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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Shock Generation and Control Using DBD Plasma Actuators

This report is the final report of a NASA Phase I SBIR contract, with some revisions to remove company proprietary data. The Shock Boundary Layer Interaction (SBLI) phenomena in a supersonic inlet involve mutual interaction of oblique shocks with boundary layers, forcing the boundary layer to separate from the inlet wall. To improve the inlet efficiency, it is desired to prevent or delay shock-induced boundary layer separation. In this effort, Innovative Technology Applications Company (ITAC), LLC and the University of Notre Dame (UND) jointly investigated the use of dielectric-barrier-discharge (DBD) plasma actuators for control of SBLI in a supersonic inlet. The research investigated the potential for DBD plasma actuators to suppress flow separation caused by a shock in a turbulent boundary layer. The research involved both numerical and experimental investigations of plasma flow control for a few different SBLI configurations: (a) a 12 wedge flow test case at Mach 1.5 (numerical and experimental), (b) an impinging shock test case at Mach 1.5 using an airfoil as a shock generator (numerical and experimental), and (c) a Mach 2.0 nozzle flow case in a simulated 15 X 15 cm wind tunnel with a shock generator (numerical). Numerical studies were performed for all three test cases to examine the feasibility of plasma flow control concepts. These results were used to guide the wind tunnel experiments conducted on the Mach 1.5 12 degree wedge flow (case a) and the Mach 1.5 impinging shock test case (case b) which were at similar flow conditions as the corresponding numerical studies to obtain experimental evidence of plasma control effects for SBLI control. The experiments also generated data that were used in validating the numerical studies for the baseline cases (without plasma actuators). The experiments were conducted in a Mach 1.5 test section in the University of Notre Dame Hessert Laboratory. The simulation results from cases a and b indicated that multiple spanwise actuators in series and at a voltage of 75 kVp-p could fully suppress the flow separation downstream of the shock. The simulation results from case c showed that the streamwise plasma actuators are highly effective in creating pairs of counter-rotating vortices, much like the mechanical vortex generators, and could also potentially have beneficial effects for SBLI control. However, to achieve these effects, the positioning and the quantity of the DBD actuators used must be optimized. The wind tunnel experiments mapped the baseline flow with good agreement to the numerical simulations. The experimental results were conducted with spanwise actuators for cases a and b, but were limited by the inability to generate a sufficiently high voltage due to arcing in the wind-tunnel test-section. The static pressure in the tunnel was lower than the static pressure in an inlet at flight conditions, promoting arching and degrading the actuator performance.

Patel, Mehul P.↗
NASA NTRS2009

Active Fail-Safe Micro-Array Flow Control for Advanced Embedded Propulsion Systems

The primary objective of this research effort was to develop and analytically demonstrate enhanced first generation active "fail-safe" hybrid flow-control techniques to simultaneously manage the boundary layer on the vehicle fore-body and to control the secondary flow generated within modern serpentine or embedded inlet S-duct configurations. The enhanced first-generation technique focused on both micro-vanes and micro-ramps highly-integrated with micro -jets to provide nonlinear augmentation for the "strength' or effectiveness of highly-integrated flow control systems. The study focused on the micro -jet mass flow ratio (Wjet/Waip) range from 0.10 to 0.30 percent and jet total pressure ratios (Pjet/Po) from 1.0 to 3.0. The engine bleed airflow range under study represents about a 10 fold decrease in micro -jet airflow than previously required. Therefore, by pre-conditioning, or injecting a very small amount of high-pressure jet flow into the vortex generated by the micro-vane and/or micro-ramp, active flow control is achieved and substantial augmentation of the controlling flow is realized.

