SEARCH · Search NASA

Results for “Airflow configurations”

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 19 records

All sources NASA NTRS DOE OSTI

NASA NTRS1991

V/STOL gets a lift

The concept of a supersonic STOVL that could offer enhanced mission capability, survivability, operational flexibility, and utility over conventional aircraft is presented. Emphasis is currently on design studies, CFD work, small- and large-scale wind tunnel tests, simulation activities, flight experiments, and ground environment experiments. Propulsion system technology centers about the adaptation of existing or off-the-shelf engines. Concepts under study include separate flow in hover, gas-driven lift fan, and shaft-driven lift fan. NASA is examining generic valve and ducting configurations with airflow at ambient temperature and at temperatures up to 1000 F to gather pressure loss and heat transfer data. Advanced civil rotorcraft technologies examined include high-efficiency/dual-mode components such as torque converters; lightweight, quiet transmissions; and variable geometry power turbines; along with dual-function or convertible engines.

Biesiadny, Tom↗

NASA NTRS1992

Tests characterizing bioprocessor hardware for analytical modeling

The tests outlined in this paper were used to characterize the hardware components of the Salad Machine, a small NASA-developed bioprocessor. The data from these tests are presented, and the methods by which this data can be integrated into system mathematical models are briefly discussed. The subsystems and physical processes discussed include the lighting system, the air loop (condensing heat exchanger and the blower), heat transfer to the surroundings, and leakage. Through this effort it was learned that in the development of a test protocol, care should be taken to order the tests such that environmental parameters, particularly humidity, require as few large adjustments as possible. Sensor calibration and installation take a substantial amount of time, which should be built into the test schedule. Two properties were particularly hard to quantify: the air flow rate and the energy from the lighting system entering into the growth volume. Flow rate can be measured using the appropriate device for the system configuration and airflow. Lighting system radiation level was measured using three methods. The results of these methods varied substantially, putting off conclusive quantification of this value.

Gustavino, S.↗

NASA NTRS1977

Experimental evaluation of fuel preparation systems for an automotive gas turbine catalytic combustor

Spatial fuel distributions, degree of vaporization, pressure drop and air velocity profiles were measured. Three airblast injectors and an air-assist nozzle were tested. Air swirlers were used to improve the spatial fuel-air distribution. The work was done in a 12 cm tubular duct. Test conditions were: a pressure of 0.3 and 0.5 MPa, inlet air temperatures up to 800 K, air velocities of 10 20 m/s and fuel-air ratios up to 0.020. The fuel was Jet A. The best results were obtained with an airblast configuration that used multiple cones to provide high velocity air for atomization and also straightened the inlet airflow. With this configuration, uniform spatial fuel-air distributions were obtained with mixing lengths greater than 17.8 cm. In this length, vaporization of the fuel was 98.5 percent complete at an inlet air temperature of 700 K. The total pressure loss was 1.0 percent with a reference velocity of 20 m/s and 0.25 percent at 10m/s. The air velocity was uniform across the duct and no autoignition reactions were observed.

Tacina, R. R.↗

NASA NTRS1977

Experimental evaluation of fuel preparation systems for an automotive gas turbine catalytic combustor

Premixing-prevaporizing fuel systems were evaluated for use with a catalytic reactor for possible automotive gas turbine application. Spatial fuel-air distributions, degree of vaporization, pressure drop and air velocity profiles were measured. Three airblast injectors and an air-assist nozzle were tested. Air swirlers were used to improve the spatial fuel-air distribution. The work was done in a 12 cm tubular duct. Test conditions were: a pressure of 0.3 and 0.5 MPa, inlet air temperatures up to 800 K, air velocities of 10 and 20 m/s and fuel-air ratios up to 0.020. The fuel was Jet A. The best results were obtained with an air-blast configuration that used multiple cones to provide high velocity air for atomization and also straightened the inlet airflow. With this configuration, uniform spatial fuel-air distributions were obtained with mixing lengths greater than 17.8 cm. In this length, vaporization of the fuel was 98.5 percent complete at an inlet air temperature of 700 K. The total pressure loss was 1.0 percent with a reference velocity of 20 m/s and 0.25 percent at 10 m/s. The air velocity was uniform across the duct and no autoignition reactions were observed.

