cruise efficiency. A collaboration in this project includes work by NASA research engineers, whereas CFD validation and flow physics experimental research are part of NASA s systematic approach to developing design and optimization tools for CC applications to fixed-wing aircraft. The design space for CESTOL type aircraft is focusing on geometries that depend on advanced flow control technologies that include Circulation Control aerodynamics. The ability to consistently predict advanced aircraft performance requires improvements in design tools to include these advanced concepts. Validation of these tools will be based on experimental methods applied to complex flows that go beyond conventional aircraft modeling techniques. This paper focuses on recent/ongoing benchmark high-lift experiments and CFD efforts intended to provide 2-D CFD validation data sets related to NASA s Cruise Efficient Short Take Off and Landing (CESTOL) study. Both the experimental data and related CFD predictions are discussed.

Description: This viewgraph presentation describes the high temperature structural measurements developments for hypersonic airframes applications.

Description: Comparison metrics can be established to reliably and repeatedly establish the health of the joggle region of the Orbiter Wing Leading Edge reinforced carbon carbon (RCC) panels. Using these metrics can greatly reduced the man hours needed to perform, wing leading edge scanning for service induced damage. These time savings have allowed for more thorough inspections to be preformed in the necessary areas with out affecting orbiter flow schedule. Using specialized local inspections allows for a larger margin of safety by allowing for more complete characterizations of panel defects. The presence of the t-seal during thermographic inspection can have adverse masking affects on ability properly characterize defects that exist in the joggle region of the RCC panels. This masking affect dictates the final specialized inspection should be preformed with the t-seal removed. Removal of the t-seal and use of the higher magnification optics has lead to the most effective and repeatable inspection method for characterizing and tracking defects in the wing leading edge. Through this study some inadequacies in the main health monitoring system for the orbiter wing leading edge have been identified and corrected. The use of metrics and local specialized inspection have lead to a greatly increased reliability and repeatable inspection of the shuttle wing leading edge.

Description: As world emissions are further scrutinized to identify areas for improvement, aviation s contribution to the problem can no longer be ignored. Previous studies for zero or near-zero emissions aircraft suggest aircraft and propulsion system sizes that would perform propulsion system and subsystems layout and propellant tankage analyses to verify the weight-scaling relationships. These efforts could be used to identify and guide subsequent work on systems and subsystems to achieve viable aircraft system emissions goals. Previous work quickly focused these efforts on propulsion systems for 70- and 100-passenger aircraft. Propulsion systems modeled included hydrogen-fueled gas turbines and fuel cells; some preliminary estimates combined these two systems. Hydrogen gas-turbine engines, with advanced combustor technology, could realize significant reductions in nitrogen emissions. Hydrogen fuel cell propulsion systems were further laid out, and more detailed analysis identified systems needed and weight goals for a viable overall system weight. Results show significant, necessary reductions in overall weight, predominantly on the fuel cell stack, and power management and distribution subsystems to achieve reasonable overall aircraft sizes and weights. Preliminary conceptual analyses for a combination of gas-turbine and fuel cell systems were also performed, and further studies were recommended. Using gas-turbine engines combined with fuel cell systems can reduce the fuel cell propulsion system weight, but at higher fuel usage than using the fuel cell only.

Description: An investigation of the aeroheating environment of the Project Orion Crew Entry Vehicle has been performed in the Langley Research Center 20-Inch Mach 6 Air Tunnel. Data were measured on a approx.3.5% scale model (0.1778-m/7-inch diameter) of the vehicle using coaxial thermocouples at free stream Reynolds numbers of 2.0 10(exp 6)/ft to 7.30 10(exp 6)/ft and computational predictions were generated for all test conditions. The primary goals of this test were to obtain convective heating data for use in assessing the accuracy of the computational technique and to validate test methodology and heating data from a test of the same wind tunnel model in the Arnold Engineering Development Center Tunnel 9. Secondary goals were to determine the extent of transitional/turbulent data which could be produced on a CEV model in this facility, either with or without boundary-layer trips, and to demonstrate continuous pitch-sweep operation in this tunnel for heat transfer testing.

