ALBERT

Description: The Stratospheric Observatory For Infrared Astronomy (SOFIA) is an international cooperative development and operations program between the United States National Aeronautics and Space Administration (NASA) and the German Space Agency, DLR (Deutsches Zentrum fuer Luft-und Raumfahrt). SOFIA is a 2.5 meter, optical/infrared/sub-millimeter telescope mounted in a Boeing model 747SP-21 aircraft and will be used for many basic astronomical observations performed at stratospheric altitudes. It will accommodate installation of different focal plane instruments with in-flight accessibility provided by investigators selected from the international science community. The Facility operational lifetime is planned to be greater than 20 years. This presentation will present the results of developmental testing of SOFIA, including analysis, envelope expansion and the first operational mission. It will describe a brief history of open cavities in flight, how NASA designed and tested SOFIAs cavity, as well as flight test results. It will focus on how the test team achieved key milestones by systematically and efficiently reducing the number of test points to only those absolutely necessary to achieve mission requirements, thereby meeting all requirements and saving the potential loss of program funding. Finally, it will showcase examples of the observatory in action and the first operational mission of the observatory, illustrating the usefulness of the system to the international scientific community. Lessons learned on how to whittle a mountain of test points into a manageable sum will be presented at the conclusion.

Description: Aircraft design has been progressing toward reduced structural weight to improve fuel efficiency, increase performance, and reduce cost. Lightweight aircraft structures are more flexible than conventional designs and require new design considerations. Intelligent sensing allows for enhanced control and monitoring of aircraft, which enables increased structurally efficiency. The NASA Dryden Flight Research Center (DFRC) has developed an instrumentation system and analysis techniques that combine to make distributed structural measurements practical for lightweight vehicles. Dryden's Fiber Optic Strain Sensing (FOSS) technology enables a multitude of lightweight, distributed surface strain measurements. The analysis techniques, referred to as the Displacement Transfer Functions (DTF) and Load Transfer Functions (LTF), use surface strain values to calculate structural deflections and operational loads. The combined system is useful for real-time monitoring of aeroelastic structures, along with many other applications. This paper describes how the capabilities of the measurement system were demonstrated using subscale test articles that represent simple aircraft structures. Empirical FOSS strain data were used within the DTF to calculate the displacement of the article and within the LTF to calculate bending moments due to loads acting on the article. The results of the tests, accuracy of the measurements, and a sensitivity analysis are presented.

Description: The Space Shuttle Orbiters Discovery and Endeavor have been digitally scanned to produce post-flight configuration outer mold line surfaces. Very detailed scans of the windward side of these vehicles provide resolution of the detailed tile step and gap geometry, as well as the reinforced carbon carbon nose cap and leading edges. Lower resolution scans of the upper surface provide definition of the crew cabin windows, wing upper surfaces, payload bay doors, orbital maneuvering system pods and the vertical tail. The process for acquisition of these digital scans as well as post-processing of the very large data set will be described.

Description: The presentation describes supersonic flight testing accomplished on a novel mixed compression axisymmetric inlet utilizing channels for off design flow matching rather than a translating centerbody concept.

Description: The design and performance of compound helicopters utilizing lift-offset rotors are examined, in the context of short-haul, medium-size civil and military missions. The analysis tools used are the comprehensive analysis CAMRAD II and the sizing code NDARC. Following correlation of the comprehensive analysis with existing lift-offset aircraft flight test data, the rotor performance model for the sizing code was developed, and an initial estimate was made of the rotor size and key hover and cruise flight conditions. The rotor planform and twist were optimized for those conditions, and the sizing code rotor performance model updated. Two models for estimating the blade and hub weight of lift-offset rotors are discussed. The civil and military missions are described, along with the aircraft design assumptions. The aircraft are sized for 30 passengers or 6600 lb payload, with a range of 300 nm. Civil and military aircraft designs are described for each of the rotor weight models. Disk loading and blade loading were varied to optimize the designs, based on gross weight and fuel burn. The influence of technology is shown, in terms of rotor hub drag and rotor weight.

Description: Two full-scale crash tests of an MD-500 helicopter were conducted in 2009 and 2010 at NASA Langley's Landing and Impact Research Facility in support of NASA s Subsonic Rotary Wing Crashworthiness Project. The first crash test was conducted to evaluate the performance of an externally mounted composite deployable energy absorber under combined impact conditions. In the second crash test, the energy absorber was removed to establish baseline loads that are regarded as severe but survivable. Accelerations and kinematic data collected from the crash tests were compared to a system integrated finite element model of the test article. Results from 19 accelerometers placed throughout the airframe were compared to finite element model responses. The model developed for the purposes of predicting acceleration responses from the first crash test was inadequate when evaluating more severe conditions seen in the second crash test. A newly developed model calibration approach that includes uncertainty estimation, parameter sensitivity, impact shape orthogonality, and numerical optimization was used to calibrate model results for the second full-scale crash test. This combination of heuristic and quantitative methods was used to identify modeling deficiencies, evaluate parameter importance, and propose required model changes. It is shown that the multi-dimensional calibration techniques presented here are particularly effective in identifying model adequacy. Acceleration results for the calibrated model were compared to test results and the original model results. There was a noticeable improvement in the pilot and co-pilot region, a slight improvement in the occupant model response, and an over-stiffening effect in the passenger region. This approach should be adopted early on, in combination with the building-block approaches that are customarily used, for model development and test planning guidance. Complete crash simulations with validated finite element models can be used to satisfy crash certification requirements, thereby reducing overall development costs.

Description: This paper investigates the use of proof (or acceptance) test data during the reliability based design optimization of structural components. It is assumed that every component will be proof tested and that the component will only enter into service if it passes the proof test. The goal is to reduce the component weight, while maintaining high reliability, by exploiting the proof test results during the design process. The proposed procedure results in the simultaneous design of the structural component and the proof test itself and provides the designer with direct control over the probability of failing the proof test. The procedure is illustrated using two analytical example problems and the results indicate that significant weight savings are possible when exploiting the proof test results during the design process.

Description: Many theoretical and experimental studies have shown that aircraft flying in formation could experience significant reductions in fuel use compared to solo flight. To date, formation flight for aerodynamic benefit has not been thoroughly explored in flight for large transport-class vehicles. This paper summarizes flight data gathered during several two ship, C-17 formation flights at a single flight condition of 275 knots, at 25,000 ft MSL. Stabilized test points were flown with the trail aircraft at 1,000 and 3,000 ft aft of the lead aircraft at selected crosstrack and vertical offset locations within the estimated area of influence of the vortex generated by the lead aircraft. Flight data recorded at test points within the vortex from the lead aircraft are compared to data recorded at tare flight test points outside of the influence of the vortex. Since drag was not measured directly, reductions in fuel flow and thrust for level flight are used as a proxy for drag reduction. Estimated thrust and measured fuel flow reductions were documented at several trail test point locations within the area of influence of the leads vortex. The maximum average fuel flow reduction was approximately 7-8%, compared to the tare points flown before and after the test points. Although incomplete, the data suggests that regions with fuel flow and thrust reduction greater than 10% compared to the tare test points exist within the vortex area of influence.

