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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 NTRS1992

NASA F-16XL supersonic laminar flow control program overview

The viewgraphs and discussion of the NASA supersonic laminar flow control program are provided. Successful application of laminar flow control to a High Speed Civil Transport (HSCT) offers significant benefits in reductions of take-off gross weight, mission fuel burn, cruise drag, structural temperatures, engine size, emissions, and sonic boom. The ultimate economic success of the proposed HSCT may depend on the successful adaption of laminar flow control, which offers the single most significant potential improvements in lift drag ratio (L/D) of all the aerodynamic technologies under consideration. The F-16XL Supersonic Laminar Flow Control (SLFC) Experiment was conceived based on the encouraging results of in-house and NASA supported industry studies to determine if laminar flow control is feasible for the HSCT. The primary objective is to achieve extensive laminar flow (50-60 percent chord) on a highly swept supersonic wing. Data obtained from the flight test will be used to validate existing Euler and Navier Stokes aerodynamic codes and transition prediction boundary layer stability codes. These validated codes and developed design methodology will be delivered to industry for their use in designing supersonic laminar flow control wings. Results from this experiment will establish preliminary suction system design criteria enabling industry to better size the suction system and develop improved estimates of system weight, fuel volume loss due to wing ducting, turbocompressor power requirements, etc. so that benefits and penalties can be more accurately assessed.

Fischer, Michael C.↗
NASA NTRS1988

Evaluation of FIDAP on some classical laminar and turbulent benchmarks

The numerical accuracy of the fluid-dynamics code FIDAP is investigated by means of test computations on two-dimensional flows. The theoretical basis of the algorithm is briefly reviewed, and results for laminar flow in a wall-driven cavity and for both laminar and turbulent flows over a backward-facing step are presented in extensive tables and graphs and characterized in detail. Good agreement with published experimental data and previous computations is obtained for laminar flows using a version of FIDAP without streamlined upwinding (STU). For turbulent flows, the addition of STU is found to be useful, although it causes separation-zone reattachment length to be underpredicted.

Sohn, J. L.↗
NASA NTRS1971

Fundamental ignition study for material fire safety improvement, part 2

The autoignition behavior of polymeric compositions in oxidizing media was investigated as well as the nature and relative concentration of the volatiles produced during oxidative decomposition culminating in combustion. The materials investigated were Teflon, Fluorel KF-2140 raw gum and its compounded versions Refset and Ladicote, 45B3 intumenscent paint, and Ames isocyanurate foam. The majority of the tests were conducted using a stagnation burner arrangement which provided a laminar gas flow and allowed the sample block and gas temperatures to be varied independently. The oxidizing atmospheres were essentially air and oxygen, although in the case of the Fluorel family of materials, due to partial blockage of the gas inlet system, some tests were performed unintentionally in enriched air (not oxygen). The 45B3 paint was not amenable to sampling in a dynamic system, due to its highly intumescent nature. Consequently, selected experiments were conducted using a sealed tube technique both in air and oxygen media.

Paciorek, K. L.↗
NASA NTRS1988

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1030:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

Smith, Tamara A.↗
NASA NTRS1988

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1039:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

Smith, Tamara A.↗
NASA NTRS2013

Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft s swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First AIAA CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

Bui, Trong↗
NASA NTRS2014

Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft's swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First American Institute of Aeronautics and Astronautics (AIAA) CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

Bui, Trong T.↗
NASA NTRS2014

Analysis of Low Speed Stall Aerodynamics of a Swept Wing with Laminar Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft's swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First American Institute of Aeronautics and Astronautics (AIAA) CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