Description: United States Air Force Research Laboratory (AFRL) ground tests at the NASA Transonic Dynamics Tunnel (TDT) and NASA flight tests provide a basis and methodology for in-flight characterization of the aeroelastic performance through the monitoring of the fluid-structure interaction using surface flow sensors. NASA NF-15B flight tests provided a unique opportunity to test the correlation of aerodynamic loads with sectional flow attachment/detachment points, also known as flow bifurcation points (FBPs), as observed in previous wind tunnel tests. The NF-15B tail was instrumented with hot-film sensors and strain gages for measuring root-bending strains. These data were gathered via selected sideslip maneuvers performed at level flight and subsonic speeds. The aerodynamic loads generated by the sideslip maneuver resulted in root-bending strains and hot-film sensor signals near the stagnation region that were highly correlated. For the TDT tests, a flexible wing section developed under the AFRL SensorCraft program was instrumented with strain gages, accelerometers, and hot-film sensors at multiple span stations. The TDT tests provided data showing a gradual phase change between the FBP and the structural mode occurred during a resonant condition as the wings structural modes were excited by the tunnel-generated gusts.

Description: This document describes the development of further extensions and improvements to the jet noise model developed by Modern Technologies Corporation (MTC) for the National Aeronautics and Space Administration (NASA). The noise component extraction and correlation approach, first used successfully by MTC in developing a noise prediction model for two-dimensional mixer ejector (2DME) nozzles under the High Speed Research (HSR) Program, has been applied to dual-stream nozzles, then extended and improved in earlier tasks under this contract. Under Task 6, the coannular jet noise model was formulated and calibrated with limited scale model data, mainly at high bypass ratio, including a limited-range prediction of the effects of mixing-enhancement nozzle-exit chevrons on jet noise. Under Task 9 this model was extended to a wider range of conditions, particularly those appropriate for a Supersonic Business Jet, with an improvement in simulated flight effects modeling and generalization of the suppressor model. In the present task further comparisons are made over a still wider range of conditions from more test facilities. The model is also further generalized to cover single-stream nozzles of otherwise similar configuration. So the evolution of this prediction/analysis/correlation approach has been in a sense backward, from the complex to the simple; but from this approach a very robust capability is emerging. Also from these studies, some observations emerge relative to theoretical considerations. The purpose of this task is to develop an analytical, semi-empirical jet noise prediction method applicable to takeoff, sideline and approach noise of subsonic and supersonic cruise aircraft over a wide size range. The product of this task is an even more consistent and robust model for the Footprint/Radius (FOOTPR) code than even the Task 9 model. The model is validated for a wider range of cases and statistically quantified for the various reference facilities. The possible role of facility effects will thus be documented. Although the comparisons that can be accomplished within the limited resources of this task are not comprehensive, they provide a broad enough sampling to enable NASA to make an informed decision on how much further effort should be expended on such comparisons. The improved finalized model is incorporated into the FOOTPR code. MTC has also supported the adaptation of this code for incorporation in NASA s Aircraft Noise Prediction Program (ANOPP).

Description: The present study, which is the first of a series of investigations of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary layer flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed on a large-scale, high-subsonic unsteady turbine cascade research facility with an integrated wake generator and test section unit. Blade Pak B geometry was used in the cascade. The wakes were generated by continuously moving cylindrical bars device. Boundary layer investigations were performed using hot wire anemometry at Reynolds number of 110,000, based on the blade suction surface length and the exit velocity, for one steady and two unsteady inlet flow conditions, with the corresponding passing frequencies, wake velocities, and turbulence intensities. The reduced frequencies cover the entire operation range of LP-turbines. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re = 50,000, 75,000, 100,000, 110,000, and 125,000. For each Reynolds number, surface pressure measurements are carried out at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extension of the separation zone as well as its behavior under unsteady wake flow. The results, presented in ensemble-averaged and contour plot forms, help to understand the physics of the separation phenomenon under periodic unsteady wake flow.

Description: This paper examines the aeroelastic stability of an on-orbit installable Space Shuttle patch panel. CFD flutter solutions were obtained for thick and thin boundary layers at a free stream Mach number of 2.0 and several Mach numbers near sonic speed. The effect of structural damping on these flutter solutions was also examined, and the effect of structural nonlinearities associated with in-plane forces in the panel was considered on the worst case linear flutter solution. The results of the study indicated that adequate flutter margins exist for the panel at the Mach numbers examined. The addition of structural damping improved flutter margins as did the inclusion of nonlinear effects associated with a static pressure difference across the panel.