Description: This paper describes the interim progress for an in-house study that is directed toward innovative structural analysis and design of next-generation advanced aircraft concepts, such as the Hybrid Wing-Body (HWB) and the Advanced Mobility Concept-X flight vehicles, for structural weight reduction and associated performance enhancement. Unlike the conventional, skin-stringer-frame construction for a cylindrical fuselage, the box-type pressurized fuselage panels in the HWB undergo significant deformation of the outer aerodynamic surfaces, which must be minimized without significant structural weight penalty. Simple beam and orthotropic plate theory is first considered for sizing, analytical verification, and possible equivalent-plate analysis with appropriate simplification. By designing advanced composite stiffened-shell configurations, significant weight reduction may be possible compared with the sandwich and ribbed-shell structural concepts that have been studied previously. The study involves independent analysis of the advanced composite structural concepts that are presently being developed by The Boeing Company for pressurized HWB flight vehicles. High-fidelity parametric finite-element models of test coupons, panels, and multibay fuselage sections, were developed for conducting design studies and identifying critical areas of potential failure. Interim results are discussed to assess the overall weight/strength advantages.

Description: Integrally stitched composite technology is an area that shows promise in enhancing the structural integrity of aircraft and aerospace structures. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The goal of the PRSEUS concept relevant to this test is to provide damage containment capability for composite structures while reducing overall structural weight. The National Aeronautics and Space Administration (NASA), the Federal Aviation Administration (FAA), and The Boeing Company have partnered in an effort to assess the damage containment features of a full-scale curved PRSEUS panel using the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. A single PRSEUS test panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure loads. The test results showed excellent performance of the PRSEUS concept. No growth of Barely Visible Impact Damage (BVID) was observed after ultimate loads were applied. With a two-bay notch severing the central stringer, damage was contained within the two-bay region well above the required limit load conditions. Catastrophic failure was well above the ultimate load level. Information describing the test panel and procedure has been previously presented, so this paper focuses on the experimental procedure, test results, nondestructive inspection results, and preliminary test and analysis correlation.

Description: The goal of this ongoing study is to develop and demonstrate the feasibility of a blade actuation system to dynamically change the twist, and/or the camber, of an airfoil section and, consequently, alter the in-flight aerodynamic loading on the blade for efficient flight control. The required analytical and finite element tools are under development to enable an accurate and comprehensive aeroelastic assessment of the current Full-Blade Warping and 3D Warping Actuated Trailing Edge Flap concepts. The feasibility of the current concepts for swashplateless rotors and higher harmonic blade control is also being investigated. In particular, the aim is to complete the following objectives, some of which have been completed (as noted below) and others that are currently ongoing: i) Develop a Vlasov finite element model and validate against the ABAQUS shell models (completed). ii) Implement the 3D warping actuation concept within the comprehensive analysis code DYMORE. iii) Perform preliminary aeroelastic simulations of blades using DYMORE with 3D warping actuation: a) Investigate the blade behavior under 1 per/rev actuation. Determine whether sufficient twist can be generated and sustained to achieve primary blade control. b) Investigate the behavior of a trailing edge flap configuration under higher harmonic excitations. Determine how much twist can be obtained at the harmonics 2-5 per/rev. iv) Determine actuator specifications such as the power required, load and displacements, and identify the stress and strain distributions in the actuated blades. In general, the completion of Item ii) above will give an additional research capability in rotorcraft dynamics analyses, i.e., the capability to calculate the rotor blade twist due to warping, something that is not currently available in any of the existing comprehensive rotorcraft analyses.

Description: NASA Glenn Research Center, in collaboration with GE Aviation, has begun the development of a smart adaptive structure system with piezoelectric (PE) transducers to improve composite fan blade damping at resonances. Traditional resonant damping approaches may not be realistic for rotating frame applications such as engine blades. The limited space in which the blades reside in the engine makes it impossible to accommodate the circuit size required to implement passive resonant damping. Thus, a novel digital shunt scheme has been developed to replace the conventional electric passive shunt circuits. The digital shunt dissipates strain energy through the load resistor on a power amplifier. General Electric (GE) designed and fabricated a variety of polymer matrix fiber composite (PMFC) test specimens. Investigating the optimal topology of PE sensors and actuators for each test specimen has revealed the best PE transducer location for each target mode. Also a variety of flexible patches, which can conform to the blade surface, have been tested to identify the best performing PE patch. The active damping control achieved significant performance at target modes. This work has been highlighted by successful spin testing up to 5000 rpm of subscale GEnx composite blades in Glenn s Dynamic Spin Rig.

Description: No abstract available

Description: The Environmentally Responsible Aviation project seeks to accomplish the simultaneous reduction of fuel burn, noise, and emissions. A project at NASA Dryden Flight Research Center is contributing to ERAs goals by exploring the practical application of real-time trim configuration optimization for enhanced performance and reduced fuel consumption. This peak-seeking control approach is based on Newton-Raphson algorithm using a time-varying Kalman filter to estimate the gradient of the performance function. In real-time operation, deflection of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of a modified F-18 are directly optimized, and the horizontal stabilators and angle of attack are indirectly optimized. Preliminary results from three research flights are presented herein. The optimization system found a trim configuration that required approximately 3.5% less fuel flow than the baseline trim at the given flight condition. The algorithm consistently rediscovered the solution from several initial conditions. These preliminary results show the algorithm has good performance and is expected to show similar results at other flight conditions and aircraft configurations.

Description: The presentation will provide a brief overview of the project with specific focus on Dryden's role and capabilities exercised during X-43A.

Description: The primary focus of this paper is how the flight test team for the Stratospheric Observatory For Infrared Astronomy (SOFIA) re-cast an extensive developmental test program to meet key milestones while simultaneously ensuring safe certification of the airframe and delivery of an operationally relevant platform, ultimately saving the overall program from financial demise. Following a brief introduction to the observatory and what it is designed to do, SOFIAs planned developmental test program is summarized, including analysis and design philosophy, envelope expansion, model validation and airframe certification. How NASA used lessons learned from other aircraft that employed open cavities in flight is explained as well as how and why the chosen design was selected. The approach to aerodynamic analysis, including bare airframe testing, wind tunnel testing, computational fluid dynamics and finite element modeling proved absolutely critical. Despite a solid analytical foundation, many unknowns remained. History provides several examples of disastrous effects on both systems and flight safety if cavity design is not approached properly. For these reasons, an extensive test plan was developed to ensure a safe and thorough build-up for envelope expansion, airframe certification and early science missions. Unfortunately, as is often the case, because of chronic delays in overall program execution, severe schedule and funding pressures were present. If critical milestones were not met, domestic as well as international funding was in serious jeopardy, and the demise of the entire program loomed large. Concentrating on rigorous model validation, the test team challenged certification requirements, increased test efficiency and streamlined engineering analysis. This resulted in the safe reduction of test point count by 72%, meeting all program milestones and a platform that soundly satisfied all operational science requirements. Results from early science missions are shown and a proof of concept mission for which SOFIA was opportunely positioned is showcased. Success on this time-critical mission to observe a rare astronomical event proved the usefulness of an airborne observatory and the value in waiting for the capability provided by SOFIA. Finally, lessons learned in the test program are presented with emphasis on how lessons from previous aircraft and successful test programs were applied to SOFIA. Effective application of these lessons was crucial to the success of the SOFIA flight test program. SOFIA is an international cooperative program between NASA and the German Space Agency, DLR. It is a 2.5 meter (100-inch) telescope mounted in a Boeing 747SP aircraft used for astronomical observations at altitudes above 35,000 feet. SOFIA will accommodate a host of scientific instruments from the international science community and has a planned operational lifespan of more than 20 years.

