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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

NASA Langley laminar flow control airfoil experiment

The design and construction of an advanced swept supercritical airfoil for commercial aircraft to be tested in a transonic wind tunnel is described. The swept LFC airfoil was designed for a given thickness ratio and lift coefficient, with emphasis placed on high critical Mach number with shock-free flow. It is compatible with satisfactory low speed and buffeting characteristics and minimizing the suction laminarization. Further emphasis was placed on achieving shock-free flow over a wide range of off-design conditions including trailing edge flap control. The requirements and design of the suction system and modifications to the Langley 8 foot transonic pressure tunnel is briefly described. Contouring of nonporous test section walls for free air simulation and flow quality improvements is included.

Harvey, W. D.↗
NASA NTRS1978

Summary report of the second wind tunnel test of the Boeing LFC model

An 8-ft span, 20-ft chord, 30 deg swept wing section having provisions for laminar boundary control over the first 30% of the upper surface and the first 15% of the lower surface was tested in a 5-ft by 8-ft wind tunnel to explore the sensitivity of laminar flow to various forms of disturbances such as surface imperfections, contamination, off-design pressure distribution (increased crossflow), and imposed noise. The test equipment used and instrumentation of the model are described. Typical results obtained from configurations with spanwise ridges and spanwise rows of disks are discussed as well as suction flow characteristics at reduced incidence.

George-Falvy, D.↗
DOE OSTI2020

How equation of state selection impacts accuracy near the critical point: Forced convection supercritical CO 2 flow over a cylinder

Increasing application of supercritical fluids leads to increased use of numerical simulations to guide designs. However, the impact of Equations of State (EoS) selection on computational cost and accuracy is not well understood. Laminar flow around a heated cylinder is used as a test case. A robust matrix of 99-simulations is used to analyze the flow of carbon dioxide just above the critical point at the isobaric condition of 8 MPa and temperatures between 305 and 390 K. Four EoS were studied: Ideal Gas Law, cubic EoS’s Peng-Robinson and Soave-Redlich-Kwong, and the Span-Wagner EoS. Results were identical for isothermal cases, but for heated or cooled cases, the high-fidelity Span-Wagner EoS simulations presented drastically different wake structures. Compared to the high-fidelity SW EoS, cEoS’s underestimated Nu values by up to 38% and cEoS’s underestimated CD values by 18–42%. In conclusion, the Span-Wagner EoS required only 17% more computation time then the cubic EoS simulations, making the Span-Wagner EoS recommended for engineering design.

42 ENGINEERING↗
DOE OSTI2025

A Novel Optical Instrument for On-Line Measurement of Particle Size Distribution—Application to Clean Coal Technologies

A flow cell is a critical measurement interface for many optical instruments. However, the flows are often sampled under harsh conditions, such as under high pressure and/or high temperature, in the presence of particles, moisture, vapors with high dew points or corrosive gases. Therefore, obtaining a high-optical-quality flow cell that does not perturb the measurement is a significant challenge. To address this challenge, we proposed a new flow cell that employs a unique laminar coaxial flow field (for the purge and sample flows). A test system was built to conduct particle size distribution (PSD) measurements with no sampling bias using a state-of-the-art analyzer (Malvern Panalytical Insitec). The results revealed that the measurement zone is well defined solely by the sample flow, and the optical windows are well protected by the purge flow, with minimal risk of any depositions from the sample flow. Using this flow cell, the Insitec can successfully measure PSD under high pressure and temperature under moist, corrosive conditions without generating any sampling bias. Importantly, we successfully applied this flow cell for on-line PSD measurement for the flue gas of a 100 kWth pressurized oxy-coal combustor operating at 15 bara.