Anderson, Bernhard H.↗
NASA NTRS1993

Underwing compression vortex attenuation device

A vortex attenuation device is presented which dissipates a lift-induced vortex generated by a lifting aircraft wing. The device consists of a positive pressure gradient producing means in the form of a compression panel attached to the lower surface of the wing and facing perpendicular to the airflow across the wing. The panel is located between the midpoint of the local wing cord and the trailing edge in the chord-wise direction and at a point which is approximately 55 percent of the wing span as measured from the fuselage center line in the spanwise direction. When deployed in flight, this panel produces a positive pressure gradient aligned with the final roll-up of the total vortex system which interrupts the axial flow in the vortex core and causes the vortex to collapse.

Patterson, James C., Jr.↗
NASA NTRS1987

Lift distributions for a 3-dimensional steady blade-vortex interaction

The interaction of a horizontally mounted V23010-1.58 semispan airfoil (simulating a helicopter rotor blade) with a tip vortex shed by a vertically mounted upstream vortex-generating wing (VGW) is investigated experimentally at 60 m/s (dynamic pressure 2.2 kPa, Reynolds number 850,000, and Mach number 0.17) in the NASA Ames 7 x 10-ft wind tunnel. The velocity field near the blade is determined using a three-dimensional zoom LDV; the spanwise lift distribution is measured by strain gages; and the results are compared with the predictions of the panel computer code VSAERO (Maskew, 1982) in graphs. Features noted include localized loss of lift due to the presence of residual VGW wake, loss of lift far inboard on the blade (indicating the large domain of VGW vorticity), little change in total lift with variations in vortex strength, and good agreement between VSAERO and experiment in overall lift distribution but not in all geometric variations).

Dunagan, Stephen E.↗
NASA NTRS2014

3D Visualization of Cooperative Trajectories

Aerodynamicists and biologists have long recognized the benefits of formation flight. When birds or aircraft fly in the upwash region of the vortex generated by leaders in a formation, induced drag is reduced for the trail bird or aircraft, and efficiency improves. The major consequence of this is that fuel consumption can be greatly reduced. When two aircraft are separated by a large enough longitudinal distance, the aircraft are said to be flying in a cooperative trajectory. A simulation has been developed to model autonomous cooperative trajectories of aircraft; however it does not provide any 3D representation of the multi-body system dynamics. The topic of this research is the development of an accurate visualization of the multi-body system observable in a 3D environment. This visualization includes two aircraft (lead and trail), a landscape for a static reference, and simplified models of the vortex dynamics and trajectories at several locations between the aircraft.

aerodynamic drag↗
NASA NTRS1993

Effect of tabs on the flow and noise field of an axisymmetric jet

The effect of vortex generators, in the form of small tabs projecting normally into the flow at the nozzle exit, on the characteristics of an axisymmetric jet is investigated experimentally over the jet Mach number range of 0.3-1.81. The tabs eliminate screech noise from supersonic jets and alter the shock structure drastically. They distort the jet cross section and increase the jet spread rate significantly. The distortion produced is essentially the same at subsonic and underexpanded supersonic conditions. Thus, the underlying mechanism must be independent of compressibility effects. A tab with a height as small as 2 percent of the jet diameter, but larger than the efflux boundary-layer thickness, is found to produce a significant effect. Flow visualization reveals that each tab introduces an 'indentation' into the high speed side of the shear layer via the action of streamwise vortices. These vortices are inferred to be of the 'trailing vortex' type rather than of the 'necklace vortex' type. It is apparent that a substantial pressure differential must exist between the upstream and the downstream sides of the tab to effectively produce these trailing vortices. This explains why the tabs are ineffective in the overexpanded flow, as in that case an adverse pressure gradient exists near the nozzle exit which reduces the pressure differential produced by the tab.

Samimy, M.↗
NASA NTRS2006

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a miss alignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery that was 0.01 lower than the experiment as was seen in the baseline. Comparison of the flow features the jet case revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Allan Brian G.↗
NASA NTRS1968

Trailing vortex hazard.

