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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Study of casing treatment stall margin improvement phenomena

The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot, and blade angle slot treatments were tested. These yielded, for reduction in stalling flow and loss in peak efficiency, 5.8% and 0 points, 15.3% and 2.0 points, and 15.0% and 1.2 points, respectively. These values are consistent with other experience. The favorable stalling flow situations correlated well with observations of higher-then-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annular wall pressure gradients, especially in the 50-75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations. Rotor blade wakes may be somewhat thinner in the presence of good treatments, particularly under operating conditions close to the baseline stall.

Prince, D. C., Jr.↗
NASA NTRS1975

A study of casing treatment stall margin improvement phenomena

The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot and blade angle slot treatments were tested at low speeds. With the circumferential groove treatment the stalling flow was reduced 5.8% at negligible efficiency sacrifice. The axial-skewed slot treatment improved the stalling flow by 15.3%; 1.8 points in peak efficiency were sacrificed. The blade angle slot treatment improved the stalling flow by 15.0%; 1.4 points in peak efficiency were sacrificed. The favorable stalling flow situations correlated well with observations of higher-than-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annulus wall pressure gradients, especially in the 50 to 75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations.

Prince, D. C., Jr.↗
NASA NTRS1986

The flowfield structure of the 3-D shock wave - Boundary layer interaction generated by a 20 deg sharp fin at Mach 3

The three-dimensional oblique shock wave-turbulent boundary layer interaction generated by a sharp fin attached to a flat plate is investigated experimentally and theoretically for Mach 3 and Reynolds number 9 x 10 to the 5th using two different models. Both models employ the three-dimensional compressible Navier-Stokes equations in mass-averaged variables; one model utilizes the algebraic turbulent eddy viscosity model of Baldwin and Lomax (1978), while the other model employs the two-equation turbulence model of Jones and Launder (1972) coupled with the wall function model of Viegas and Rubesin (1985). The computed surface pressure, surface streamlines, pitot pressure, and yaw angle profiles are found to be in good agreement with experimental data. The three-dimensional velocity fields computed by both models are in close agreement, although the eddy viscosity profiles differ significantly within the three-dimensional interaction. This result indicates that the overall structure of this three-dimensional sharp fin interaction is insensitive to the turbulence model.

Knight, D. D.↗
NASA NTRS2004

Application of Excitation from Multiple Locations on a Simplified High-Lift System

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

Melton, LaTunia Pack↗
NASA NTRSDate not supplied

Validation of the Hypersolve CFD Solver for Entry Descent and Landing Applications

The functional equivalence of the HyperSolve unstructured edge-based, finite-volume computational fluid dynamics code to the Langley Aerothermodynamic Upwind Relaxation Algorithm multiblock structured grid code is documented for applications of interest to the Entry, Descent, and Landing community. A suite of cases using a range of thermochemical gas models on relevant vehicle configurations were analyzed with both codes and the results compared. A tolerance of ±4% difference in surface pressure and surface heat flux from a benchmark LAURA solution was used as the criterion for functional equivalence, and comparisons of flowfield quantities are also included to verify that the thermochemical nonequilibrium capabilities in HyperSolve match those of the LAURA code. The functional equivalence of the HyperSolve unstructured edge-based finite-volume computational fluid dynamics (CFD) code to the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) multiblock structured grid code is documented for applications of interest to the Entry, Descent, and Landing (EDL) community. A suite of cases using a range of thermochemical gas models on EDL-relevant vehicle configurations were analyzed with both codes and the results compared. A tolerance of ±4% difference in surface pressure and surface heat flux from a benchmark LAURA solution was used as the criterion for functional equivalence, and comparisons of flow field quantities are also included to verify that the thermochemical nonequilibrium capabilities in HyperSolve match those of the LAURA code. In general, HyperSolve predictions for surface pressure and surface heat flux are in close agreement with those predicted by LAURA.

hypersolve↗
NASA NTRSDate not supplied

Validation of the HyperSolve CFD Solver for Entry Descent and Landing Applications