Description: The Synthetic and Enhanced Vision Systems for NextGen (SEVS) simulation and flight tests are jointly sponsored by NASA's Aviation Safety Program, Vehicle Systems Safety Technology project and the Federal Aviation Administration (FAA). The flight tests were conducted by a team of Honeywell, Gulfstream Aerospace Corporation and NASA personnel with the goal of obtaining pilot-in-the-loop test data for flight validation, verification, and demonstration of selected SEVS operational and system-level performance capabilities. Nine test flights (38 flight hours) were conducted over the summer and fall of 2011. The evaluations were flown in Gulfstream.s G450 flight test aircraft outfitted with the SEVS technology under very low visibility instrument meteorological conditions. Evaluation pilots flew 108 approaches in low visibility weather conditions (600 ft to 2400 ft visibility) into various airports from Louisiana to Maine. In-situ flight performance and subjective workload and acceptability data were collected in collaboration with ground simulation studies at LaRC.s Research Flight Deck simulator.

Description: An area that shows promise in enhancing structural integrity of aircraft and aerospace structures is the integrally stitched composite technology. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept developed by Boeing Research and Technology and the National Aeronautics and Space Administration. A joint test program on the assessment of damage containment capabilities of the PRSEUS concept for curved fuselage structures was conducted recently at the Federal Aviation Administration William J. Hughes Technical Center. The panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure load conditions up to fracture, with a through-the-thickness, two-bay notch severing the central stiffener. For the purpose of future progressive failure analysis development and verification, extensive post failure nondestructive and teardown inspections were conducted. Detailed inspections were performed directly ahead of the notch tip where stable damage progression was observed. These examinations showed: 1) extensive delaminations developed ahead of the notch tip, 2) the extent and location of damage, 3) the typical damage mechanisms observed in composites, and 4) the role of stitching and warp-knitting in the failure mechanisms. The objective of this paper is to provide a summary of results from these posttest inspections.

Description: No abstract available

Description: As unmanned air vehicles (UAVs) continue to expand their flight envelopes into areas of high angular rate and high angle of attack, modeling the complex unsteady aerodynamics for simulation in these regimes has become more difficult using traditional methods. The goal of this experiment was to improve the current six degree-of-freedom aerodynamic model of a small UAV by replacing the analytically derived damping derivatives with experimentally derived values. The UAV is named the Free-flying Aircraft for Sub-scale Experimental Research, FASER, and was tested in the NASA Langley Research Center 12- Foot Low-Speed Tunnel. The forced oscillation wind tunnel test technique was used to measure damping in the roll and yaw axes. By imparting a variety of sinusoidal motions, the effects of non-dimensional angular rate and reduced frequency were examined over a large range of angle of attack and side-slip combinations. Tests were performed at angles of attack from -5 to 40 degrees, sideslip angles of -30 to 30 degrees, oscillation amplitudes from 5 to 30 degrees, and reduced frequencies from 0.010 to 0.133. Additionally, the effect of aileron or elevator deflection on the damping coefficients was examined. Comparisons are made of two different data reduction methods used to obtain the damping derivatives. The results show that the damping derivatives are mainly a function of angle of attack and have dependence on the non-dimensional rate and reduced frequency only in the stall/post-stall regime

Description: Measured data from the slowed rotor part of the 2010 UH-60A Airloads Rotor test in the NASA Ames 40- by 80- Foot Wind Tunnel are compared with CAMRAD II calculations. The emphasis in this initial study is to correlate overall trends. This analytical effort considers advance ratios from 0.3 to 1.0, with the rotor rotational speed at 40%NR. The rotor performance parameters considered are the thrust coefficient, power coefficient, L/DE, torque, and H-force. The blade loads considered are the half peak-to-peak, mid-span and outboard torsion, flatwise, and chordwise moments, and the pitch link load. For advance ratios . 0.7, the overall trends for the performance and loads (excluding the pitch link load) could be captured, but with substantial overprediction or underprediction. The correlation gradually deteriorates as the advance ratio is increased and for advance ratios . 0.8 there is no correlation. The pitch link load correlation is not good. There is considerable scope for improvement in the prediction of the blade loads. Considering the modeling complexity associated with the unconventional operating condition under consideration, the current predictive ability to capture overall trends is encouraging.

Description: Supersonic aircraft designers must shape the outer mold line of the aircraft to improve multiple objectives, such as mission performance, cruise efficiency, and sonic-boom signatures. Conceptual designers have demonstrated an ability to assess these objectives for a large number of candidate designs. Other critical objectives and constraints, such as weight, fuel volume, aeroelastic effects, and structural soundness, are more difficult to address during the conceptual design process. The present research adds both static structural analysis and sizing to an existing conceptual design framework. The ultimate goal is to include structural analysis in the multidisciplinary optimization of a supersonic aircraft. Progress towards that goal is discussed and demonstrated.

Description: Technical challenges of compressors for future rotorcraft engines are driven by engine-level and component-level requirements. Cycle analyses are used to highlight the engine-level challenges for 3000, 7500, and 12000 SHP-class engines, which include retention of performance and stability margin at low corrected flows, and matching compressor type, axial-flow or centrifugal, to the low corrected flows and high temperatures in the aft stages. At the component level: power-to-weight and efficiency requirements impel designs with lower inherent aerodynamic stability margin; and, optimum engine overall pressure ratios lead to small blade heights and the associated challenges of scale, particularly increased clearance-to-span ratios. The technical challenges associated with the aerodynamics of low corrected flows and stability management impel the compressor aero research and development efforts reviewed herein. These activities include development of simple models for clearance sensitivities to improve cycle calculations, full-annulus, unsteady Navier-Stokes simulations used to elucidate stall, its inception, and the physics of stall control by discrete tip-injection, development of an actuator-duct-based model for rapid simulation of nonaxisymmetric flow fields (e.g., due inlet circumferential distortion), advanced centrifugal compressor stage development and experimentation, and application of stall control in a T700 engine.

