SEARCH · Search NASA

Results for “Inlet diffusers”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 235 records · Page 13

All sources NASA NTRS DOE OSTI

NASA NTRS1983

Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine

The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.

Mclallin, K. L.↗

NASA NTRS1983

Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine

Mclallin, K. L.↗

NASA NTRS1950

Characteristics of Perforated Diffusers at Free-stream Mach Number 1.90

An investigation was conducted at Mach number 1.90 to determine pressure recovery and mass-flow characteristics of series of perforated convergent-divergent supersonic diffusers. Pressure recoveries as high as 96 percent were obtained, but at reduced mass flows through the diffuser. Theoretical considerations of effect of perforation distribution on shock stability in converging section of diffuser are presented and correlated with experimental data. A method of estimating relative importance of pressure recovery and mass flow on internal thrust coefficient basis is given and a comparison of various diffusers investigated is made.

FLOW, SUPERSONIC↗

NASA NTRS1984

Numerical simulation of flow through inlets/diffusers with terminal shocks

Two-dimensional Euler and Navier-Stokes solutions of the flow through three inlet/diffuser configurations with terminal shock systems are reported. Calculations without bleed indicate that the terminal shock location is very sensitive to the outflow back pressure. For cases where there are no available experimental results, it becomes difficult to estimate the back pressure that will result in a terminal shock. Estimates based on quasi-one-dimensional analysis are not found adequate for complex two-dimensional flows. It is found that since the flow downstream of the terminal shock is subsonic, and what happens at the outflow boundary affects the flow inside the inlet, enough of the subsonic diffuser must be modeled to accurately predict the terminal shock region. The diffuser portion should be fairly long with the outflow boundary occurring in a region of more or less uniform flow to be able to prescribe a uniform back pressure. The third configuration studied was investigated with and without incorporating bleed in the code. It is found that the use of bleed stabilizes the shock location and allows solutions which without bleed result in unstarting of the inlet. Comparisons are made with available experimental data.

Talcott, N. A., Jr.↗

NASA NTRS1990

Control of flow separation and mixing by aerodynamic excitation

The recent research in the control of shear flows using unsteady aerodynamic excitation conducted at the NASA Lewis Research Center is reviewed. The program is of a fundamental nature, concentrating on the physics of the unsteady aerodynamic processes. This field of research is a fairly new development with great promise in the areas of enhanced mixing and flow separation control. Enhanced mixing research includes influence of core turbulence, forced pairing of coherent structures, and saturation of mixing enhancement. Separation flow control studies included are for a two-dimensional diffuser, conical diffusers, and single airfoils. Ultimate applications include aircraft engine inlet flow control at high angle of attack, wide angle diffusers, highly loaded airfoils as in turbomachinery, and ejector/suppressor nozzles for the supersonic transport. An argument involving the Coanda Effect is made that all of the above mentioned application areas really only involve forms of shear layer mixing enhancement. The program also includes the development of practical excitation devices which might be used in aircraft applications.

Rice, Edward J.↗

NASA NTRS1981

A summary of V/STOL inlet analysis methods

Methods to analyze the aerodynamic performance of V/STOL inlets at the NASA Lewis Research Center are discussed with emphasis on recent extensions and applications. Methods include the specification of the Kutta condition for a slotted inlet, the calculation of suction and tangential blowing for boundary layer control, and the analysis of auxiliary inlet geometries at angles of attack. An optimum diffuser velocity distribution is also developed, which can result in the shortest no-boundary layer control inlet and the lowest loss for the required amount of diffusion.

Hwang, D. P.↗

NASA NTRS1981

Extension to an analysis of turbulent swirling compressible flow for application to axisymmetric small gas turbine ducts

An existing computer program, the Axisymmetric Diffuser Duct Code (ADD code), which calculates compressible turbulent swirling flow through axisymmetric ducts was modified to permit calculation of flows through small gas turbine ducts with struts, guide vanes and large degrees of turning. The improvements include a coordinate generator, an end-wall loss model, and a generalized geometry capability to describe struts and guide vanes in ducts which turn more than 90 degrees. An improved output format was developed to provide the solution on any arbitrary plane in the duct and an extensive literature survey of calculation procedures used in gas turbine technology was completed which suggests improvements in the computer code. Calculations are presented for the flow through the AGT101 small gas turbine inlet duct and turbine exhaust diffuser which demonstrate the ADD code modifications implemented in the investigation. The computed results compare favorably with experimental results.