Description: The influence of lift offset on the performance of several rotorcraft configurations is explored. A lift-offset rotor, or advancing blade concept, is a hingeless rotor that can attain good efficiency at high speed by operating with more lift on the advancing side than on the retreating side of the rotor disk. The calculated performance capability of modern-technology coaxial rotors utilizing a lift offset is examined, including rotor performance optimized for hover and high-speed cruise. The ideal induced power loss of coaxial rotors in hover and twin rotors in forward flight is presented. The aerodynamic modeling requirements for performance calculations are evaluated, including wake and drag models for the high-speed flight condition. The influence of configuration on the performance of rotorcraft with lift-offset rotors is explored, considering tandem and side-by-side rotorcraft as well as wing-rotor lift share.

Description: Data-Parallel Line Relaxation (DPLR) code is a computational fluid dynamic (CFD) solver that was developed at NASA Ames Research Center to help mission support teams generate high-value predictive solutions for hypersonic flow field problems. The DPLR Code Package is an MPI-based, parallel, full three-dimensional Navier-Stokes CFD solver with generalized models for finite-rate reaction kinetics, thermal and chemical non-equilibrium, accurate high-temperature transport coefficients, and ionized flow physics incorporated into the code. DPLR also includes a large selection of generalized realistic surface boundary conditions and links to enable loose coupling with external thermal protection system (TPS) material response and shock layer radiation codes.

Description: A request was submitted on September 2, 2004 concerning the uncertainties regarding the acoustic environment within the Stratospheric Observatory for Infrared Astronomy (SOFIA) cavity, and the potential for structural damage from acoustical resonance or tones, especially if they occur at or near a structural mode. The requestor asked for an independent expert opinion on the approach taken by the SOFIA project to determine if the project's analysis, structural design and proposed approach to flight test were sound and conservative. The findings from this assessment are recorded in this document.

Description: JeNo (Version 1.0) is a Fortran90 computer code that calculates the far-field sound spectral density produced by axisymmetric, unheated jets at a user specified observer location and frequency range. The user must provide a structured computational grid and a mean flow solution from a Reynolds-Averaged Navier Stokes (RANS) code as input. Turbulence kinetic energy and its dissipation rate from a k-epsilon or k-omega turbulence model must also be provided. JeNo is a research code, and as such, its development is ongoing. The goal is to create a code that is able to accurately compute far-field sound pressure levels for jets at all observer angles and all operating conditions. In order to achieve this goal, current theories must be combined with the best practices in numerical modeling, all of which must be validated by experiment. Since the acoustic predictions from JeNo are based on the mean flow solutions from a RANS code, quality predictions depend on accurate aerodynamic input.This is why acoustic source modeling, turbulence modeling, together with the development of advanced measurement systems are the leading areas of research in jet noise research at NASA Glenn Research Center.

Description: A plan is presented for the development of a high fidelity multidisciplinary optimization process for rotorcraft. The plan formulates individual disciplinary design problems, identifies practical high-fidelity tools and processes that can be incorporated in an automated optimization environment, and establishes statements of the multidisciplinary design problem including objectives, constraints, design variables, and cross-disciplinary dependencies. Five key disciplinary areas are selected in the development plan. These are rotor aerodynamics, rotor structures and dynamics, fuselage aerodynamics, fuselage structures, and propulsion / drive system. Flying qualities and noise are included as ancillary areas. Consistency across engineering disciplines is maintained with a central geometry engine that supports all multidisciplinary analysis. The multidisciplinary optimization process targets the preliminary design cycle where gross elements of the helicopter have been defined. These might include number of rotors and rotor configuration (tandem, coaxial, etc.). It is at this stage that sufficient configuration information is defined to perform high-fidelity analysis. At the same time there is enough design freedom to influence a design. The rotorcraft multidisciplinary optimization tool is built and substantiated throughout its development cycle in a staged approach by incorporating disciplines sequentially.

