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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Low speed performance of a supersonic axisymmetric mixed compression inlet with auxiliary inlets

The aerodynamic performance of a representative supersonic cruise inlet was investigated using a fan simulator coupled to the inlet to provide characteristic noise signatures and to pump the inlet flow. Data were obtained at Mach numbers from 0 to 0.2 for the inlet equipped with an auxiliary inlet system that provided 20 to 40 percent of the fan flow. Results show that inlet performance improved when the inlet bleed systems were sealed; when the freestream Mach number was increased; and when the auxiliary inlets were opened. The inlet flow could not be choked by either centerbody translation or by increasing the fan speed when the 40 percent auxiliary inlet was incorporated. Previously announced in STAR as N83-27992

Wasserbauer, J. F.↗
NASA NTRS1986

An evaluation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

The aerodynamic performance of leading-edge flaps on three delta double-delta wing planforms having aspect ratios of 1.75, 2.11, and 2.50, have been investigated experimentally. The wings were mounted on a generic fuselage without an inlet canopy, or a vertical tail. The Mach numbers of the flow over the wings were 1.60, 1.90 and 2.16. A primary set of full-span leading-edge flaps with similar root and tip chords were tested on each wing, and several alternate flap planforms were tested on the aspect ratio 1.75 wings. It is found that all leading edge geometries were effective in reducing drag lifting over the range of wing aspect ratios and Mach numbers tested. Greater flap performance was obtained when primary flaps were applied to the delta planform. In general, the primary flap geometry yielded better performance than the alternative geometries tested. Flow visualization techniques were found to be useful for identifying the beneficial effects of leading-edge flap deflection on flow separation as well as fuselage interference effects. Black and white photographs of the delta and double-delta planforms are provided.

Covell, P. F.↗
NASA NTRS2007

Performance of Advanced Heavy-Lift, High-Speed Rotorcraft Configurations

The aerodynamic performance of rotorcraft designed for heavy-lift and high-speed cruise is examined. Configurations considered include the tiltrotor, the compound helicopter, and the lift-offset rotor. Design conditions are hover and 250-350 knot cruise, at 5k/ISA+20oC (civil) or 4k/95oF (military); with cruise conditions at 4000 or 30,000 ft. The performance was calculated using the comprehensive analysis CAMRAD II, emphasizing rotor optimization and performance, including wing-rotor interference. Aircraft performance was calculated using estimates of the aircraft drag and auxiliary propulsion efficiency. The performance metric is total power, in terms of equivalent aircraft lift-to-drag ratio L/D = WV/P for cruise, and figure of merit for hover.

Johnson, Wayne↗
NASA NTRS1988

Hypersonic aerodynamics at North Carolina State University

Hypersonic Aerodynamics has attracted outstanding graduate students and faculty at NCSU. It has grown to 6 faculty and 32 graduate students for fall, 1988. A significant feature of the program is that students and faculty spend time at government laboratories which currently include NASA Langley Research Center, Naval Surface Warfare Center at Silver Spring, and the Wright Aeronautical Laboratories. Research projects include analysis, computational and experimental aerodynamics. The combined program produces graduates with the background needed to perform aerodynamics investigations of hypersonic aircraft and spacecraft. The research performed helps advance the state of the art as well as assist government laboratories in meeting their goals.

Dejarnette, Fred R.↗
NASA NTRS2003

Effect of Intercycle Ice Accretions on Airfoil Performance

This paper presents the results of an experimental study designed to characterize and evaluate the aerodynamic performance penalties of residual and intercycle ice accretions that result from the cyclic operation of a typical aircraft deicing system. Icing wind tunnel tests were carried out on a 36-inch chord NACA 23012 airfoil section equipped with a pneumatic deicer for several different FAR 25 Appendix C cloud conditions. Results from the icing tests showed that the intercycle ice accretions were much more severe in terms of size and shape than the residual ice accretions. Molds of selected intercycle ice shapes were made and converted to castings that were attached to the leading edge of a 36-inch chord NACA 23012 airfoil model for aerodynamic testing. The aerodynamic testing revealed that the intercycle ice shapes caused a significant performance degradation. Maximum lift coefficients were typically reduced about 60% from 1.8 (clean) to 0.7 (iced) and stall angles were reduced from 17 deg. (clean) to 9 deg. (iced). Changes in the Reynolds number (from 2.0 x 10(exp 6) to 10.5 x 10(exp 6) and Mach number (from 0.10 to 0.28) did not significantly affect the iced-airfoil performance.

Broeren, Andy P.↗
NASA NTRS1972

The NASA/GE quiet engine 'A.'