computational fluid dynamics↗
NASA NTRSDate not supplied

A Preliminary Assessment of Step Effects on the BOLT Geometry in Mach 6 Flow

An infrared thermography measurement technique has been installed in the Langley Aerothermodynamic Laboratory. The system is calibrated for use with commercially available thermoplastics and ceramic materials, and to the surface temperature ranges when exposed to the facility hypersonic flow. Herein an uncertainty quantification is performed, revealing 5% surface temperature and 20% heat transfer uncertainty bands to two-sigma confidence. As a first application of the test measurement technique, an investigation was conducted to explore the correlation between the height of forward facing (FF) and rearward facing (RF) two-dimensional steps, free stream Reynolds number, and the onset of boundary layer transition from laminar to turbulent in Mach 6 flow on a 50% scale Boundary Layer Transition (BOLT) flight test vehicle geometry. Computational predictions were generated to identify potential x/L locations of transition onset for various flow conditions with a smooth outer mold line geometry and are compared to experimental results. The BOLT Step Test was conducted in the NASA Langley Aerothermodynamic Laboratory 20-inch Mach 6 wind tunnel at Reynolds numbers between 1.4 and 11.3 million conditions for various step heights. Data were collected using a developmental infrared thermography test technique and reduced to heat transfer for comparison with CFD solutions. This is a preliminary assessment of the developmental infrared thermography test technique used to acquire and reduce the experimental data. insight into the ongoing comparative analysis of the effects of FF and RF steps on boundary layer transition for the BOLT geometry in Mach 6 flow.

BOLT↗
NASA NTRSDate not supplied

A Preliminary Assessment of Step Effects on the BOLT Geometry in Mach 6 Flow

An infrared thermography measurement technique has been installed in the Langley Aerothermodynamic Laboratory. The system is calibrated for use with commercially available thermoplastics and ceramic materials, and to the surface temperature ranges when exposed to the facility hypersonic flow. Herein an uncertainty quantification is performed, revealing 5% surface temperature and 20% heat transfer uncertainty bands to two-sigma confidence. As a first application of the test measurement technique, an investigation was conducted to explore the correlation between the height of forward facing (FF) and rearward facing (RF) two-dimensional steps, free stream Reynolds number, and the onset of boundary layer transition from laminar to turbulent in Mach 6 flow on a 50% scale Boundary Layer Transition (BOLT) flight test vehicle geometry. Computational predictions were generated to identify potential x/L locations of transition onset for various flow conditions with a smooth outer mold line geometry and are compared to experimental results. The BOLT Step Test was conducted in the NASA Langley Aerothermodynamic Laboratory 20-inch Mach 6 wind tunnel at Reynolds numbers between 1.4 and 11.3 million conditions for various step heights. Data were collected using a developmental infrared thermography test technique and reduced to heat transfer for comparison with CFD solutions. This is a preliminary assessment of the developmental infrared thermography test technique used to acquire and reduce the experimental data. insight into the ongoing comparative analysis of the effects of FF and RF steps on boundary layer transition for the BOLT geometry in Mach 6 flow.

BOLT↗
NASA NTRS1975

Effect of rod gap spacing on a suction panel for laminar flow and noise control in supersonic wind tunnels

Results are presented of a coordinated experimental and theoretical study of a sound shield concept which aims to provide a means of noise reduction in the test section of supersonic wind tunnels at high Reynolds numbers. The model used consists of a planar array of circular rods aligned with the flow, with adjustable gaps between them for boundary layer removal by suction, i.e., laminar flow control. One of the basic requirements of the present sound shield concept is to achieve sonic cross flow through the gaps in order to prevent lee-side flow disturbances from penetrating back into the shielded region. Tests were conducted at Mach 6 over a local unit Reynolds number range from about 1.2 x 10 to the 6th power to 13.5 x 10 to the 6th power per foot. Measurements of heat transfer, static pressure, and sound levels were made to establish the transition characteristics of the boundary layer on the rod array and the sound shielding effectiveness.

Harvey, W. D.↗
NASA NTRS2007

Turbulent Aeroheating Testing of Mars Science Laboratory Entry Vehicle in Perfect-Gas Nitrogen

An experimental investigation of turbulent aeroheating on the Mars Science Laboratory entry vehicle heat shield has been conducted in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9. Testing was performed on a 6-in. (0.1524 m) diameter MSL model in pure N2 gas in the tunnel s Mach 8 and Mach 10 nozzles at free stream Reynolds numbers of 4.1x10(exp 6)/ft to 49x10(exp 6)/ft (1.3x10(exp 7)/m to 16x10(exp 7)/m) and 1.2x10(exp 6)/ft to 19x10(exp 6)/ft (0.39x10(exp 7)/m to 62x10(exp 7)/m), respectively. These conditions were sufficient to span the regime of boundary-layer flow from completely laminar to fully-developed turbulent flow over the entire forebody. A supporting aeroheating test was also conducted in the Langley Research Center 20-Inch Mach 6 Air Tunnel at free stream Reynolds number of 1x10(exp 6)/ft to 7x10(exp 6)/ft (0.36x10(exp 7)/m to 2.2x10(exp 7)/m) in order to help corroborate the Tunnel 9 results. A complementary computational fluid dynamics study was conducted in parallel to the wind tunnel testing. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins on predictions for aeroheating environments during entry into the Martian atmosphere. Data from both wind tunnel tests and comparisons with the predictions are presented herein. It was concluded from these comparisons that for perfect-gas conditions, the computational tools could predict fully-laminar or fully-turbulent heating conditions to within 10% of the experimental data