Description: In 2006, NASA Dryden Flight Research Center, Edwards, Calif., obtained a civil version of the General Atomics MQ-9 unmanned aircraft system and modified it for research purposes. Proposed missions included support of Earth science research, development of advanced aeronautical technology, and improving the utility of unmanned aerial systems in general. The project team named the aircraft Ikhana a Native American Choctaw word meaning intelligent, conscious, or aware in order to best represent NASA research goals. Building on experience with these and other unmanned aircraft, NASA scientists developed plans to use the Ikhana for a series of missions to map wildfires in the western United States and supply the resulting data to firefighters in near-real time. A team at NASA Ames Research Center, Mountain View, Calif., developed a multispectral scanner that was key to the success of what became known as the Western States Fire Missions. Carried out by team members from NASA, the U.S. Department of Agriculture Forest Service, National Interagency Fire Center, National Oceanic and Atmospheric Administration, Federal Aviation Administration, and General Atomics Aeronautical Systems Inc., these flights represented an historic achievement in the field of unmanned aircraft technology.

Description: This paper presents an integrated flight dynamic modeling method for flexible aircraft that captures coupled physics effects due to inertial forces, aeroelasticity, and propulsive forces that are normally present in flight. The present approach formulates the coupled flight dynamics using a structural dynamic modeling method that describes the elasticity of a flexible, twisted, swept wing using an equivalent beam-rod model. The structural dynamic model allows for three types of wing elastic motion: flapwise bending, chordwise bending, and torsion. Inertial force coupling with the wing elasticity is formulated to account for aircraft acceleration. The structural deflections create an effective aeroelastic angle of attack that affects the rigid-body motion of flexible aircraft. The aeroelastic effect contributes to aerodynamic damping forces that can influence aerodynamic stability. For wing-mounted engines, wing flexibility can cause the propulsive forces and moments to couple with the wing elastic motion. The integrated flight dynamics for a flexible aircraft are formulated by including generalized coordinate variables associated with the aeroelastic-propulsive forces and moments in the standard state-space form for six degree-of-freedom flight dynamics. A computational structural model for a generic transport aircraft has been created. The eigenvalue analysis is performed to compute aeroelastic frequencies and aerodynamic damping. The results will be used to construct an integrated flight dynamic model of a flexible generic transport aircraft.

Description: Within this paper, control-relevant vehicle design concepts are examined using a widely used 3 DOF (plus flexibility) nonlinear model for the longitudinal dynamics of a generic carrot-shaped scramjet powered hypersonic vehicle. Trade studies associated with vehicle/engine parameters are examined. The impact of parameters on control-relevant static properties (e.g. level-flight trimmable region, trim controls, AOA, thrust margin) and dynamic properties (e.g. instability and right half plane zero associated with flight path angle) are examined. Specific parameters considered include: inlet height, diffuser area ratio, lower forebody compression ramp inclination angle, engine location, center of gravity, and mass. Vehicle optimizations is also examined. Both static and dynamic considerations are addressed. The gap-metric optimized vehicle is obtained to illustrate how this control-centric concept can be used to "reduce" scheduling requirements for the final control system. A classic inner-outer loop control architecture and methodology is used to shed light on how specific vehicle/engine design parameter selections impact control system design. In short, the work represents an important first step toward revealing fundamental tradeoffs and systematically treating control-relevant vehicle design.

Description: In this paper, we investigate the use of Bayesian networks to construct large-scale diagnostic systems. In particular, we consider the development of large-scale Bayesian networks by composition. This compositional approach reflects how (often redundant) subsystems are architected to form systems such as electrical power systems. We develop high-level specifications, Bayesian networks, clique trees, and arithmetic circuits representing 24 different electrical power systems. The largest among these 24 Bayesian networks contains over 1,000 random variables. Another BN represents the real-world electrical power system ADAPT, which is representative of electrical power systems deployed in aerospace vehicles. In addition to demonstrating the scalability of the compositional approach, we briefly report on experimental results from the diagnostic competition DXC, where the ProADAPT team, using techniques discussed here, obtained the highest scores in both Tier 1 (among 9 international competitors) and Tier 2 (among 6 international competitors) of the industrial track. While we consider diagnosis of power systems specifically, we believe this work is relevant to other system health management problems, in particular in dependable systems such as aircraft and spacecraft. (See CASI ID 20100021910 for supplemental data disk.)