Light aircraft hazards due to trailing vortex generated by heavy transport aircraft, suggesting specific procedures for takeoff, landing and flight phase

Mcgowan, W. A.↗
NASA NTRS1969

Systems and Controls Laboratory

Jet spreading in fluid amplifiers, thermal switching of jet flows, and nozzle thrust modulation by vortex generation

Source record↗
NASA NTRS1972

Performance of a bicone inlet designed for Mach 2.5 with internal distributed compression and 40 percent internal contraction

The inlet was designed to have the minimum internal contraction consistent with high total-pressure recovery and low cowl drag. Without a bypass system, the peak pressure recoveries increased from 0.890 to 0.936 when the supercritical bleed mass flow ratio was varied from 0.035 to 0.060. With an operating bypass system and installed centerbody vortex generators, a slight increase in peak pressure recovery was obtained. The values of steady-state distortion and dynamic distortion were below 0.10 and 0.02, respectively, near critical operation. Simulation of a turbofan engine with concentric pipes showed no effect on compressor face flow profiles with varying bypass flow ratio.

Wasserbauer, J. F.↗
NASA NTRS1973

High loading, low speed fan study, 5

A low speed, low noise, single stage fan was designed and tested. Design pressure ratio was 1.5 at a rotor tip speed of 1000 ft/sec. No inlet guide vane was used, the rotor stator was spaced and the number of rotor and stator airfoils was selected for low noise. Tests were conducted with uniform and distorted inlet flows. Stall margin of the initial design was too low for practical application. Airfoil slots and boundary layer and endwall devices did not improve stall margin sufficiently. A redesigned stator with reduced loadings increased stall margin, giving a fan efficiency of 0.883, 15% stall margin, and a 1.474 pressure radio at a specific flow of 41.7 lb/sec sq ft. Casing treatment over rotor tips improved stall margin with distorted inlet flow; vortex generators did not. Blade passing frequency noise increased with rotor relative Mach number. No supersonic fan noise was measured below 105% of design speed. Slotting airfoils, casing treatments, and a reduction of the ratio (number-stators/number-rotors) from (2n + 16) to (2n + 2) had no significant effects on noise.

Keenan, M. J.↗
NASA NTRS1973

Experimental investigation of a large-scale, two-dimensional, mixed-compression inlet system: Internal performance and drag at transonic conditions, free stream Mach equals 0.6 to 1.28

A large scale, variable-geometry inlet system with a design Mach number of 3.0 was tested at Mach numbers from 0.6 to 1.28. Variable features for off-design operation are an adjustable-height ramp system and a translating cowl. Experimental results are presented for transonic ramp and cowl positions showing the effect of throat boundary layer bleed and vortex generators on engine-face performance. Detailed pressure and force-balance data are used to evaluate transonic drag characteristics.

Wong, N. D.↗
NASA NTRS1973

A 727 airplane center duct inlet low speed performance confirmation model test for refanned JT8D engines, phase 2

The results from testing of a 0.3 scale model center duct inlet (S duct) for the Pratt and Whitney Aircraft JT8D-100 engines are presented. The objective of this test was to demonstrate that the required airflow of the JT8D-100 engine (480 lb/sec as compared to 334 lb/sec for JT8D-15) can be achieved with minimum modifications to the existing 727 airplane structure at acceptable levels of total pressure recovery and distortion. Steady-state pressure recovery, steady-state pressure distortion, and dynamic pressure measurements were taken at the engine face station. Surface static pressure measurements were taken along the duct. Test results indicated that the required airflow was achieved with acceptable pressure recovery (comparable to the current 727-200 S duct). Inlet inflow angle variation within the 727 airplane operating regime (minus 5 to 5 degrees) had no effect on the inlet performance. Pressure distortion at static and forward speed at takeoff airflow conditions are within P and WA limits for the Phase II duct when equipped with vortex generators. Static crosswind operation between 10 knots and 25 knots appears feasible at full takeoff power.

Kaldschmidt, G.↗