Cheng, Mao (ORCID:0000000213273900)↗
NASA NTRS1981

Advanced Aerodynamics. Selected NASA Research

This Conference Publication contains selected NASA papers that were presented at the Fifth Annual Status Review of the NASA Aircraft Energy Efficiency (ACEE) Energy Efficient Transport (EET) Program held at Dryden Flight Research Center in Edwards, California on September 14 to 15, 1981. These papers describe the status of several NASA in-house research activities in the areas of advanced turboprops, natural laminar flow, oscillating control surfaces, high-Reynolds-number airfoil tests, high-lift technology, and theoretical design techniques.

Source record↗
NASA NTRS1987

Transitional Boundary Layers Under the Influence of High Free Stream Turbulence, Intensive Wall Cooling and High Pressure Gradients in Hot Gas Circulation

The requirements for fundamental experimental studies of the influence of free stream turbulence, pressure gradients and wall cooling are discussed. Under turbine-like free stream conditions, comprehensive tests of transitional boundary layers with laminar, reversing and turbulent flow increments were performed to decouple the effects of the parameters and to determine the effects during mutual interaction.

Rued, Klaus↗
NASA NTRS1986

Laminar Flow Aircraft Certification

Various topics telative to laminar flow aircraft certification are discussed. Boundary layer stability, flaps for laminar flow airfoils, computational wing design studies, manufacturing requirements, windtunnel tests, and flow visualization are among the topics covered.

Williams, Louis J.↗
NASA NTRS1989

Upwind algorithm for the parabolized Navier-Stokes equations

A new upwind algorithm based on Roe's scheme has been developed to solve the two-dimensional parabolized Navier-Stokes equations. This method does not require the addition of user-specified smoothing terms for the capture of discontinuities such as shock waves. Thus, the method is easy to use and can be applied without modification to a wide variety of supersonic flowfields. The advantages and disadvantages of this adaptation are discussed in relation to those of the conventional Beam-Warming (1978) scheme in terms of accuracy, stability, computer time and storage requirements, and programming effort. The new algorithm has been validated by applying it to three laminar test cases, including flat-plate boundary-layer flow, hypersonic flow past a 15-deg compression corner, and hypersonic flow into a converging inlet. The computed results compare well with experiment and show a dramatic improvement in the resolution of flowfield details when compared with results obtained using the conventional Beam-Warming algorithm.

Lawrence, Scott L.↗
NASA NTRS1991

An experimental study of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack

An experimental study of the effects of aft blowing on the forebody vortex asymmetry over a 3.0 caliber tangent ogive body at high angles of attack was conducted. The tip of the ogive body was equipped with a single blowing nozzle whose position could be adjusted. The tests were conducted in a subsonic wind tunnel at laminar flow conditions. The effects of model roll, angle of attack, blowing coefficient, and blowing nozzle axial position were independently studied. Surface pressure measurements and flow visualization results were obtained. Aft blowing was observed to alleviate the degree of vortex asymmetry at all angles of attack. The blowing was found to be more effective at the higher angles of attack. However, proportional control of the degree of vortex asymmetry was not observed, because the initial flowfield was highly asymmetric.

Gittner, Nathan M.↗
NASA NTRS1994

Instabilities orginating from suction holes used for Laminar Flow Control (LFC)