Description: Objectives: a) Regain Stable Platform: 1) Metrics analogous to stability margins needed for adaptive control systems; 2) Avoid adverse structural interactions. b) Maneuverability: 1) Control vehicle within new constraints; 2) Respect structural limitations; 3) Inform pilot of new performance limitations. c) Provide ability to safely land airplane: 1) Develop safest recovery trajectory.

Description: Experimental tests have been completed which recorded the ability of two combination steady state and high response time varying Pitot probe designs to accurately measure steady stagnation pressure at a single location in a flow field. Tests were conducted of double-barreled and coannular Prati probes in a 3.5 in. diameter free jet probe calibration facility from Mach 0.1 to 0.9. Geometric symmetry and pitch (-40 deg to 40 deg) and yaw (0 deg to 40 deg) angle actuation were used to fully evaluate the probes. These tests revealed that the double-barreled configuration induced error in its steady state measurement at zero incidence that increased consistently with jet Mach number to 1.1 percent at Mach 0.9. For all Mach numbers, the double-barreled probe nulled at a pitch angle of approximately 7.0 deg and provided inconsistent measurements when yawed. The double-barreled probe provided adequate measurements via both its steady state and high response tubes (within +/- 0.15 percent accuracy) over unacceptable ranges of biased pitch and inconsistent yaw angles which varied with Mach number. By comparison, the coannular probe provided accurate measurements (at zero incidence) for all jet Mach numbers as well as over a flow angularity range which varied from +/- 26.0 deg at Mach 0.3 deg to +/- 14.0 deg at Mach 0.9. Based on these results, the Prati probe is established as the preferred design. Further experimental tests are recommended to document the frequency response characteristics of the Prati probe.

Description: The effectiveness of microramp flow control devices in controlling an oblique shock interaction was tested in the 15- by 15-Centimeter Supersonic Wind Tunnel at NASA Glenn Research Center. Fifteen microramp geometries were tested varying the height, chord length, and spacing between ramps. Measurements of the boundary layer properties downstream of the shock reflection were analyzed using design of experiments methods. Results from main effects, D-optimal, full factorial, and central composite designs were compared. The designs provided consistent results for a single variable optimization.

Description: High quality jet noise spectral data measured at the Aeroacoustic Propulsion Laboratory at the NASA Glenn Research Center is used to examine a number of jet noise scaling laws. Configurations considered in the present study consist of convergent and convergent-divergent axisymmetric nozzles. Following the work of Viswanathan, velocity power factors are estimated using a least squares fit on spectral power density as a function of jet temperature and observer angle. The regression parameters are scrutinized for their uncertainty within the desired confidence margins. As an immediate application of the velocity power laws, spectral density in supersonic jets are decomposed into their respective components attributed to the jet mixing noise and broadband shock associated noise. Subsequent application of the least squares method on the shock power intensity shows that the latter also scales with some power of the shock parameter. A modified shock parameter is defined in order to reduce the dependency of the regression factors on the nozzle design point within the uncertainty margins of the least squares method.

Description: This slide presentation reviews several projects that NASA Dryden personnel are involved with: Integrated Resilient Aircraft Controls Project (IRAC), NASA G-III Research Aircraft, X-48B Blended Wing Body aircraft, Stratospheric Observatory for Infrared Astronomy (SOFIA), and the Orion CEV Launch Abort Systems Tests.

Description: This video, accompanying material for a talk, shows portions of the F-15 flight test. It shows the in-flight movement of the plane, and the landing.

Description: Presentation for Aging Aircraft conference covering chafing fault diagnostics using Time Domain Reflectometry. Laboratory setup and experimental methods are presented, along with initial results that summarize fault modeling and detection capabilities.