Hollis, Brian R.↗
NASA NTRS2008

Turbulent Aeroheating Testing of Mars Science Laboratory Entry Vehicle

An experimental investigation of turbulent aeroheating on the Mars Science Laboratory entry vehicle heat shield has been conducted in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9. Testing was performed on a 6-in. (0.1524 m) diameter MSL model in pure N2 gas in the tunnel's Mach 8 and Mach 10 nozzles at free stream Reynolds numbers of 4.1 x 10(exp 6)/ft to 49 x 10(exp 6)/ft (1.3 x 10(exp 7)/m to 19 x 10(exp 6/ft) and 1.2 x 10(exp 6)/ft to 19 x 10(exp 6)/ft (0.39 x 10(exp 7)/m to 62 x 10(exp 7)/m), respectively. These conditions were sufficient to span the regime of boundary-layer flow from completely laminar to fully-developed turbulent flow over the entire forebody. A supporting aeroheating test was also conducted in the Langley Research Center 20-Inch Mach 6 Air Tunnel at free stream Reynolds number of 1 x 10(exp 6)/ft to 7 x 10(exp 6)/ft (0.36 x 10(exp 7)/m to 2.2 x 10(exp 7)/m) in order to help corroborate the Tunnel 9 results. A complementary computational fluid dynamics study was conducted in parallel to the wind tunnel testing. Laminar and turbulent predictions were generated for the wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins on predictions for aeroheating environments during entry into the Martian atmosphere. Data from both wind tunnel tests and comparisons with the predictions are presented herein. It was concluded from these comparisons that for perfect-gas conditions, the computational tools could predict fully-laminar or fully-turbulent heating conditions to within 12% or better of the experimental data.

Hollis, Brian R.↗
NASA NTRSDate not supplied

Transition Modeling Based on the Dual N-factor Method for the CRM-NLF Wind Tunnel Configuration

The dual N-factor method is used to model the boundary-layer transition over the common research model with natural laminar flow (CRM-NLF) aircraft configuration. The flow conditions match selected test conditions from a wind tunnel experiment in the National Transonic Facility at the NASA Langley Research Center. The paper presents a systematic methodology for transition prediction in the presence of a dual shock system and extends the prior capability for iteratively coupled computational fluid dynamics (CFD) predictions to incorporate three-dimensional, transonic wings. The method employs stability computations based on the linear parabolized stability equations (PSE), along with a dual N-factor criterion. The iterative process begins with the fully turbulent Reynolds-averaged-Navier-Stokes (RANS) mean flow solution. For the first iteration, a mean flow solution is calculated with an imposed transition front that aligns with the shock front from the fully turbulent solution. Subsequently, stability computations are performed along a set of streamlines across the wing to calculate the amplification of planar Tollmien-Schlichting (TS) and stationary crossflow (CF) modes. The transition criterion based on the dual N-factor method is used to infer the updated transition front and the process is successively repeated until convergence of the solution. Within three iterations, the predicted fronts for angles of attack of 1.45, 1.98, 2.46 and 2.94 degrees and a mean-aerodynamic-chord Reynolds number equal to 15 million, approach visual convergence in most regions of the studied cases, and the resulting predictions are in good agreement with the transition fronts deduced from measurements of temperature-sensitive paint. Even though surface pressure measurements based on fully-turbulent flow agree well with the measured pressure coefficient distributions, strong viscous-inviscid interaction effects cause significant shifts in the shock locations based on the imposed transition front, underscoring the intrusive nature of static pressure measurements using surface mounted ports on the CRM-NLF configuration.