The functional equivalence of the HyperSolve unstructured edge-based, finite-volume computational fluid dynamics code to the Langley Aerothermodynamic Upwind Relaxation Algorithm multiblock structured grid code is documented for applications of interest to the Entry, Descent, and Landing community. A suite of cases using a range of thermochemical gas models on relevant vehicle configurations were analyzed with both codes and the results compared. A tolerance of ±4% difference in surface pressure and surface heat flux from a benchmark LAURA solution was used as the criterion for functional equivalence, and comparisons of flowfield quantities are also included to verify that the thermochemical nonequilibrium capabilities in HyperSolve match those of the LAURA code. The functional equivalence of the HyperSolve unstructured edge-based finite-volume computational fluid dynamics (CFD) code to the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) multiblock structured grid code is documented for applications of interest to the Entry, Descent, and Landing (EDL) community. A suite of cases using a range of thermochemical gas models on EDL-relevant vehicle configurations were analyzed with both codes and the results compared. A tolerance of ±4% difference in surface pressure and surface heat flux from a benchmark LAURA solution was used as the criterion for functional equivalence, and comparisons of flow field quantities are also included to verify that the thermochemical nonequilibrium capabilities in HyperSolve match those of the LAURA code. In general, HyperSolve predictions for surface pressure and surface heat flux are in close agreement with those predicted by LAURA.

hypersolve↗
NASA NTRS1990

Experimental and computational surface and flow-field results for an all-body hypersonic aircraft

The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA/Ames 3.5-foot Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pitot-pressure flow-field surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at Ames demonstrate the capabilities of this code.

Lockman, William K.↗
NASA NTRS1992

Flow over an all-body hypersonic aircraft - Experiment and computation

The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA Ames 3.5-ft Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pilot-pressure flowfield surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at NASA Ames demonstrate the capabilities of this code.

William K. Lockman↗
NASA NTRS1990

Experimental aerothermodynamic research of hypersonic aircraft

Wind tunnel tests were conducted to establish a benchmark experimental data base for a genetic hypersonic aircraft vehicle. Comprehensive measurements were made at Mach 7 to give flow visualization, surface pressure, surface convective heat transfer, and flow field Pitot pressure for a delta platform all-body vehicle. The tests were conducted in the NASA/Ames 3.5-Foot Hypersonic Wind Tunnel at Reynolds numbers sufficient to give turbulent flow. Comparisons are made of the experimental results with computational solutions of the flow by an upwind parabolized Navier-Stokes code developed at Ames. Good agreement of experiment with solutions by the code is demonstrated.

Cleary, Joseph W.↗
NASA NTRS1973

Rotation of the upper atmosphere of Venus.

A fundamental aspect of the Venus wind system is discussed that has apparently escaped notice. If Venus has a negligible meridional gradient of surface pressure, as the Venera data suggest, but has a small decrease in the average temperature from the equator to both poles, then the pressure surfaces at high elevations are likely to display an equatorial bulge. The pressure force associated with the bulge can only be balanced by an excess centrifugal force due to a rotation rate that increases with height. The implications of this aspect are investigated.

Leovy, C. B.↗
NASA NTRS1992

Quasiconical flowfield structure of the three-dimensional single fin interaction

A series of conical and three-dimensional computations have been performed for the swept oblique shock wave/turbulent boundary-layer interaction generated by a 20-deg sharp fin at Mach 4 and freestream Reynolds number of 2.18 x 10 exp 5 based on the incoming boundary-layer thickness. The Reynolds-averaged compressible Navier-Stokes equations are employed with turbulence incorporated using the Baldwin-Lomax and Jones-Launder models. The computed results are basically similar for both turbulence models and display general agreement with experimental data for surface pressure and surface flow direction, although underestimating the size of the primary vortex. The computed three-dimensional flowfield displays quasiconical behavior of the surface pressure, surface flow direction, and flowfield contours of static pressure, density, and Mach number over the extent of the computational domain except for an inception region near the fin leading edge. Certain features of the flowfield model are not observed in the computations, namely, a 'normal' shock near the attachment line, transonic shocklets in the expansion region, and secondary separation. The absence of these features in the computation is believed to be indirectly attributable to limitations in the turbulence models.

