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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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61 records · Page 4
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NASA NTRS2001

Summary Report of the Orbital X-34 Wing Static Aeroelastic Study

This report documents the results of a computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid aerodynamic characteristics taking into account the wing structural flexibility. This was a joint exercise between LaRC and SDRC of California. SDRC modeled the structural details of the wing, and provided the structural deformation for a given pressure distribution on its surfaces. This study was done for a Mach number of 1.35 and an angle of attack of 9 deg.; the freestream dynamic pressure was assumed to be 607 lb/sq ft. Only the wing and the body were simulated in the CFD computations. Two wing configurations were examined. The first had the elevons in the undeflected position and the second had the elevons deflected 20 deg. up. The results indicated that with elevon undeflected, the wing twists by about 1.5 deg. resulting in a reduction in the angle of attack at the wing tip to by 1.5 deg. The maximum vertical deflection of the wing is about 3.71 inches at the wing tip. For the wing with the undeflected elevons, the effect of this wing deformation is to reduce the normal force coefficient (C(sub N)) by 0.012 and introduce a noise up pitching moment coefficient (C(sub m)) of 0.042.

Prabhn, Ramadas K.↗
NASA NTRS2004

Characterization of Unsteady Flow Structures Near Leading-Edge Slat: PIV Measurements - Part 1

A comprehensive computational and experimental study has been performed at the NASA Langley Research Center as part of the Quiet Aircraft Technology (QAT) Program to investigate the unsteady flow near a leading-edge slat of a two-dimensional, high-lift system. This paper focuses on the experimental effort conducted in the NASA Langley Basic Aerodynamics Research Tunnel (BART) where Particle Image Velocimetry (PIV) data was acquired in the slat cove and at the slat trailing edge of a three-element, high-lift model at 4, 6, and 8 degrees angle of attack and a freestream Mach Number of 0.17. Instantaneous velocities obtained from PIV images are used to obtain mean and fluctuating components of velocity and vorticity. The data show the recirculation in the cove, reattachment of the shear layer on the slat lower surface, and discrete vortical structures within the shear layer emanating from the slat cusp and slat trailing edge. Detailed measurements are used to examine the shear layer formation at the slat cusp, vortex shedding at the slat trailing edge, and convection of vortical structures through the slat gap. Selected results are discussed and compared with unsteady, Reynolds-Averaged Navier-Stokes (URANS) computations for the same configuration in a companion paper by Khorrami, Choudhari, and Jenkins (2004). The experimental dataset provides essential flow-field information for the validation of near-field inputs to noise prediction tools.

Jenkins, Luther N.↗
NASA NTRS1992

Acoustic characteristics and dynamic structural loading of an ASTOVL aircraft in hover

Measurements of surface dynamic loading and freestream acoustics were made for an ASTOVL model in hover, to quantify the effects of elevated temperature on the acoustic field and surface loading. Data were acquired for a many combinations of operating parameters: model height above the ground plane, nozzle pressure ratio, and jet exit stagnation temperature. For many conditions, strong tones were observed, with amplitudes up to 150 dB. The frequencies of the strongest tones were well predicted by a model assuming feedback between the nozzle exit and the ground plane. The model also accounts for many of the variations in frequency observed with changes in model height, nozzle pressure ratio, and jet temperature. Broadband sound pressure levels up to 170 dB were also recorded. The maximum levels occurred at approximately 3 equivalent jet diameters above the ground plane. For the majority of the cases, the increase in noise due to temperature was less than expected based on free jet correlations.

Mitchell, L. K.↗
NASA NTRS2001

Qualitative Assessment of the Acoustic Disturbance Environment in the NASA LaRC 20-Inch MACH 6 Wind Tunnel

An experimental investigation was conducted on a 5-degree-half-angle cone with a flare in a conventional Mach 6 wind tunnel to examine the effect of facility noise on boundary layer transition. The effect of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a quiet hypersonic wind tunnel using a common test model. In the present conventional hypersonic tunnel experiment, adiabatic wall temperatures were measured and heat transfer distributions were inferred on the cone flare model at zero degree angle of attack over a range of length Reynolds numbers (2 x 10(exp 6) to 10 x 10(exp 6)) which resulted in laminar and turbulent flow. Wall-to-total temperature ratio for the transient heating measurements and the adiabatic wall temperature measurements were 0.69 and 0.86, respectively. The cone flare nosetip radius was varied from 0.0001 to 0.125-inch to examine the effects of bluntness on transition onset. At comparable freestream conditions the transition onset Reynolds number obtained on the cone flare model in the conventional "noisy" tunnel was approximately 25% lower than that measured in the low disturbance tunnel.

Horvath, Thomas J.↗
NASA NTRS2003

Flap-Edge Blowing Experiments

This Appendix documents the salient results from an effort to mitigate the so-called flap-edge noise generated at the split between a flap edge that is deployed and the undeployed flap. Utilizing a Coanda surface installed at the flap edge, steady blowing was used in an attempt to diminish the vortex strength resulting from the uneven lift distribution. The strength of this lifting vortex was augmented by steady blowing over the deployed flap. The test article for this study was the same 2D airfoil used in the steady blowing program reported earlier (also used in pulsed blowing tests, see Appendix G), however its trailing edge geometry was modified. An exact duplicate of the airfoil shape was made out of fiberglass with no flap, and in the clean configuration. It was attached to the existing airfoil to make an airfoil that has half of its flap deployed and half un-deployed. Figure 1 shows a schematic of the planform showing the two areas where steady blowing was introduced. The flap-edge blowing or the auxiliary blowing was in the direction normal to the freestream velocity vector. Slot heights for the blowing chambers were on the order of 0.0 14 inches.

Gaeta, R. J.↗
NASA NTRS2002

Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

An experimental investigation was conducted on a 5-degree half-angle cone and a 5-degree half-angle flared cone in a conventional Mach 6 wind tunnel to examine the effects of facility noise on boundary layer transition. The influence of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 low disturbance quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a low disturbance wind tunnel using a common test model and transition detection technique. In the present conventional hypersonic tunnel experiment, separate measurements of heat transfer and adiabatic wall temperatures were obtained on the conical models at small angles of attack over a range of Reynolds numbers, which resulted in laminar, transitional, and turbulent flow. Smooth model turbulent heating distributions are compared to that obtained with transition forced via discrete surface roughness. The model nosetip radius was varied to examine the effects of bluntness on transition onset. Despite wall to total temperature differences between the transient heating measurements and the adiabatic wall temperature measurement, the two methods for determining sharp cone transition onset generally yielded equivalent locations. In the 'noisy' mode of the hypersonic low disturbance tunnel, transition onset occurred earlier than that measured in the conventional hypersonic tunnel, suggesting higher levels of freestream acoustic radiation relative to the conventional tunnel. At comparable freestream conditions, the transition onset Reynolds number under low disturbance conditions was a factor of 1.3 greater than that measured on flared cone in the LaRC conventional hypersonic tunnel and a factor of 1.6 greater that the flared cone run in the low disturbance tunnel run 'noisy'. Navier-Stokes mean flow computations and linear stability analysis were conducted to assess the experimental results and have indicated N factors associated with sharp flared cone transition onset to be approximately a factor of 2 lower than that inferred from the corresponding low disturbance tunnel measurements.

Horvath, Thomas J.↗
NASA NTRS2009

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

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

Felder, James L.↗