Description: NASA sets aggressive, strategic, civil aircraft performance and environmental goals and develops ambitious technology roadmaps to guide its research efforts. NASA has adopted a phased approach for community noise reduction of civil aircraft. While the goal of the near-term first phase focuses primarily on source noise reduction, the goal of the second phase relies heavily on presumed architecture changes of future aircraft. The departure from conventional airplane configurations to designs that incorporate some type of propulsion noise shielding is anticipated to provide an additional 10 cumulative EPNdB of noise reduction. One candidate propulsion system for these advanced aircraft is the open rotor engine. In some planned applications, twin open rotor propulsors are located on the aft fuselage, with the vehicle s empennage shielding some of their acoustic signature from observers on the ground. This study focuses on predicting the noise certification benefits of a notional open rotor aircraft with tail structures shielding a portion of the rotor noise. The measured noise of an open rotor test article--collected with and without an acoustic barrier wall--is the basis of the prediction. The results are used to help validate NASA s reliance on acoustic shielding to achieve the second phase of its community noise reduction goals. The noise measurements are also compared to a popular empirical diffraction correlation often used at NASA to predict acoustic shielding.

Description: One of the fundamental problems in flight dynamics is the formulation of aerodynamic forces and moments acting on an aircraft in arbitrary motion. Classically, conventional stability derivatives are used for the representation of aerodynamic loads in the aircraft equations of motion. However, for modern aircraft with highly nonlinear and unsteady aerodynamic characteristics undergoing maneuvers at high angle of attack and/or angular rates the conventional stability derivative model is no longer valid. Attempts to formulate aerodynamic model equations with unsteady terms are based on several different wind tunnel techniques: for example, captive, wind tunnel single degree-of-freedom, and wind tunnel free-flying techniques. One of the most common techniques is forced oscillation testing. However, the forced oscillation testing method does not address the systematic and systematic correlation errors from the test apparatus that cause inconsistencies in the measured oscillatory stability derivatives. The primary objective of this study is to identify the possible sources and magnitude of systematic error in representative dynamic test apparatuses. Sensitivities of the longitudinal stability derivatives to systematic errors are computed, using a high fidelity simulation of a forced oscillation test rig, and assessed using both Design of Experiments and Monte Carlo methods.

Description: In a previous study by the authors it was shown that the N3-X, a 300 passenger hybrid wing body (HWB) aircraft with a turboelectric distributed propulsion (TeDP) system, was able to meet the NASA Subsonic Fixed Wing (SFW) project goal for N+3 generation aircraft of at least a 60% reduction in total energy consumption as compared to the best in class current generation aircraft. This previous study combined technology assumptions that represented the highest anticipated values that could be matured to technology readiness level (TRL) 4-6 by 2030. This paper presents the results of a sensitivity analysis of the total mission energy consumption to reductions in each key technology assumption. Of the parameters examined, the mission total energy consumption was most sensitive to changes to total pressure loss in the propulsor inlet. The baseline inlet internal pressure loss is assumed to be an optimistic 0.5%. An inlet pressure loss of 3% increases the total energy consumption 9%. However changes to reduce inlet pressure loss can result in additional distortion to the fan which can reduce fan efficiency or vice versa. It is very important that the inlet and fan be analyzed and optimized as a single unit. The turboshaft hot section is assumed to be made of ceramic matrix composite (CMC) with a 3000 F maximum material temperature. Reducing the maximum material temperature to 2700 F increases the mission energy consumption by only 1.5%. Thus achieving a 3000 F temperature in CMCs is important but not central to achieving the energy consumption objective of the N3-X/TeDP. A key parameter in the efficiency of superconducting motors and generators is the size of the superconducting filaments in the stator. The size of the superconducting filaments in the baseline model is assumed to be 10 microns. A 40 micron filament, which represents current technology, results in a 200% increase in AC losses in the motor and generator stators. This analysis shows that for a system with 40 micron filaments the higher stator losses plus the added weight and power of larger cryocoolers results in a 4% increase in mission energy consumption. If liquid hydrogen is used to cool the superconductors the 40 micron fibers results in a 200% increase in hydrogen required for cooling. Each pound of hydrogen used as fuel displaces 3 pounds of jet fuel. For the N3-X on the reference mission the additional hydrogen due to the increase stator losses reduces the total fuel weight 10%. The lighter fuel load and attendant vehicle resizing reduces the total energy consumption more than the higher stator losses increase it. As a result with hydrogen cooling there is a slight reduction in mission energy consumption with increasing stator losses. This counter intuitive result highlights the need to consider the full system impact of changes rather than just at the component or subsystem level.

Description: An 11% scale-model of a Cruise-Efficient Short Take-off and Landing (CESTOL) scalemodel test was recently completed. The test was conducted in the AEDC National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center. The model included two over-wing pod-mounted turbine propulsion simulators (TPS). The hybrid blended wing-body used a circulation control wing (CCW) with leadingand trailing-edge blowing. The bulk of the test matrix included three forward velocities (40 kts, 60 kts, and 100kts), angle-of-attack variation between -5 and 25 , and CCW mass flow variation. Seven strut-mounted microphones outboard of the left wing provided source directivity. A phased microphone array was mounted outboard of the right wing for source location. The goal of this paper is to provide a preliminary look at the acoustic data acquired during the Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA) test for 0 angle-of-attack and 0 sideslip conditions. Data presented provides a good overview of the test conditions and the signal-to-noise quality of the data. TPS height variation showed a difference of 2 dB to 3 dB due to wing shielding. Variation of slot mass flow showed increases of 12 dB to 26 dB above the airframe noise and the TPS increased the overall levels an additional 5 dB to 10 dB.

Description: A study was undertaken at NASA Langley Research Center to establish, demonstrate, and apply methodology for modeling and implementing the aerodynamic effects of MANPADS damage to a transport aircraft into real-time flight simulation, and to demonstrate a preliminary capability of using such a simulation to conduct an assessment of aircraft survivability. Key findings from this study include: superpositioning of incremental aerodynamic characteristics to the baseline simulation aerodynamic model proved to be a simple and effective way of modeling damage effects; the primary effect of wing damage rolling moment asymmetry may limit minimum airspeed for adequate controllability, but this can be mitigated by the use of sideslip; combined effects of aerodynamics, control degradation, and thrust loss can result in significantly degraded controllability for a safe landing; and high landing speeds may be required to maintain adequate control if large excursions from the nominal approach path are allowed, but high-gain pilot control during landing can mitigate this risk.

Description: A conceptual design study of a large civil compound helicopter is presented. The objective is to determine how a compound helicopter performs when compared to both a conventional helicopter and a tiltrotor using a design mission that is shorter than optimal for a tiltrotor and longer than optimal for a helicopter. The designs are generated and analyzed using conceptual design software and are further evaluated with a comprehensive rotorcraft analysis code. Multiple metrics are used to determine the suitability of each design for the given mission. Plots of various trade studies and parameter sweeps as well as comprehensive analysis results are presented. The results suggest that the compound helicopter examined for this study would not be competitive with a tiltrotor or conventional helicopter, but multiple possibilities are identified for improving the performance of the compound helicopter in future research.