Tacina, R. R.↗

NASA NTRS1991

3-D CFD in a day - The laser digitizer project

The computation of airflow over complex configurations requires a complete description of the geometry. This can be obtained from CAD data, from blueprints, or from actual models. In any case, the time required is currently estimated at 4 to 6 months. It is proposed to shorten this time by a factor of 10 to 100 through the use of automated software, a fast, highly parallel computer and a three-dimensional laser digitizer. This device can provide (x,y,z) coordinates of surface points at rates exceeding 14,500/sec. Thus, it is possible to digitize an entire model in a few minutes. The accuracy of measurement on a flat white surface is better than 0.005 inches. Higher accuracy is available at higher cost. This work discusses the challenges which remain to be addressed. In particular, the surface point data need to be converted into a surface description, the surface description needs to be made into a surface grid, and the surface grid used to make a volume grid for the flow solver. Algorithms are kept in place or in mind for all of these problems. Integration of the more mature flow solution and visualization algorithms then allows generation of solution graphics directly from a wind tunnel model.

Merriam, Marshal↗

NASA NTRS1974

JT8D-100 turbofan engine, phase 1

The JT8D turbofan engine, widely used in short and medium range transport aircraft, contributes substantially to airport community noise. The jet noise is predominant in the JT8D engine and may be reduced in a modified engine, without loss of thrust, by increasing the airflow to reduce jet velocity. A configuration study evaluated the effects of fan airflow, fan pressure ratio, and bypass ratio on noise, thrust, and fuel comsumption. The cycle selected for the modified engine was based upon an increased diameter, single-stage fan and two additional core engine compressor stages, which replace the existing two-stage fan. Modifications were also made to the low pressure turbine to provide the increased torque required by the larger diameter fan. The resultant JT8D-100 engine models have the following characteristics at take-off thrust, compared to the current JT8D engine: Airflow and bypass ratio are increased, and fan pressure ratio and engine speed are reduced. The resultant engine is also longer, larger in diameter, and heavier than the JT8D base model, but these latter changes are compensated by the increased thrust and decreased fuel comsumption of the modified engine, thus providing the capability for maintaining the performance of the current JT8D-powered aircraft.

Source record↗

NASA NTRS2012

Air monitoring device

An air monitoring device (100) includes an outer casing (101) configured to receive an airflow (102) comprising particulate; a bore (103) located inside the outer casing (101); and a collection probe (104) located inside the outer casing (101), the collection probe (104) being configured such that there is a gap (105) between an exit of the bore (103) and an entrance of the collection probe (104), such that particulate in the airflow (102) having a diameter larger than a threshold flows through an interior of the collection probe (104).

Tissandier, Michael D.↗

NASA NTRS1977

Jet engine air intake system

An axisymmetric air intake system for a jet aircraft engine comprising a fixed cowl extending outwardly from the face of the engine, a centerbody coaxially disposed within the cowl, and an actuator for axially displacing the centerbody within the cowl was developed. The cowl and centerbody define a main airflow passageway therebetween, the configuration of which is changed by displacement of the centerbody. The centerbody includes a forwardly-located closeable air inlet which communicates with a centerbody auxiliary airflow passageway to provide auxiliary airflow to the engine. In one embodiment, a system for opening and closing the centerbody air inlet is provided by a dual-member centerbody, the forward member of which may be displaced axially with respect to the aft member.

Sorensen, N. E.↗

NASA NTRS1975

Poppet valve control of throat stability bypass to increase stable airflow range of a Mach 2.5. inlet with 60 percent internal contraction

The throat of a Mach 2.5 inlet with a coldpipe termination was fitted with a stability-bypass system. System variations included several stability bypass entrance configurations. Poppet valves controlled the bypass airflow. The inlet stable airflow range achieved with each configuration was determined for both steady state conditions and internal pulse transients. Results are compared with those obtained without a stability bypass system. Transient results were also obtained for the inlet with a choke point at the diffuser exit and for the inlet with large and small stability bypass plenum volumes. Poppet valves at the stability bypass exit provided the inlet with a stable airflow range of 20 percent or greater at all static and transient conditions.