Description: Large complex aerospace systems are generally validated in regions local to anticipated operating points rather than through characterization of the entire feasible operational envelope of the system. This is due to the large parameter space, and complex, highly coupled nonlinear nature of the different systems that contribute to the performance of the aerospace system. We have addressed the factors deterring such an analysis by applying a combination of technologies to the area of flight envelop assessment. We utilize n-factor (2,3) combinatorial parameter variations to limit the number of cases, but still explore important interactions in the parameter space in a systematic fashion. The data generated is automatically analyzed through a combination of unsupervised learning using a Bayesian multivariate clustering technique (AutoBayes) and supervised learning of critical parameter ranges using the machine-learning tool TAR3, a treatment learner. Covariance analysis with scatter plots and likelihood contours are used to visualize correlations between simulation parameters and simulation results, a task that requires tool support, especially for large and complex models. We present results of simulation experiments for a cold-gas-powered hover test vehicle.

Description: Boeing and a team from Air Force, NASA, Army, DARPA, MIT, UCLA, and U. of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero-acoustic analysis, and quantified the effects of open and closed loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed significant reductions up to 6dB in blade-vortex-interaction and in-plane noise, as well as reductions in vibratory hub loads up to 80%. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind-tunnel testing.

Description: From the fall of 1993 to late winter of 1994, NASA Ames and the U.S. Army flew a flight test program using a UH-60A helicopter with extensive instrumentation on the rotor and blades, including 242 pressure transducers. Over this period, approximately 30 flights were made, and data were obtained in level flight, maneuver, ascents, and descents. Coordinated acoustic measurements were obtained with a ground-acoustic array in cooperation with NASA Langley, and in-flight acoustic measurements with a YO-3A aircraft. NASA has sponsored the creation of a "tutorial' which covers the depth and breadth of the flight test program with a mixture of text and graphics. The primary purpose of this tutorial is to introduce the student to what is known about rotor aerodynamics based on the UH-60A measurements. The tutorial will also be useful to anyone interested in helicopters who would like to have more detailed knowledge about helicopter aerodynamics.

Description: This slide presentation reviews the preliminary Flight tests of the X-48B development program. The X-48B is a blended wing body aircraft that is being used to test various features of the BWB concept. The research concerns the following: (1) Turbofan Development, (2) Intelligent Flight Control and Optimization, (3) Airdata Calibration (4) Parameter Identification (i.e., Determination of the parameters of a mathematical model of a system based on observation of the system inputs and response.)

Description: Boeing and a team from Air Force, NASA, Army, Massachusetts Institute of Technology, University of California at Los Angeles, and University of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center, figure 1. The SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing-edge flap on each blade. The development effort included design, fabrication, and component testing of the rotor blades, the trailing-edge flaps, the piezoelectric actuators, the switching power amplifiers, the actuator control system, and the data/power system. Development of the smart rotor culminated in a whirl-tower hover test which demonstrated the functionality, robustness, and required authority of the active flap system. The eleven-week wind tunnel test program evaluated the forward flight characteristics of the active-flap rotor, gathered data to validate state-of-the-art codes for rotor noise analysis, and quantified the effects of open- and closed-loop active-flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness and the reliability of the flap actuation system were successfully demonstrated in more than 60 hours of wind-tunnel testing. The data acquired and lessons learned will be instrumental in maturing this technology and transitioning it into production. The development effort, test hardware, wind-tunnel test program, and test results will be presented in the full paper.

Description: The desire for higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets, to low bypass ratio, first generation turbofans, to today's high bypass ratio turbofans. It is possible that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines. Although increased bypass ratio has clear benefits in terms of propulsion system metrics such as specific fuel consumption, these benefits may not translate into aircraft system level benefits due to integration penalties. In this study, the design trade space for advanced turbofan engines applied to a single-aisle transport (737/A320 class aircraft) is explored. The benefits of increased bypass ratio and associated enabling technologies such as geared fan drive are found to depend on the primary metrics of interest. For example, bypass ratios at which fuel consumption is minimized may not require geared fan technology. However, geared fan drive does enable higher bypass ratio designs which result in lower noise. Regardless of the engine architecture chosen, the results of this study indicate the potential for the advanced aircraft to realize substantial improvements in fuel efficiency, emissions, and noise compared to the current vehicles in this size class.