Description: The Large Civil Tiltrotor (LCTR) was developed as part of the NASA Heavy Lift Rotorcraft Systems Investigation in order to establish a consistent basis for evaluating the benefits of advanced technology for large tiltrotors. The concept has since evolved into the second-generation LCTR2, designed to carry 90 passengers for 1,000 nm at 300 knots, with vertical takeoff and landing. This paper examines the impact of advanced propulsion system concepts on LCTR2 sizing. Two concepts were studied: an advanced, single-speed engine with a conventional power turbine layout (Advanced Conventional Engine, or ACE), and a variable-speed power turbine engine (VSPT). The ACE is the lighter engine, but requires a multi-speed (shifting) gearbox, whereas the VSPT uses a lighter, fixed-ratio gearbox. The NASA Design and Analysis of Rotorcraft (NDARC) design code was used to study the trades between rotor and engine efficiency and weight. Rotor performance was determined by Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD II), and engine performance was estimated with the Numerical Propulsion System Simulation (NPSS). Design trades for the ACE vs. VSPT are presented in terms of vehicle weight empty for variations in mission altitude and range; the effect of different One Engine Inoperative (OEI) criteria are also examined. Because of its strong effect on gearbox weight and on both rotor and engine efficiency, rotor speed was chosen as the reference design variable for comparing design trades. The two propulsion concepts had nearly identical vehicle weights and mission fuel consumption, and their relative advantages varied little with cruise altitude, mission range, or OEI criteria; high cruise altitude and low cruise tip speed were beneficial for both concepts.

Description: In preparation for upcoming wind tunnel acoustic experiments of a Hybrid Wing Body (HWB) vehicle with two jet engine simulator units, a series of twin jet aeroacoustic investigations were conducted leading to increased understanding and risk mitigation. A previously existing twin jet nozzle system and a fabricated HWB aft deck fuselage are combined for a 1.9% model scale study of jet nozzle spacing and jet cant angle effects, elevon deflection into the jet plume, and acoustic shielding by the fuselage body. Linear and phased array microphone measurements are made, and data processing includes the use of DAMAS (Deconvolution Approach for the Mapping of Acoustic Sources). Closely-spaced twin jets with a 5 inward cant angle exhibit reduced noise levels compared to their parallel flow counterparts at similar and larger nozzle spacings. A 40 elevon deflection into the twin jet plume, which is required for HWB ground rotation, can significantly increase upstream noise levels (more than 5 dB OASPL) with only minimal increases in the downstream direction. Lastly, DAMAS processing can successfully measure the noise source distribution of multiple shielded jet sources.

Description: The Hybrid Wing Body (HWB) configuration is a subsonic transport aircraft concept with the potential to simultaneously reduce fuel burn, noise and emissions compared to conventional concepts. Initial studies focused on very large applications with capacities for up to 800 passengers. More recent studies have focused on the large, twin-aisle class with passenger capacities in the 300-450 range. Efficiently scaling this concept down to the single aisle or smaller size is challenging due to geometric constraints, potentially reducing the desirability of this concept for applications in the 100-200 passenger capacity range or less. In order to quantify this scaling challenge, five advanced conventional (tube-and-wing layout) concepts were developed, along with equivalent (payload/range/technology) HWB concepts, and their fuel burn performance compared. The comparison showed that the HWB concepts have fuel burn advantages over advanced tube-and-wing concepts in the larger payload/range classes (roughly 767-sized and larger). Although noise performance was not quantified in this study, the HWB concept has distinct noise advantages over the conventional tube-and-wing configuration due to the inherent noise shielding features of the HWB. NASA s Environmentally Responsible Aviation (ERA) project will continue to investigate advanced configurations, such as the HWB, due to their potential to simultaneously reduce fuel burn, noise and emissions.

Description: In this paper we describe efforts to obtain canonical data sets to assist computational modelers in their development of models for the prediction of mixing and combustion in scramjet combustors operating in the ramjet-scramjet transition regime. The CARS technique is employed to acquire temporally and spatially resolved measurements of temperature and species mole-fraction at four planes, one upstream of an H2 fuel injector and three downstream. The technique is described and results are presented for cases with and without chemical reaction. The vibrational energy mode in the heated airstream of the combustor was observed to be frozen at near facility heater conditions and significant nonuniformities in temperature were observed, attributed to nonuniformities of temperature exiting the heater. The measurements downstream of fuel injection show development of mixing and combustion, and are already proving useful to the modelers.

Description: This paper discusses a proposed framework for the safe integration of small unmanned aerial systems (sUAS) into the National Airspace System (NAS). The paper briefly examines the potential uses of sUAS to build an understanding of the location and frequency of potential future flight operations based on the future applications of the sUAS systems. The paper then examines the types of systems that would be required to meet the application-level demand to determine "classes" of platforms and operations. A framework for categorization of the "intelligence" level of the UAS is postulated for purposes of NAS integration. Finally, constraints on the intelligent systems are postulated to ensure their ease of integration into the NAS.

Description: A rapid shape parameterization tool called PROTEUS is developed for aircraft shape optimization. This tool can be applied directly to any aircraft geometry that has been defined in PLOT3D format, with the restriction that each aircraft component must be defined by only one data block. PROTEUS has eight types of parameterization schemes: planform, wing surface, twist, body surface, body scaling, body camber line, shifting/scaling, and linear morphing. These parametric schemes can be applied to two types of components: wing-type surfaces (e.g., wing, canard, horizontal tail, vertical tail, and pylon) and body-type surfaces (e.g., fuselage, pod, and nacelle). These schemes permit the easy setup of commonly used shape modification methods, and each customized parametric scheme can be applied to the same type of component for any configuration. This paper explains the mathematics for these parametric schemes and uses two supersonic configurations to demonstrate the application of these schemes.

Description: No abstract available

Description: Results of the Aeronautics Research Mission Directorate Seedling Project Phase I research project entitled "Nonlinear Aerodynamics Modeling using Fuzzy Logic" are presented. Efficient and rapid flight test capabilities were developed for estimating highly nonlinear models of airplane aerodynamics over a large flight envelope. Results showed that the flight maneuvers developed, used in conjunction with the fuzzy-logic system identification algorithms, produced very good model fits of the data, with no model structure inputs required, for flight conditions ranging from cruise to departure and spin conditions.

Description: No abstract available

Description: Federal Aviation Administration will certainly require the Heavy Lift Rotorcraft to be operated under Category A performance and operations requirements. Because of the weight, no operation will be allowed except Category A according to FAA Part 29.1(c). This means that any where along the flight path, the aircraft must be able to land safely following an engine failure or continue flight. A repeatable flight profile must be developed and executed to ensure that the aircraft can be safely landed or flown away depending on its location on the flight profile. This means that there will be no Height-Velocity testing required as is currently required for Part 29 Category B. Since all the configurations shown to date are different than existing rotorcraft, each type would have to develop their individual requirements under existing special conditions FAA Part 21.17(b). This means the FAA will take the opportunity to negotiate additional requirements or change requirements to ensure safety. For example, since the tiltrotor did not fit normal rotorcraft category, new rules were negotiated between the applicant and the FAA. As a result of this negotiation, performance requirements for Category A were increased. The rules were written in terms of guaranteed performance instead of Category A requirements. Detailed discussion will follow later. The proposed tiltrotor would likely follow along with the current tiltrotor rules with the possibility of increase Category A performance requirements. Compounding with addition of wing and auxiliary thrust to both the tandem and coaxial rotor would result in new special condition aircraft. To my knowledge, no compound tandem or compound coaxial rotor has ever been certified by FAA.