Knight, Doyle D.↗
NASA NTRS1981

Fluid mechanics mechanisms in the stall process of helicopters

Recent experimental results from airfoils in the Mach number, Reynolds number, or reduced frequency ranges typical of helicopter rotor blades have identified the most influential flow mechanisms in the dynamic stall process. The importance of secondary shed vortices, downstream wake action, and the flow in the separated region is generally acknowledged but poorly understood. By means of surface pressure cross-correlations and flow field measurements in static stall, several new hypotheses have been generated. It is proposed that vortex shedding may be caused by acoustic disturbances propagating forward in the lower (pressure) surface boundary layer, that wake closure is a misnomer, and that the shed vortex leaves a trail of vorticity that forms a turbulent free shear layer. The known dynamic stall flow mechanisms are reviewed and the potential importance of recently proposed and hypothetical flow phenomena with respect to helicopter blade aeroelastic response are assessed.

Young, W. H., Jr.↗
NASA NTRS1990

Unsteady analysis of hot streak migration in a turbine stage

Two- and three-dimensional Navier-Stokes analyses are used to predict unsteady viscous rotor-stator interacting flow in the presence of a combustor hot streak. Predicted results are presented for a two-dimensional three-stator/four-rotor, a two-dimensional one-stator/one-rotor, and a three-dimensional one-stator/one-rotor simulation of hot streak migration through a turbine stage. Comparison of these results with experimental data demonstrates the capability of the three-dimensional procedure to capture most of the flow physics associated with hot streak migration including the effects of combustor hot streaks on turbine rotor surface temperatures. It is noted that blade count ratio has little effect on predicted time-averaged surface pressure and temperature distributions, but a substantial effect on the unsteady flow characteristics. It is shown that high-temperature hot streak fluid accumulates on the pressure surface of the rotor blades, resulting in a high time-averaged surface temperature 'hot spots'.

Dorney, Daniel J.↗
NASA NTRS1991

Numerical prediction of the unsteady flowfield around the F-18 aircraft at large incidence

This paper describes a numerical method capable of solving the steady and unsteady viscous flow around complete aircraft configurations at high angles of attack. This method is used to simulate the external flow around the F-18 aircraft, including deflected control surfaces. The current technique employs a generalized overset zonal grid scheme to decompose the computational space around the aircraft. The grid around various components of the aircraft are created numerically using a three-dimensional hyperbolic grid generation procedure. The Reynolds-averaged Navier-Stokes equations are integrated using a time-accurate, implicit procedure. Results for the turbulent flow around the F-18 aircraft at 30 degrees angle of attack show the details of the flowfield structure, including the unsteadiness created by the vortex burst and the resulting fluctuating airloads exerted on the vertical tail. The computed results agree fairly well with flight data for surface pressure, surface flow pattern, vortex burst location, and the dominant frequency for tail load fluctuations.

Rizk, Yehia M.↗
NASA NTRS1992

Hypersonic aerodynamic/aerothermodynamic testing capabilities at Langley Research Center

Capabilities of the Langley Hypersonic Facilities Complex that consists of nine blowdown-to-vacuum wind tunnels are reviewed. These tunnels complement one another to provide ranges of Mach number from 6 to 20, unit Reynolds number from 0.03 to 40 million per ft, and normal shock density ratio from 2 to 12. Topics under consideration include the basic components of each facility, status, and upgrades recently performed; instrumentation and testing techniques routinely used to measure forces and moments, surface pressures, surface temperature-time histories, flow properties within the shock layer about the model via survey probes, and flow visualization; and data acquisition systems and upgrades performed via a NASA revitalization program.

Miller, C. G., III↗