Description: Preliminary aerodynamic and performance predictions for an active twist rotor for a HART-II type of configuration are performed using a computational fluid dynamics (CFD) code, OVERFLOW2, and a computational structural dynamics (CSD) code, CAMRAD -II. These codes are loosely coupled to compute a consistent set of aerodynamics and elastic blade motions. Resultant aerodynamic and blade motion data are then used in the Ffowcs-Williams Hawkins solver, PSU-WOPWOP, to compute noise on an observer plane under the rotor. Active twist of the rotor blade is achieved in CAMRAD-II by application of a periodic torsional moment couple (of equal and opposite sign) at the blade root and tip at a specified frequency and amplitude. To provide confidence in these particular active twist predictions for which no measured data is available, the rotor system geometry and computational set up examined here are identical to that used in a previous successful Higher Harmonic Control (HHC) computational study. For a single frequency equal to three times the blade passage frequency (3P), active twist is applied across a range of control phase angles at two different amplitudes. Predicted results indicate that there are control phase angles where the maximum mid-frequency noise level and the 4P non -rotating hub vibrations can be reduced, potentially, both at the same time. However, these calculated reductions are predicted to come with a performance penalty in the form of a reduction in rotor lift-to-drag ratio due to an increase in rotor profile power.

Description: Overview: Provide validation of adaptive control law concepts through full scale flight evaluation in a representative avionics architecture. Develop an understanding of aircraft dynamics of current vehicles in damaged and upset conditions Real-world conditions include: a) Turbulence, sensor noise, feedback biases; and b) Coupling between pilot and adaptive system. Simulated damage includes 1) "B" matrix (surface) failures; and 2) "A" matrix failures. Evaluate robustness of control systems to anticipated and unanticipated failures.

Description: A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Description: Surface roughness can influence laminar-turbulent transition in many different ways. This paper outlines selected analyses performed at the NASA Langley Research Center, ranging in speed from subsonic to hypersonic Mach numbers and highlighting the beneficial as well as adverse roles of the surface roughness in technological applications. The first theme pertains to boundary-layer tripping on the forebody of a hypersonic airbreathing configuration via a spanwise periodic array of trip elements, with the goal of understanding the physical mechanisms underlying roughness-induced transition in a high-speed boundary layer. The effect of an isolated, finite amplitude roughness element on a supersonic boundary layer is considered next. The other set of flow configurations examined herein corresponds to roughness based laminar flow control in subsonic and supersonic swept wing boundary layers. A common theme to all of the above configurations is the need to apply higher fidelity, physics based techniques to develop reliable predictions of roughness effects on laminar-turbulent transition.

Description: Nitric oxide (NO) planar laser-induced fluorescence (PLIF) has been use to investigate the hypersonic flow over a flat plate with and without a 2-mm (0.08-in) radius hemispherical trip. In the absence of the trip, for all angles of attack and two different Reynolds numbers, the flow was observed to be laminar and mostly steady. Boundary layer thicknesses based on the observed PLIF intensity were measured and compared with a CFD computation, showing agreement. The PLIF boundary layer thickness remained constant while the NO flowrate was varied by a factor of 3, indicating non-perturbative seeding of NO. With the hemispherical trip in place, the flow was observed to be laminar but unsteady at the shallowest angle of attack and lowest Reynolds number and appeared vigorously turbulent at the steepest angle of attack and highest Reynolds number. Laminar corkscrew-shaped vortices oriented in the streamwise direction were frequently observed to transition the flow to more turbulent structures.

Description: A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Description: The modeling of porous bleed regions as boundary conditions in computational fluid dynamics (CFD) simulations of supersonic inlet flows has been improved through a scaling of sonic flow coefficient data for 90deg bleed holes. The scaling removed the Mach number as a factor in computing the sonic flow coefficient and allowed the data to be fitted with a quadratic equation, with the only factor being the ratio of the plenum static pressure to the surface static pressure. The implementation of the bleed model into the Wind-US CFD flow solver was simplified by no longer requiring the evaluation of the flow properties at the boundary-layer edge. The quadratic equation can be extrapolated to allow the modeling of small amounts of blowing, which can exist when recirculation of the bleed flow occurs within the bleed region. The improved accuracy of the bleed model was demonstrated through CFD simulations of bleed regions on a flat plate in supersonic flow with and without an impinging oblique shock. The bleed model demonstrated good agreement with experimental data and three-dimensional CFD simulations of bleed holes.