Description: Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter. Projections near a support pylon and/or associated heat shield can have particular configurations, including greater flow immersion than other projections.

Description: During mass properties testing on the X-48B Blended Wing Body aircraft (The Boeing Company, Chicago, Illinois) at the National Aeronautics and Space Administration Dryden Flight Research Center, Edwards, California, large inertia measurement errors were observed in results from compound pendulum swings when compared to analytical models. By comparing periods of oscillations as measured from an average over the test period versus the period of each oscillation, it was noticed that the frequency of oscillation was shifting significantly throughout the test. This phenomenon was only noticed during compound pendulum swings, and not during bifilar pendulum swings. The frequency shift was only visible upon extensive data analysis of the frequency for each oscillation, and did not appear in averaged frequency data over the test period. Multiple test articles, test techniques, and hardware setups were used in attempts to eliminate or identify the cause of the frequency shift. Plotting the frequency of oscillation revealed a region of minimal shift that corresponded to a larger amplitude range. This region of minimal shift provided the most accurate results compared to a known test article; however, the amplitudes that produce accurate inertia measurements are amplitudes larger than those generally accepted in mass properties testing. This paper examines two case studies of the frequency shift, using mass properties testing performed on a dummy test article, and on the X-48B Blended Wing Body aircraft.

Description: A vane has an airfoil shell and a spar within the shell. The vane has an outboard shroud at an outboard end of the shell and an inboard platform at an inboard end of the shell. The shell includes a region having a depth-wise coefficient of thermal expansion and a second coefficient of thermal expansion transverse thereto, the depth-wise coefficient of thermal expansion being greater than the second coefficient of thermal expansion.

Description: Considerable interest surrounds the design of the next generation of single-aisle commercial transports in the Boeing 737 and Airbus A320 class. Aircraft designers will depend on advanced, next-generation turbofan engines to power these airplanes. The focus of this study is to apply single- and multi-objective optimization algorithms to the conceptual design of ultrahigh bypass turbofan engines for this class of aircraft, using NASA s Subsonic Fixed Wing Project metrics as multidisciplinary objectives for optimization. The independent design variables investigated include three continuous variables: sea level static thrust, wing reference area, and aerodynamic design point fan pressure ratio, and four discrete variables: overall pressure ratio, fan drive system architecture (i.e., direct- or gear-driven), bypass nozzle architecture (i.e., fixed- or variable geometry), and the high- and low-pressure compressor work split. Ramp weight, fuel burn, noise, and emissions are the parameters treated as dependent objective functions. These optimized solutions provide insight to the ultrahigh bypass engine design process and provide information to NASA program management to help guide its technology development efforts.

Description: MATLAB Stability and Control Toolbox (MASCOT) utilizes geometric, aerodynamic, and inertial inputs to calculate air vehicle stability in a variety of critical flight conditions. The code is based on fundamental, non-linear equations of motion and is able to translate results into a qualitative, graphical scale useful to the non-expert. MASCOT was created to provide the conceptual aircraft designer accurate predictions of air vehicle stability and control characteristics. The code takes as input mass property data in the form of an inertia tensor, aerodynamic loading data, and propulsion (i.e. thrust) loading data. Using fundamental nonlinear equations of motion, MASCOT then calculates vehicle trim and static stability data for the desired flight condition(s). Available flight conditions include six horizontal and six landing rotation conditions with varying options for engine out, crosswind, and sideslip, plus three take-off rotation conditions. Results are displayed through a unique graphical interface developed to provide the non-stability and control expert conceptual design engineer a qualitative scale indicating whether the vehicle has acceptable, marginal, or unacceptable static stability characteristics. If desired, the user can also examine the detailed, quantitative results.

Description: Plans are underway to convert an A-10 combat attack aircraft into a research aircraft for thunderstorm research. This aircraft would be configured and instrumented for flights into large, convective thunderstorms. It would have the capabilities of higher altitude performance and protection for thunderstorm conditions that exceed those of aircraft now in use for this research. One area of investigation for this aircraft would be terrestrial gamma ]ray flashes (TGFs), building on the pioneering observations made by the Airborne Detector for Energetic Lightning Emissions (ADELE) project several years ago. A new and important component of the planned investigations are the continuous, detailed correlations of TGFs with the electric fields near the aircraft, as well as detailed measurements of nearby lightning discharges. Together, the x-and gamma-radiation environments, the electric field measurements, and the lightning observations (all measured on microsecond timescales) should provide new insights into this TGF production mechanism. The A -10 aircraft is currently being modified for thunderstorm research. It is anticipated that the initial test flights for this role will begin next year.

Description: It has been more than 50 years since NASA engineers have attempted to design a manned space vehicle with the capability to return from beyond low Earth orbit. In this interval, our methodologies for designing the thermal protection system (TPS) to protect humans from the extremely high temperatures of re-entry have changed significantly. With these considerations in mind, we return to the Apollo IV (AS-501) flight data. This incredible data set allows us to assess the current tools and methodologies being used to design Orion MPCV. In particular, our ability to predict the aftbody separated region convective heating environments for MPCV is critical. The design uses reusable TPS in this area, whereas Apollo designers used ablative TPS which can withstand much more severe environments. This presentation will revisit the flight data, summarize the assumptions going into the analysis, present the results and draw conclusions regarding how accurately we can currently predict the heating in the aftbody separated region of a re-entry capsule.

Description: A flight experiment was conducted to investigate the pressure distribution, local-flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.00. The tests used a NASA testbed aircraft with a bottom centerline mounted test fixture. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating. Boundary-layer transition was captured in both analog and digital formats using an onboard infrared imaging system. Surface pressures were measured on the surface of the flat plate. Flow field measurements near the leading edge of the test fixture revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration.

Description: This report will present the results obtained from the Pressure Sensitive Paint (PSP) technique on a circulation control concept model. This test was conducted at the National Transonic Facility (NTF) at the NASA Langley Research Center. PSP was collected on the upper wing surface while the facility was operating in cryogenic mode at 227 K (-50 oF). The test envelope for the PSP portion included Mach numbers from 0.7 to 0.8 with angle of attack varying between 0 and 8 degrees and a total pressure of approximately 168 kPa (24.4 psi), resulting in a chord Reynolds number of approximately 15 million. While the PSP results did exhibit high levels of noise in certain conditions (where the oxygen content of the flow was very small), some conditions provided good correlation between the PSP and pressure taps, showing the ability of the PSP technique. This work also served as a risk reduction opportunity for future testing in cryogenic conditions at the NTF.