Mitchell, G. A.↗

NASA NTRS1975

An investigation of airflow over the aft portions of a variable sweep fighter configuration

An investigation of air flow over the aft portions of a variable sweep fighter aircraft configuration was made. Tests conducted in the unitary plan wind tunnel at Mach number 2.16 included measurements of forces, moments, and local static pressures as well as visual recordings of the air flow. An aerodynamic analytical prediction method was evaluated when used in data comparison at angles of attack of 0, 5, and 15 degrees. The results indicate that in supersonic flow the typical outboard located twin vertical tail arrangement tends to provide a more positive increment in normal-force on the afterbody fuselage and the horizontal tail than a single center-mounted vertical tail of similar planform shape. In addition, the results indicate that a method for aerodynamic analysis of wing-body-tail configurations currently available can provide reasonable estimates of pressure coefficient distributions on configurations in regions of complex supersonic flow. At this time, however, the available analytical method cannot adequately replace experimental wind tunnel tests for determining the supersonic flow environment of a given configuration.

Graves, E. B.↗

NASA NTRS2011

Blended Wing Body Concept Development with Open Rotor Engine Intergration

The purpose of this study is to perform a systems analysis of a Blended Wing Body (BWB) open rotor concept at the conceptual design level. This concept will be utilized to estimate overall noise and fuel burn performance, leveraging recent test data. This study will also investigate the challenge of propulsion airframe installation of an open rotor engine on a BWB configuration. Open rotor engines have unique problems relative to turbofans. The rotors are open, exposed to flow conditions outside of the engine. The flow field that the rotors are immersed in may be higher than the free stream flow and it may not be uniform, both of these characteristics could increase noise and decrease performance. The rotors sometimes cause changes in the flow conditions imposed on aircraft surfaces. At high power conditions such as takeoff and climb out, the stream tube of air that goes through the rotors contracts rapidly causing the boundary layer on the body upper surface to go through an adverse pressure gradient which could result with separated airflow. The BWB / Open Rotor configuration must be designed to mitigate these problems.

Pitera, David M.↗

NASA NTRS1974

Low-speed test of translating lip axisymmetric inlets for subsonic transports

Translating lip, axisymmetric inlets have been tested at freestream Mach numbers between 0.0 and 0.30 and at angles of attack between 0 and 90 degrees. Three isolated inlet models were tested. One model represented the inlet geometry at cruise, that is, with no lip translation. The other two models had forward translating lips with different contours. The low forward speed inlet angle-of-attack capability with the translating lip was increased more than twenty degrees over the basic untranslated lip configuration at the higher inlet airflows. The static inlet performance with the translating lip was also greatly improved. The design translation distance, defined by potential flow analysis, was near optimum.

Henne, P. A.↗

NASA NTRS1988

An analysis for high speed propeller-nacelle aerodynamic performance prediction. Volume 1: Theory and application

A computer program, the Propeller Nacelle Aerodynamic Performance Prediction Analysis (PANPER), was developed for the prediction and analysis of the performance and airflow of propeller-nacelle configurations operating over a forward speed range inclusive of high speed flight typical of recent propfan designs. A propeller lifting line, wake program was combined with a compressible, viscous center body interaction program, originally developed for diffusers, to compute the propeller-nacelle flow field, blade loading distribution, propeller performance, and the nacelle forebody pressure and viscous drag distributions. The computer analysis is applicable to single and coaxial counterrotating propellers. The blade geometries can include spanwise variations in sweep, droop, taper, thickness, and airfoil section type. In the coaxial mode of operation the analysis can treat both equal and unequal blade number and rotational speeds on the propeller disks. The nacelle portion of the analysis can treat both free air and tunnel wall configurations including wall bleed. The analysis was applied to many different sets of flight conditions using selected aerodynamic modeling options. The influence of different propeller nacelle-tunnel wall configurations was studied. Comparisons with available test data for both single and coaxial propeller configurations are presented along with a discussion of the results.