A small-scale wind tunnel previously used for turbulent boundary layer studies has been modified for experiments in laminar flow control. The facility incorporates suction through interchangeable porous test surfaces which are used to stabilize the boundary layer and delay transition to turbulent flow. The thin porous test surfaces are supported by a baffled plenum chamber box which also acts to gather the flow through the surface into tubes which are routed to a high pressure fan. An elliptic leading edge is attached to the assembly to establish a new layer on the test plate. A slot is used to remove the test section flow below the leading edge. The test section was lengthened and fitted with a new ceiling. Substantial modifications were also made to the 3D probe traverse. Detailed studies have been made using isolated holes to explore the underlying instability mechanisms. The suction is perturbed harmonically and data are averaged on the basis of the phase of the disturbance. Conditions corresponding to strong suction and without suction have been studied. In both cases, 3D contour surfaces in the vicinity of the hole show highly three-dimensional T-S waves that fan out away from the hole with streamwise distance. With suction, the perturbations on the centerline are much stronger and decay less rapidly, while the far field is similar to the case without suction. Downstream the contour surfaces of the bow-shaped TS waves develop spanwise irregularities which eventually form into clumps. The contours remain smooth when suction is not applied. Even without suction, the harmonic point source is challenging for CFD; e.g. DNS has been used for streamwise growth. With suction, grid resources are consumed by the hole and this makes DNS even more expensive. Limited DNS results so far indicate that the vortices which emanate from suction holes appear to be stable. The spanwise clumping observed in the experiment is evidence of a secondary instability that could be associated with suction vortices. A typical porous surface for LFC consists of 0.002 inch diameter holes with 0.020 inch grid spacing L, which is too small to resolve disturbances. A 20:1 scale porous test surface has been machined for improved spatial resolution while the L/d is still representative of flight conditions. Designers of porous surfaces use Goldsmith's criterion to minimize crossstream interaction. However nothing is known about the streamwise interactions. Results using two holes, aligned but displaced in the streamwise direction, indicate that partial TS wave cancellation is possible, depending on the hole spacing and disturbance frequency. Using DNS for streamwise interaction studies will be prohibitively expensive if linear superposition cannot be used for the multiple holes.

Watmuff, Jonathan H.↗
NASA NTRS1997

A Numerical Simulation of a Normal Sonic Jet into a Hypersonic Cross-Flow

This study involves numerical modeling of a normal sonic jet injection into a hypersonic cross-flow. The numerical code used for simulation is GASP (General Aerodynamic Simulation Program.) First the numerical predictions are compared with well established solutions for compressible laminar flow. Then comparisons are made with non-injection test case measurements of surface pressure distributions. Good agreement with the measurements is observed. Currently comparisons are underway with the injection case. All the experimental data were generated at the Southampton University Light Piston Isentropic Compression Tube.

Jeffries, Damon K.↗
NASA NTRSDate not supplied

Bluntness and Supersaturation Effects on the Second Mode Instability in Mach 10 Flow

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 31-Inch Mach 10 Tunnel to investigate the impact of nosetip radius and freestream supersaturation on the development of the second mode instability. The model was instrumented with surface-mounted Kulite® and PCB® pressure transducers and thermocouples. Power spectral density plots demonstrate the anticipated trends of the second mode instability weakening and shifting to lower frequencies with increased bluntness. Heat transfer measurements suggest laminar flow over the bulk of the model for most testing conditions, with evidence of transition by the cone base for the sharpest nosetip (R = 0.15 mm) at the higher unit Reynolds numbers. The dynamic surface pressure and heat transfer measurements show that reducing the temperature of the freestream to a supersaturated state has a significant stabilizing effect on the model boundary layer. The surface measurements are observed to be highly dependent on the level of clustering in the freestream. Significant changes in surface pressure spectra due to total temperature reduction are first noted below T0= 948 K, while more substantial fluctuation reductions are found below T0= 810 K.

Hypersonics↗
NASA NTRS1987

Flight survey of the 757 wing noise field and its effects on laminar boundary layer transition. Volume 2: Data compilation

A flight test program was performed using the Boeing 757 flight research airplane to investigate the effect of noise from wing mounted engines on laminar boundary layer transition. An NLF glove was installed on the right wing panel just outboard of the engine. The extent of laminar flow on the glove was measured as a function of engine power setting for a range of flight conditions. A combination of surface and probe microphones was distributed over the upper and lower wing surfaces to measure sound spectra. The flight test program was completed in June 1985 and the results of preliminary analysis indicate that a maximum of about 29 percent of chord laminar flow was obtained on the upper surface and about 28 percent on the lower surface (at a high sideslip condition). The engine speed was varied from about 2600 (idle) to about 4500 (maximum continuous power) r/min. This produced changes in sound pressure level up to 20 dB on the lower surface. On the upper surface, the noise levels were independent of engine power but sensitive to airplane Mach number. No effect of engine power setting on upper surface transition location was observed, and only a small forward movement of the transition location on the lower surface was observed at the high power settings. Volume 1 of this report contains the program description and data analysis. Volume 2 is a compilation of all of the flight test data.