Description: The Soft X-Ray Telescope (SXT) is an instrument on the International X-Ray Observatory (IXO). Its flight mirror assembly (FMA) has a single mirror configuration that includes a 3.3 m diameter and 0.93 m tall mirror assembly. It consists of 24 outer modules, 24 middle modules and 12 inner modules. Each module includes more than 200 mirror segments. There are a total of nearly 14,000 mirror segments. The operating temperature requirement of the SXT FMA is 20 C. The spatial temperature gradient requirement between the FMA modules is 1 C or smaller. The spatial temperature gradient requirement within a module is 0.5 C. This paper presents thermal design considerations to meet these stringent thermal requirements.

Description: The stability and receptivity of three-dimensional supersonic boundary layers over a 7deg sharp tipped straight cone at an angle of attack of 4.2deg is numerically investigated at a free stream Mach number of 3.5 and at two high Reynolds numbers, 0.25 and 0.50x10(exp 6)/inch. The generation and evolution of stationary crossflow vortices are also investigated by performing simulations with three-dimensional roughness elements located on the surface of the cone. The flow fields with and without the roughness elements are obtained by solving the full Navier-Stokes equations in cylindrical coordinates using the fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and using the third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. Stability computations reveal that the azimuthal wavenumbers are in the range of m approx. 25-50 for the most amplified traveling disturbances and in the range of m approx. 40-70 for the stationary disturbances. The N-Factor computations predicted that transition would occur further forward in the middle of the cone compared to the transition fronts near the windward and the leeward planes. The simulations revealed that the crossflow vortices originating from the nose region propagate towards the leeward plane. No perturbations were observed in the lower part of the cone.

Description: An experimental investigation of the DLR-F6 generic transport configuration was conducted in the NASA NTF for use in the Drag Prediction Workshop. As data from this experimental investigation was collected, a large difference in drag values was seen between the NTF test and an ONERA test that was conducted several years ago. After much investigation, it was determined that this difference was likely due to a sting effect correction applied to the ONERA data which NTF does not use. This insight led to the present work. In this study, a computational assessment has been undertaken to investigate model support system interference effects on the DLR-F6 transport configuration. The configurations computed during this investigation were the isolated wing-body, the wing-body with the full support system (blade and sting), the wing-body with just the blade, and the wing-body with just the sting. The results from this investigation show the same trends that ONERA saw when they conducted a similar experimental investigation in the S2MA tunnel. Computational results suggest that the blade contributed an interference type of effect, the sting contributed a general blockage effect, and the full support system combined these effects.

Description: Scientists have eagerly anticipated the performance capability of the National Aeronautics and Space Administration (NASA) Global Hawk for over a decade. In 2009 this capability becomes operational. One of the most desired performance capabilities of the Global Hawk aircraft is very long endurance. The Global Hawk aircraft can remain airborne longer than almost all other jet-powered aircraft currently flying, and longer than all other aircraft available for airborne science use. This paper describes the NASA Global Hawk system, payload accommodations, concept of operations, and the first scientific data-gathering mission: Global Hawk Pacific 2009.

Description: Time accurate numerical simulations were performed using the Reynolds-averaged Navier-Stokes (RANS) flow solver OVERFLOW for a heavy lift, slowed-rotor, compound helicopter configuration, tested at the NASA Langley 14- by 22-Foot Subsonic Tunnel. The primary purpose of these simulations is to provide support for the development of a large field of view Particle Imaging Velocimetry (PIV) flow measurement technique supported by the Subsonic Rotary Wing (SRW) project under the NASA Fundamental Aeronautics program. These simulations provide a better understanding of the rotor and body wake flows and helped to define PIV measurement locations as well as requirements for validation of flow solver codes. The large field PIV system can measure the three-dimensional velocity flow field in a 0.914m by 1.83m plane. PIV measurements were performed upstream and downstream of the vertical tail section and are compared to simulation results. The simulations are also used to better understand the tunnel wall and body/rotor support effects by comparing simulations with and without tunnel floor/ceiling walls and supports. Comparisons are also made to the experimental force and moment data for the body and rotor.