The National Aeronautics and Space Administration and the General Electric Company are jointly developing two low noise propulsion technology demonstrators (Engine 'A' and 'C' incorporating respectively a low tip speed and high tip speed fan) under the Experimental Quiet Engine Program. The acoustic and aerodynamic performance characteristics of Engine 'C' which is currently on test will be reported at a later date. The present paper describes the design, component development and engine evaluation of the 'A' demonstrator. The aerodynamic and acoustic performance obtained on the fan component are discussed. The demonstrator engine acoustic test program is outlined. The effect of inlet geometry and duct acoustic treatment are presented. The static acoustic data are extrapolated to flight and compared with the FAA noise regulations.

Benzakein, M. J.↗
NASA NTRS2019

Establishing Best Practices for X-57 Maxwell CFD Database Generation

The X-57 Maxwell is NASA’s latest electric airplane concept that has been simulated for aerodynamic performance using the structured overset and unstructured grid solvers within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework as well as the unstructured polyhedral grid solver in Star-CCM+ for code-to-code comparison. In order to validate the predictions, comparisons were made between the CFD solutions and experimental data collected in the 12-foot Low-Speed Wind Tunnel at NASA Langley Research Center. The simulations are in preparation for the development of a comprehensive aerodynamic database which will assess aircraft performance at a variety of conditions. The findings from these simulations will establish the best practices for mesh resolution, numerical discretization, and turbulence modeling to be used for this database. Preliminary database results have shown that best-practices learned from the initial validation simulations will potentially reduce error in X-57 aerodynamic loads and moments relative to experiment by up to 14%.

Duensing, Jared C.↗
NASA NTRS2013

Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle

On August 5, 2012, the Mars Science Laboratory entry vehicle successfully entered Mars atmosphere, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the Entry Descent and Landing sequence. This paper compares the aerodynamics of the entry capsule extracted from onboard flight data, including Inertial Measurement Unit (IMU) accelerometer and rate gyro information, and heatshield surface pressure measurements. From the onboard data, static force and moment data has been extracted. This data is compared to preflight predictions. The information collected by MSL represents the most complete set of information collected during Mars entry to date. It allows the separation of aerodynamic performance from atmospheric conditions. The comparisons show the MSL aerodynamic characteristics have been identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. This data will help revise aerodynamic databases for future missions and will guide computational fluid dynamics (CFD) development to improved prediction codes.

Schoenenberger, Mark↗
NASA NTRS2013

Testing of the Trim Tab Parametric Model in NASA Langley's Unitary Plan Wind Tunnel

In support of NASA's Entry, Descent, and Landing technology development efforts, testing of Langley's Trim Tab Parametric Models was conducted in Test Section 2 of NASA Langley's Unitary Plan Wind Tunnel. The objectives of these tests were to generate quantitative aerodynamic data and qualitative surface pressure data for experimental and computational validation and aerodynamic database development. Six component force-and-moment data were measured on 38 unique, blunt body trim tab configurations at Mach numbers of 2.5, 3.5, and 4.5, angles of attack from -4deg to +20deg, and angles of sideslip from 0deg to +8deg. Configuration parameters investigated in this study were forebody shape, tab area, tab cant angle, and tab aspect ratio. Pressure Sensitive Paint was used to provide qualitative surface pressure mapping for a subset of these flow and configuration variables. Over the range of parameters tested, the effects of varying tab area and tab cant angle were found to be much more significant than varying tab aspect ratio relative to key aerodynamic performance requirements. Qualitative surface pressure data supported the integrated aerodynamic data and provided information to aid in future analyses of localized phenomena for trim tab configurations.

Murphy, Kelly J.↗
NASA NTRS2015

Optimized Off-Design Performance of Flexible Wings with Continuous Trailing-Edge Flaps

This work assesses the potential aerodynamic performance benefits of a variable-camber, continuous-trailing-edge flap system on a generic transport aircraft at off-design conditions. A process to optimize transport wings while addressing static aeroelastic effects is presented. To establish a proper baseline, a transport wing is first aerodynamically optimized at a mid-cruise flight condition using an inviscid, aeroelastic analysis tool. The optimized wing is then analyzed at off-design cruise conditions. The optimization is repeated at these off-design conditions to determine how much performance is lost by the wing optimized solely for the mid-cruise condition. The full-span flap system is then adapted to improve performance of the mid-cruise-optimized wing at these off-design conditions. The measured improvement is quantified by a comparison with wings designed specifically for the off-design conditions. To evaluate the effects of aeroelasticity on the effectiveness of the flap system, this entire process is performed on both a conventionally stiff wing and a modern, more flexible wing. The results indicate that the flap system allows for recovery of near-optimal performance throughout cruise. The flap system is found to be advantageous even for modern wings with increased flexibility.