Anderson, O. L.↗

NASA NTRS1974

Distributed porous throat stability bypass to increase the stable airflow range of a Mach 2.5 inlet with 60 percent internal contraction

The results of an experimental investigation to increase the stable airflow operating range of a supersonic, mixed-compression inlet with 60-percent internal contraction are presented. Various distributed-porous, throat stability-bypass entrance configurations were tested. In terms of diffuser-exit corrected airflow, a large inlet stable airflow range of about 25 percent was obtained with the optimum configuration if a constant pressure was maintained in the by-pass plenum. The location of the centerbody bleed region had a decided effect on the overall inlet performance. Limited unstart angle-of-attack data are presented.

Shaw, R. J.↗

NASA NTRS1952

Experimental Investigation of a 90 Degree Cascade Diffusing Bend with an Area Ratio of 1.45:1 and with Several Inlet Boundary Layers

Explore the source record for details and available documents.

Friedman, Daniel↗

NASA NTRS1959

Investigation on the use of a Freely Rotating Rotor at the Cowl Face of a Supersonic Conical Inlet to Reduce Inlet Flow Distortion

An investigation has been made on the use of a freely rotating rotor at the cowl face of a supersonic conical diffuser to determine its effectiveness in reducing inlet flow distortion and the penalty in terms of total-pressure loss imposed by such a device when distortions are negligible. Tests were made with a rotor having an inlet tip diameter of 2.18 inches and a ratio of hub radius to tip radius of 0.52, in conjunction with a conical inlet having a 25 deg semi-vertex cone angle, at a Mach number of 2.1 over an angle-of-attack range of 0 deg to 8 deg. A simplified analysis showing that a supersonic, freely rotating rotor with maximum solidity for noninterference between blades will operate in an undistorted flow with a total-pressure defect of 1 percent or less was experimentally verified. Overall total-pressure distortions of 0.1 to 0.4 and Mach number distortions of 0.4 to 1.4, obtained at 4 deg to 8 deg angle of attack, were reduced about 30 percent and 23 percent, respectively, because of the presence of the rotor, with no measurable total-pressure loss. The rotor increased the peak total-pressure recovery at the simulated combustion chamber 1 1/2 and 3 1/2 percent at 6 deg and 8 deg angles of attack, respectively. This increase is attributed to lower diffusion duct losses as a consequence of a more uniform flow created by the rotor.

Goldberg, Theodore J.↗

NASA NTRS1981

Flow separation in inlets at incidence angles

Wind-tunnel pressure data and flow pictures obtained for two two-dimensional inlet models have been examined to study the internal flow structure and separation at large incidence angles. The inlet models were 12-in. high (diffuser exit height) and had internal contraction ratio of 1.21 and 1.17. They were tested at low forward speeds over a wide range of throat Mach numbers (inlet mass flow rates) and angles of incidence. Characteristic features of the internal flow such as a drastic change of pressure gradient near the highlight, local separation bubbles and shock/boundary-layer interactions have been indicated and discussed. For a few specific cases, the experimental surface pressure distributions have been compared with theoretical predictions.

Jakubowski, A. K.↗

NASA NTRS2016

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Jones, Daniel R.↗

NASA NTRS1976

Aerodynamic design guidelines and computer program for estimation of subsonic wind tunnel performance

General guidelines are given for the design of diffusers, contractions, corners, and the inlets and exits of non-return tunnels. A system of equations, reflecting the current technology, has been compiled and assembled into a computer program (a user's manual for this program is included) for determining the total pressure losses. The formulation presented is applicable to compressible flow through most closed- or open-throat, single-, double-, or non-return wind tunnels. A comparison of estimated performance with that actually achieved by several existing facilities produced generally good agreement.

Eckert, W. T.↗

NASA NTRS1979

A throat-bypass stability-bleed system using relief valves to increase the transient stability of a mixed-compression inlet

A stability-bleed system was installed in a YF-12 flight inlet that was subjected to internal and external airflow disturbances in the NASA Lewis 10 by 10 foot supersonic wind tunnel. The purpose of the system is to allow higher inlet performance while maintaining a substantial tolerance (without unstart) to internal and external disturbances. At Mach numbers of 2.47 and 2.76, the inlet tolerance to decreases in diffuser-exit corrected airflow was increased by approximately 10 percent of the operating-point airflow. The stability-bleed system complemented the terminal-shock-control system of the inlet and did not show interaction problems. For disturbances which caused a combined decrease in Mach number and increase in angle of attack, the system with valves operative kept the inlet started 4 to 28 times longer than with the valves inoperative. Hence, the stability system provides additional time for the inlet control system to react and prevent unstart. This was observed for initial Mach numbers of 2.55 and 2.68. For slow increase in angle of attack at Mach 2.47 and 2.76, the system kept the inlet started beyond the steady-state unstart angle. However, the maximum transient angles of attack without unstart could not be determined because wind-tunnel mechanical-stop limits for angle of attack were reached.