Description: This final report documents the work of the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team on Task 1 of the Phase II effort. The team consisted of Boeing Research and Technology, Boeing Commercial Airplanes, General Electric, and Georgia Tech. Using a quantitative workshop process, the following technologies, appropriate to aircraft operational in the N+4 2040 timeframe, were identified: Liquefied Natural Gas (LNG), Hydrogen, fuel cell hybrids, battery electric hybrids, Low Energy Nuclear (LENR), boundary layer ingestion propulsion (BLI), unducted fans and advanced propellers, and combinations. Technology development plans were developed.

Description: This presentation describes the current status of the joint NASA/Boeing collaboration on the development of a variable camber continuous trailing edge flap system for use in wing shaping control for cruise drag reduction.

Description: An aircraft is configured for flight in an atmosphere having a low density. The aircraft includes a fuselage, a pair of wings, and a rear stabilizer. The pair of wings extends from the fuselage in opposition to one another. The rear stabilizer extends from the fuselage in spaced relationship to the pair of wings. The fuselage, the wings, and the rear stabilizer each present an upper surface opposing a lower surface. The upper and lower surfaces have X, Y, and Z coordinates that are configured for flight in an atmosphere having low density.

Description: Embodiments of the invention permit flight paths (current and planned) to be viewed from various orientations to provide improved path and terrain awareness via graphical two-dimensional or three-dimensional perspective display formats. By coupling the flight path information with a terrain database, uncompromising terrain awareness relative to the path and ownship is provided. In addition, missed approaches, path deviations, and any navigational path can be reviewed and rehearsed before performing the actual task. By rehearsing a particular mission, check list items can be reviewed, terrain awareness can be highlighted, and missed approach procedures can be discussed by the flight crew. Further, the use of Controller Pilot Datalink Communications enables data-linked path, flight plan changes, and Air Traffic Control requests to be integrated into the flight display of the present invention.

Description: This presentation will give information on Multi-Disciplinary Analysis and Technology Development, including it's objectives and how they will be met. In addition, it will also present recent highlights including the Lift-Offset Civil Design and it's study conclusions, as well as, the LCTR2 Propulsion Concept's study conclusions. Recent publications and future publications will also be discussed.

Description: This presentation is part of panel discussion on Evaluating Models of Human Performance. The purpose of this panel is to discuss the increasing use of models in the world today and specifically focus on how to describe and evaluate models of human performance. My presentation will focus on discussions of generating distributions of performance, and the evaluation of different strategies for humans performing tasks with mixed initiative (Human-Automation) systems. I will also discuss issues with how to provide Human Performance modeling data to support decisions on acceptability and tradeoffs in the design of safety critical systems. I will conclude with challenges for the future.

Description: This presentation is part of panel discussion on Cognitive Systems Engineering. The purpose of this panel is to discuss the challenges and future directions of Cognitive Systems Engineering for the next 30 years. I intended to present the work we have been doing with the Aviation Safety program and Space Human Factors Engineering project on Work Domain Analysis and some areas of Research Focus. Specifically, I intend to focus on the shift on the need to understand and model attention in mixed-initiative systems, the need for methods which can generate results to be used in trade-off decisions, and the need to account for a range of human behavior in the design.

Description: Currently, the US Military and NASA are investigating the feasibility of a Vertical/Short Take Off and Landing (VSTOL) aircraft that can provide invaluable aid in the combat theater and significantly improve the civil transportation system. The nominal military mission requirement calls for a 28-ton payload heavy lift capability while the civilian requirements calls for a 90-passenger, 1000-nm range, airliner, as noted in Reference 1. To aid in these aircraft requirements, the present study examined the conversion of a Lockheed Martin C-130 Hercules into a compound aircraft, which would demonstrate the technology required by a much larger version. The present study examined various configurations and rotor blade designs in order to fulfill the nominal mission described previously. It was shown that the initial design of a 180 ft diameter rotor to lift 155,000 lb was not feasible due to material constraints. A revised design, in which the rotor radius was reduced to 55 ft, met the given constraints but required too much power. The decision was made to move to a twin rotor compound to take advantage of the increased disc area and drop the need for anti torque devices. Following this design shift, a new design point was found where all five constraints were met and the power requirements were deemed reasonable. This twin-rotor design was used in NDARC to provide a complete sizing analysis of the chosen design point.

Description: Computational optimization of a natural laminar flow experimental wing glove that is mounted on a business jet is presented and discussed. The process of designing a laminar flow wing glove starts with creating a two-dimensional optimized airfoil and then lofting it into a three-dimensional wing glove section. The airfoil design process does not consider the three dimensional flow effects such as cross flow due wing sweep as well as engine and body interference. Therefore, once an initial glove geometry is created from the airfoil, the three dimensional wing glove has to be optimized to ensure that the desired extent of laminar flow is maintained over the entire glove. TRANAIR, a non-linear full potential solver with a coupled boundary layer code was used as the main tool in the design and optimization process of the three-dimensional glove shape. The optimization process uses the Class-Shape-Transformation method to perturb the geometry with geometric constraints that allow for a 2-in clearance from the main wing. The three-dimensional glove shape was optimized with the objective of having a spanwise uniform pressure distribution that matches the optimized two-dimensional pressure distribution as closely as possible. Results show that with the appropriate inputs, the optimizer is able to match the two dimensional pressure distributions practically across the entire span of the wing glove. This allows for the experiment to have a much higher probability of having a large extent of natural laminar flow in flight.

Description: NASA Dryden has been engaged in exciting work that will enable lighter weight and more fuel efficient vehicles through advanced control and dynamics technologies. The main areas of emphasis are Enabling Light-weight Flexible Structures, real time control surface optimization for fuel efficiency and autonomous formation flight. This presentation provides a description of the current and upcoming work in these areas. Additionally, status is provided Dryden's work on HTV-2.