Boundary Layer Transition↗
NASA NTRS1988

Drag measurements on a laminar flow body of revolution in Langley's 13 inch magnetic suspension and balance system

Low-speed wind tunnel drag force measurements were taken on a laminar flow body of revolution free of support interference. This body was tested at zero incidence in the NASA Langley 13 inch Magnetic Suspension and Balance System (MSBS). The primary objective of these tests was to substantiate the drag force measuring capabilities of the 13 inch MSBS. A secondary objective was to obtain support interference free drag measurements on an axisymmetric body of interest. Both objectives were met. The drag force calibrations and wind-on repeatability data provide a means of assessing the drag force measuring capabilities of the 13 inch MSBS. The measured drag coefficients for this body are of interest to researchers actively involved in designing minimum drag fuselage shapes. Additional investigations included: the effects of fixing transition; the effects of fins installed in the tail; surface flow visualizations using both liquid crystals and oil flow; and base pressure measurements using a one-channel telemetry system. Two drag prediction codes were used to assess their usefulness in estimating overall body drag. These theoretical results did not compare well with the measured values because of the following: incorrect or non-existent modeling of a laminar separation bubble on the body and incorrect of non-existent estimates of base pressure drag.

Dress, David A.↗
NASA NTRS1985

Transition in a disturbed environment

The title of this presentation is the title of our research grant. While transition study is the objective of the work, the results to date are principally on the properties of turbulent boundary layers at low Reynolds numbers are discussed. Testing was done in a small return wind tunnel. Mean boundary layer development is given. The skin friction behavior of the turbulent points are considered. No standard laminar flow was observed. Furthermore, the turbulent mean flow data seem reasonable for the elevated disturbance levels of the tests in the sense that there is no discernible wake component to any of the profiles and that the variation of skin friction with R sub theta is consistent with zero wake strength. The no-grid data are in all likelihood transitional.

Reshotko, E.↗
NASA NTRS1992

Volume 2: Explicit, multistage upwind schemes for Euler and Navier-Stokes equations

The objective of this study was to develop a high-resolution-explicit-multi-block numerical algorithm, suitable for efficient computation of the three-dimensional, time-dependent Euler and Navier-Stokes equations. The resulting algorithm has employed a finite volume approach, using monotonic upstream schemes for conservation laws (MUSCL)-type differencing to obtain state variables at cell interface. Variable interpolations were written in the k-scheme formulation. Inviscid fluxes were calculated via Roe's flux-difference splitting, and van Leer's flux-vector splitting techniques, which are considered state of the art. The viscous terms were discretized using a second-order, central-difference operator. Two classes of explicit time integration has been investigated for solving the compressible inviscid/viscous flow problems--two-state predictor-corrector schemes, and multistage time-stepping schemes. The coefficients of the multistage time-stepping schemes have been modified successfully to achieve better performance with upwind differencing. A technique was developed to optimize the coefficients for good high-frequency damping at relatively high CFL numbers. Local time-stepping, implicit residual smoothing, and multigrid procedure were added to the explicit time stepping scheme to accelerate convergence to steady-state. The developed algorithm was implemented successfully in a multi-block code, which provides complete topological and geometric flexibility. The only requirement is C degree continuity of the grid across the block interface. The algorithm has been validated on a diverse set of three-dimensional test cases of increasing complexity. The cases studied were: (1) supersonic corner flow; (2) supersonic plume flow; (3) laminar and turbulent flow over a flat plate; (4) transonic flow over an ONERA M6 wing; and (5) unsteady flow of a compressible jet impinging on a ground plane (with and without cross flow). The emphasis of the test cases was validation of code, and assessment of performance, as well as demonstration of flexibility.

Elmiligui, Alaa↗
DOE OSTI2024

Overview of On-Line Optical Measurements at High Pressure for Flue Gases, Particulates and Acid-dew Point of Pressurized Oxy-Combustion

Optical flow cells are critical measurement interfaces, yet sampling under harsh conditions — high pressure, high temperature, particles, moisture, or corrosive gases — makes it difficult to maintain optical quality without perturbing the measurement. To address this challenge, a new flow cell was developed using a laminar coaxial flow field that separates the purge and sample flows. A dedicated test system was built to evaluate particle size distribution (PSD) measurements using a Malvern Panalytical Insitec analyzer. Results demonstrated that the sample flow alone defines the measurement zone, while the purge flow effectively shields the optical windows from deposition, eliminating sampling bias. The flow cell enables reliable PSD measurement under high pressure and temperature in moist, corrosive environments. As a key demonstration, the instrument was successfully deployed for on-line PSD measurement of flue gas from a 100 kWth pressurized oxy-coal combustor at 15 bara.

Cheng, Mao↗