Egolf, T. Alan↗

NASA NTRS2026

A Modular Conjugate Heat Transfer Optimization Framework for Thermal Management of Electric Aircraft

Conjugate heat transfer (CHT) analysis and optimization is a powerful method for improving thermal management, as it simultaneously resolves the temperature distribution in both fluid and solid domains. This paper presents a modular, discrete adjoint-based CHT optimization capability integrated within the OpenMDAO/MPhys framework. A unique feature of the proposed framework is its flexibility to extend to multidisciplinary optimization, including aero-structural-thermal applications. The fluid domain is modeled using a finite-volume Computational Fluid Dynamics (CFD) solver, and the solid domain with a conduction heat transfer solver. A mixed Neumann-Dirichlet boundary condition is developed to enable full submersion of the solid geometry within the fluid domain, while ensuring consistent temperature and heat flux coupling at the CHT interface. Gradient-based optimization is performed; the gradients are efficiently computed using the discrete adjoint solvers implemented in DAFoam. To demonstrate the method, this paper considers two cases related to electric aircraft thermal management: a U-bend heat exchanger and an actively cooled battery pack. The U-bend case aims to minimize pressure loss while maximizing heat flux by changing the pipe geometry. The optimized design reduces pressure loss by 52.7% and increases total heat flux by 2.3%. In the battery pack case, a 3-by-3 cell configuration is cooled by ambient airflow, with constant heat generation prescribed in the cells. The battery casing shape serves as the design variable, and the objective function is a weighted sum of pressure loss and pack weight, subject to a maximum temperature constraint. The optimized design achieves a 44.6% reduction in pressure loss and a 1.5% reduction in weight, while satisfying the thermal constraint. To ensure the reliability of the optimized designs, this study validates coarse-mesh, steady-state predictions against fine-mesh unsteady simulations, demonstrating consistency within acceptable errors. This work demonstrates the potential of the developed framework to enable rapid, high-fidelity design of thermal management systems for electric aircraft.

heat transfer↗

NASA NTRS1974

The 727 airplane target thrust reverser static performance model test for refanned JT8D engines

The results of a scale model static performance test of target thrust reverser configurations for the Pratt and Whitney Aircraft JT8D-100 series engine are presented. The objective of the test was to select a series of suitable candidate reverser configurations for the subsequent airplane model wind tunnel ingestion and flight controls tests. Test results indicate that adequate reverse thrust performance with compatible engine airflow match is achievable for the selected configurations. Tapering of the lips results in loss of performance and only minimal flow directivity. Door pressure surveys were conducted on a selected number of lip and fence configurations to obtain data to support the design of the thrust reverser system.

Chow, C. T. P.↗

NASA NTRS2002

An Aeroacoustic Study of a Leading Edge Slat Configuration

Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper describes detailed flow and acoustic measurements that have been made in order to better understand the noise generated from airflow over a wing leading edge slat configuration, and to possibly predict and reduce this noise source. The acoustic database is obtained by a moveable Small Aperture Directional Array of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.

Mendoza, J. M.↗

NASA NTRS1974

Forward slanted slot 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 are presented. Two forward-slanted slot stability-bypass entrance configurations were tested. In terms of diffuser-exit corrected airflow, a large inlet stable airflow range of 18.5 percent was obtained with the superior configuration if a constant pressure was maintained in the bypass plenum. Limited unstart angle-of-attack data are presented.

Shaw, R. J.↗

NASA NTRS1978

Experimental investigation of a 0.15-scale model of an underfuselage normal-shock inlet

A 0.15 scale model of an underfuselage inlet designed for a single-engine fighter airplane was tested. The inlet was a fixed-geometry, normal-shock configuration designed to operate at flight speeds up to Mach 2.0. Peformance data for the basic inlet and several configuration variations are presented as a function of angle of attack, angle of sideslip, and airflow in the 0.6 to 2.0 Mach number range. The configuration variations included boundary-layer diverter height, cowl and splitter-plate modifications, and inlet bleed system variations. Flow-field characteristics at the simulated engine face, at the inlet throat, at the splitter-plate leading edge, and forward of the inlet are presented. The pressure recovery of the inlet is approximately equal to the product of theoretical normal-shock and duct pressure recoverable at cruise angle of attack. Very good performance at high angle of attack was obtained. Pressure distortion and turbulence at the engine face were low, and the inlet remained stable at all engine airflows over the flight maneuver envelope of the aircraft for which the inlet was designed.

Leamer, P. C.↗