Description: Militarized versions of commercial platforms are growing in popularity due to many logistical benefits in the form of commercial off-the-shelf (COTS) parts, established production methods, and commonality for different certifications. Commercial data and best practices are often leveraged to reduce procurement and engineering development costs. While the developmental and cost reduction benefits are clear, these militarized aircraft are routinely operated in flight at significantly different conditions and in significantly different manners than for routine commercial flight. Therefore they are at a higher risk of flight envelope exceedance. This risk may lead to departure from controlled flight and/or aircraft loss1. Historically, the risk of departure from controlled flight for military aircraft has been mitigated by piloted simulation training and engineering analysis of typical aircraft response. High-agility military aircraft simulation databases are typically developed to include high angles of attack (AoA) and sideslip due to the dynamic nature of their missions and have been developed for many tactical configurations over the previous decades. These aircraft simulations allow for a more thorough understanding of the vehicle flight dynamics characteristics at high AoA and sideslip. In recent years, government sponsored research on transport airplane aerodynamic characteristics at high angles of attack has produced a growing understanding of stall/post-stall behavior. This research along with recent commercial airline training initiatives has resulted in improved understanding of simulator-based training requirements and simulator model fidelity.2-5 In addition, inflight training research over the past decade has produced a database of pilot performance and recurrency metrics6. Innovative solutions to aerodynamically model large commercial aircraft for upset conditions such as high AoA, high sideslip, and ballistic damage, as well as capability to accurately account for scaling factors, is necessary to develop realistic engineering and training simulations. Such simulations should significantly reduce the risk of departure from controlled flight, loss of aircraft, and ease the airworthiness certification process. The characteristics of commercial derivative aircraft are exemplified by the P-8A Multi-mission Maritime Aircraft (MMA) aircraft, and the largest benefits of initial investigation are likely to be yielded from this platform. The database produced would also be utilized by flight dynamics engineers as a means to further develop and investigate vehicle flight characteristics as mission tactics evolve through the years ahead. This paper will describe ongoing efforts by the U.S. Navy to develop a methodology for simulation and training for large commercial-derived transport aircraft at unusual attitudes, typically experienced during an aircraft upset. This methodology will be applied to a representative Navy aircraft (P-8A) and utilized to develop a robust simulation that should accurately represent aircraft response in these extremes. Simulation capabilities would then extend to flight dynamics analysis and simulation, as well as potential training applications. Recent evaluations of integrated academic, ground-based simulation, and in-flight upset training will be described along with important lessons learned, specific to military requirements.

Description: A flight experiment was conducted to investigate the pressure distribution, local flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.0. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. The tests used a F-15B testbed aircraft with a bottom centerline mounted test fixture. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating for future laminar flow flight tests employing infrared thermography. Boundary-layer transition was captured using an onboard infrared imaging system. The infrared imagery was captured in both analog and digital formats. Surface pressures were measured with electronically scanned pressure modules connected to 60 surface-mounted pressure orifices. The local flow field was measured with five 5-hole conical probes mounted near the leading edge of the test fixture. Flow field measurements revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration. The infrared imaging system was able to capture shock wave impingement on the surface of the flat plate in addition to indicating laminar-to-turbulent boundary-layer transition.

Description: Dr. Edwards'presentation provides an overview of aviation's economic impact in the U.S. including aviation's impact on environment and energy. The presentation discusses NASA's contributions to the advancement of commercial aircraft design highlighting the technology drivers and recent technology advancements for addressing community noise, energy efficiency and emissions. The presentation concludes with a preview of some of NASA's integrated systems solutions, such as novel aircraft concepts and advancements in propulsion that will enable the future of more environmentally compatible aviation.

Description: The Multidisciplinary Design Analysis and Optimization (MDAO) community has developed a multitude of algorithms and techniques, called architectures, for performing optimizations on complex engineering systems which involve coupling between multiple discipline analyses. These architectures seek to efficiently handle optimizations with computationally expensive analyses including multiple disciplines. We propose a new testing procedure that can provide a quantitative and qualitative means of comparison among architectures. The proposed test procedure is implemented within the open source framework, OpenMDAO, and comparative results are presented for five well-known architectures: MDF, IDF, CO, BLISS, and BLISS-2000. We also demonstrate how using open source soft- ware development methods can allow the MDAO community to submit new problems and architectures to keep the test suite relevant.

Description: This is a slide presentation regarding the Design evolution of the SR-71.

Description: This paper will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. Preliminary results show that the PSP agrees both qualitatively and quantitatively with both the expected results as well as with the pressure taps. Several areas of improvement have been indentified and are currently being developed.

Description: This paper presents a novel approach to design of the supersonic aircraft outer mold line (OML) by optimizing the A-weighted loudness of sonic boom signature predicted on the ground. The optimization process uses the sensitivity information obtained by coupling the discrete adjoint formulations for the augmented Burgers Equation and Computational Fluid Dynamics (CFD) equations. This coupled formulation links the loudness of the ground boom signature to the aircraft geometry thus allowing efficient shape optimization for the purpose of minimizing the impact of loudness. The accuracy of the adjoint-based sensitivities is verified against sensitivities obtained using an independent complex-variable approach. The adjoint based optimization methodology is applied to a configuration previously optimized using alternative state of the art optimization methods and produces additional loudness reduction. The results of the optimizations are reported and discussed.

Description: No abstract available

Description: A first-of-its-kind demonstration of the use of localized, non-harmonic active flap motions, for suppressing low frequency, in-plane rotor noise, is reported in this paper. Operational feasibility is verified via testing of the full-scale AATD/Sikorsky/UTRC active flap demonstration rotor in the NFAC's 40- by 80-Foot anechoic wind tunnel. Effectiveness of using localized, non-harmonic active flap motions are compared to conventional four-per-rev harmonic flap motions, and also active flap motions derived from closed-loop acoustics implementations. All three approaches resulted in approximately the same noise reductions over an in-plane three-by-three microphone array installed forward and near in-plane of the rotor in the nearfield. It is also reported that using an active flap in this localized, non-harmonic manner, resulted in no more that 2% rotor performance penalty, but had the tendency to incur higher hub vibration levels.

Description: The conceptual design of a short-haul civil tiltrotor aircraft is presented. The concept vehicle is designed for runway-independent operations to increase the capacity of the National Airspace System without the need for increased infrastructure. This necessitates a vehicle that is capable of integrating with conventional air traffic without interfering with established flightpaths. The NASA Design and Analysis of Rotorcraft software was used to size the concept vehicle based on the mission requirements of this market. The final configuration was selected based upon performance metrics such as acquisition and maintenance costs, fuel fraction, empty weight, and required engine power. The concept presented herein has a proposed initial operating capability date of 2035, and is intended to integrate with conventional air traffic as well as proposed future air transportation concepts.

Description: Investigate the feasibility of using piezoelectric sensors/actuators for subscale GEnx composite fan blade damping at target modes.

Description: No abstract available

Description: No abstract available

Description: A lot of composite materials are used for the modern aircraft to reduce its weight. Aircraft aeroservoelastic models are typically characterized by significant levels of model parameter uncertainty due to composite manufacturing process. Small modeling errors in the finite element model will eventually induce errors in the structural flexibility and mass, thus propagating into unpredictable errors in the unsteady aerodynamics and the control law design. One of the primary objectives of X-56A aircraft is the flight demonstration of active flutter suppression, and therefore in this study, the identification of the primary and secondary modes is based on the flutter analysis of X-56A aircraft. It should be noted that for all three Mach number cases rigid body modes and mode numbers seven and nine are participated 89.1 92.4 % of the first flutter mode. Modal participation of the rigid body mode and mode numbers seven and nine for the second flutter mode are 94.6 96.4%. Rigid body mode and the first two anti-symmetric modes, eighth and tenth modes, are participated 93.2 94.6% of the third flutter mode. Therefore, rigid body modes and the first four flexible modes of X-56A aircraft are the primary modes during the model tuning procedure. The ground vibration test-validated structural dynamic finite element model of the X-56A aircraft is to obtain in this study. The structural dynamics finite element model of X-56A aircraft is improved using the parallelized big-bang big-crunch algorithm together with a hybrid optimization technique.

Description: No abstract available

Description: No abstract available