Neiner, G. H.↗

NASA NTRS1987

Effect of shaft rotation on the incompressible flow in a labyrinth seal

The incompressible flow in a labyrinth seal at low leakage rates was computationally and experimentally investigated over a wide range of seal rotation rates. QUICK differencing was employed in the finite difference code to reduce the effects of false diffusion. The use of measured inlet boundary conditions for the axial and swirl velocity components and for the turbulent kinetic energy resulted in good agreement between velocity predictions and hot-film measurements. It was found that when the rotation rate is increased beyond a certain point, a second recirculation zone forms inside the seal cavity, altering the flow field in the cavity and resulting in a substantial increase in the pressure drop across it.

Demko, J. A.↗

NASA NTRS1998

Design, Operation, and Modeling of a Vertical APCVD Reactor for Silicon Carbide Film Growth

An atmospheric pressure chemical vapor deposition (APCVD) reactor utilizing a unique vertical geometry which enables 3C-SiC films to be grown on two, 4-inch diameter Si wafers has been constructed. Contrary to expectations, 3C-SiC films grown in this reactor are thickest at the downstream end of the substrates. To better understand the reason for the thickness distribution on the wafers, an axisymmetric finite-element model of the gas flow in the reactor was constructed. The model uses the ANSYS53 Flowtran package and includes compressible and temperature-dependent fluid properties in laminar or turbulent flow. It does not include reaction chemistry or unsteady flow. The ANSYS53 results predict that the cool, inlet fluid falls through the inlet pipe and the warm, diffuser region like a jet. This jet impinges on top of the susceptor and gets diverted to the reactor side walls, where it flows to the bottom of the reactor, turns, and slowly rises along the face of the susceptor. This may explain why the SiC films are thickest at the downstream side of the wafers, as gas containing fresh reactants first passes over this region. Modeling results are presented for both one atmosphere and one half atmosphere reactor pressure.

DeAnna, Russell G.↗

NASA NTRS2004

Algorithm for Controlling a Centrifugal Compressor

An algorithm has been developed for controlling a centrifugal compressor that serves as the prime mover in a heatpump system. Experimental studies have shown that the operating conditions for maximum compressor efficiency are close to the boundary beyond which surge occurs. Compressor surge is a destructive condition in which there are instantaneous reversals of flow associated with a high outlet-to-inlet pressure differential. For a given cooling load, the algorithm sets the compressor speed at the lowest possible value while adjusting the inlet guide vane angle and diffuser vane angle to maximize efficiency, subject to an overriding requirement to prevent surge. The onset of surge is detected via the onset of oscillations of the electric current supplied to the compressor motor, associated with surge-induced oscillations of the torque exerted by and on the compressor rotor. The algorithm can be implemented in any of several computer languages.

Benedict, Scott M.↗

NASA NTRS1961

The Development of an 8-inch by 8-inch Slotted Tunnel for Mach Numbers up to 1.28

An 8-inch by 8-inch transonic tunnel model with test section slotted on two opposite walls was constructed in which particular emphasis -was given to the development of slot geometry, slot-flow reentry section, and short-diffuser configurations for good test-region flow and minimum total-pressure losses. Center-line static pressures through the test section, wall static pressures through the other parts of the tunnel, and total-pressure distributions at the inlet and exit stations of the diffuser were measured- With a slot length equal to two tunnel heights and 1/14 open-area-ratio slotted walls) a test region one tunnel height in length was obtained in which the deviation from the mean Mach number was less than +/- 0.01 up to Mach number 1.15. With 1/7 open-area-ratio slotted walls, a test region 0.84 tunnel heights in length with deviation less than +/- O.01 was obtained up to Mach number 1.26. Increasing the tunnel diffuser angle from 6.4 to 10 deg. increased pressure loss through the tunnel at Mach number 1.20 from 15 percent to 20 percent of the total pressure. The use of other diffusers with equivalent angles of 10 deg. but contoured so that the initial diffusion angle was less than 10 deg. and the final angle was 200 reduced the losses to as low as 16 percent. A method for changing the test-section Mach number rapidly by controlling the flow through a bypass line from the tunnel settling chamber to the slot-flow plenum chamber of the test section was very effective. The test-section Mach number was reduced approximately 5 percent in 1/8 second by bleeding into the test section a flow of air equal to 2 percent of the mainstream flow and 30 percent in 1/4 second with bleed flow equal to 10 percent of the mainstream flow. The rate of reduction was largely determined by the opening rate of the bleed-flow-control valve.

Little, B. H., Jr.↗