Rodriguez, David L.↗
NASA NTRS2001

RLV Turbine Performance Optimization

A task was developed at NASA/Marshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. There are four major objectives of this task: 1) to develop, enhance, and integrate advanced turbine aerodynamic design and analysis tools; 2) to develop the methodology for application of the analytical techniques; 3) to demonstrate the benefits of the advanced turbine design procedure through its application to a relevant turbine design point; and 4) to verify the optimized design and analysis with testing. Final results of the preliminary design and the results of the two-dimensional (2D) detailed design of the first-stage vane of a supersonic turbine suitable for a reusable launch vehicle (R-LV) are presented. Analytical techniques for obtaining the results are also discussed.

Griffin, Lisa W.↗
NASA NTRS1976

Performance of a low-pressure fan stage with reverse flow

The reverse flow aerodynamic performance of a 51-centimeter-diameter fan stage is presented. The stage was tested with the variable pitch rotor blades set through feather at -75 deg, -80 deg, and -85 deg from design setting angle. Of the three tested the stage with the rotor blades set at -75 deg exhibited the highest pressure ratio and highest flow. For all three configurations, there was little or no flow in the inner third of the exit passage due to the rotor blade being almost perpendicular to the axial direction in the hub region.

Moore, R. D.↗
NASA NTRS1978

Performance of a transonic fan stage designed for a low meridional velocity ratio

The aerodynamic performance and design parameters of a transonic fan stage are presented. The fan stage was designed for a meridional velocity ratio of 0.8 across the tip of the stage, a pressure ratio of 1.57, a flow of 29.5 kilograms per second, and a tip speed of 426 meters per second. Radial surveys were obtained over the stable operating range from 50 to 100 percent of design speed. The measured, peak efficiency (0.81) of the stage occurred at a pressure ratio of 1.58 and a flow of 28.7 kilograms per second.

Moore, R. D.↗
NASA NTRS1980

Performance of annular prediffuser-combuster systems

Results of an experimental investigation of the aerodynamic performance of several annular prediffuser-combustor systems are presented. Three curved wall, dump prediffusers of different length, area ratio, and turning angle were tested with and without a simulated combustor located downstream of the prediffuser. Performance was significantly influenced by the presence of the combustor. Pressure recovery and flow losses were determined as a function of prediffuser inlet velocity profile, flow extraction at the prediffuser inlet, axial and radial location of the combustor front end, and distribution of the flow in the combustor. Axial location of the combustor was found to be the most significant parameter influencing system performance.

Wagner, W. B.↗
NASA NTRS1991

A comparison of the analytical and experimental performance of the solid version of a cooled radial turbine

An evaluation of the aerodynamic performance of the solid version of an Allison-designed cooled radial turbine was conducted at NASA Lewis' Warm Turbine Test Facility. The resulting pressure and temperature measurements are used to calculate vane, rotor, and overall stage performance. These performance results are then compared to the analytical results obtained by using NASA's MTSB (MERIDL-TSONIC-BLAYER) code.

Tirres, Lizet↗
NASA NTRS1983

Coolant passage heat transfer with rotation. A progress report on the computational aspects

Turbine airfoils are subjected to increasingly higher heat loads which escalate the cooling requirements in order to satisfy life goals for the component materials. If turbine efficiency is to be maintained, however, cooling requirements should be as low as possible. To keep the quantity of cooling air bounded, a more efficient internal cooling scheme must be developed. One approach is to employ airfoils with multipass cooling passages that contain devices to augment internal heat transfer while limiting pressure drop. Design experience with multipass cooling passage airfoils has shown that a surplus of cooling air must be provided as a margin of safety. This increased cooling air leads to a performance penalty. Reliable methods for predicting the internal thermal and aerodynamic performance of multipass cooling passage airfoils would reduce or eliminate the need for the safety margin of surplus cooling air. The objective of the program is to develop and verify improved analytical methods that will form the basis for design technology which will result in efficient turbine components with improved durability without sacrificing performance. The objective will be met by: (1) establishing a comprehensive experimental data base that can form the basis of an empirical design system; (2) developing computational fluid dynamic techniques; and (3) analyzing the information in the data base with both phenomenological modeling and mathematical modeling to derive a suitable design and analysis procedure.

Aceto, L. D.↗
NASA NTRS1989

Rotor blade aerodynamic design

Aerodynamic performance aspects of rotor blade design are presented. Design considerations, aerodynamic constraints and design variables are described.

Walsh, Joanne L.↗