Source record↗
NASA NTRS1985

Heat transfer in pipes

The heat transfer from hot water to a cold copper pipe in laminar and turbulent flow condition is determined. The mean flow through velocity in the pipe, relative test length and initial temperature in the vessel were varied extensively during tests. Measurements confirm Nusselt's theory for large test lengths in laminar range. A new equation is derived for heat transfer for large starting lengths which agrees satisfactorily with measurements for large starting lengths. Test results are compared with the new Prandtl equation for heat transfer and correlated well. Test material for 200- and to 400-diameter test length is represented at four different vessel temperatures.

Burbach, T.↗
NASA NTRS2009

Device for Measuring Low Flow Speed in a Duct

A multiple-throat venturi system has been invented for measuring laminar flow of air or other gas at low speed (1 to 30 cm/s) in a duct while preserving the laminar nature of the flow and keeping the velocity profile across the duct as nearly flat as possible. While means for measuring flows at higher speeds are well established, heretofore, there have been no reliable means for making consistent, accurate measurements in this speed range. In the original application for which this system was invented, the duct leads into the test section of a low-speed wind tunnel wherein uniform, low-speed, laminar flow is required for scientific experiments. The system could also be used to monitor a slow flow of gas in an industrial process like chemical vapor deposition. In the original application, the multiple- throat venturi system is mounted at the inlet end of the duct having a rectangular cross section of 19 by 14 cm, just upstream of an assembly of inlet screens and flow straighteners that help to suppress undesired flow fluctuations (see Figure 1). The basic venturi measurement principle is well established: One measures the difference in pressure between (1) a point just outside the inlet, where the pressure is highest and the kinetic energy lowest; and (2) the narrowest part (the throat) of the venturi passage, where the kinetic energy is highest and the pressure is lowest. Then by use of Bernoulli s equation for the relationship between pressure and kinetic energy, the volumetric flow speed in the duct can be calculated from the pressure difference and the inlet and throat widths. The design of this system represents a compromise among length, pressure recovery, uniformity of flow, and complexity of assembly. Traditionally, venturis are used to measure faster flows in narrower cross sections, with longer upstream and downstream passages to maintain accuracy. The dimensions of the passages of the present venturi system are sized to provide a readily measurable pressure drop. Multiple throats are used to minimize the length needed to recover internal energy and enable the velocity profile to recover to near flatness.

Quinn, Frank↗
NASA NTRS2009

Measuring Thermal Conductivity of a Small Insulation Sample

A multiple-throat venturi system has been invented for measuring laminar flow of air or other gas at low speed (1 to 30 cm/s) in a duct while preserving the laminar nature of the flow and keeping the velocity profile across the duct as nearly flat as possible. While means for measuring flows at higher speeds are well established, heretofore, there have been no reliable means for making consistent, accurate measurements in this speed range. In the original application for which this system was invented, the duct leads into the test section of a low-speed wind tunnel wherein uniform, low-speed, laminar flow is required for scientific experiments. The system could also be used to monitor a slow flow of gas in an industrial process like chemical vapor deposition. In the original application, the multiple- throat venturi system is mounted at the inlet end of the duct having a rectangular cross section of 19 by 14 cm, just upstream of an assembly of inlet screens and flow straighteners that help to suppress undesired flow fluctuations (see Figure 1). The basic venturi measurement principle is well established: One measures the difference in pressure between (1) a point just outside the inlet, where the pressure is highest and the kinetic energy lowest; and (2) the narrowest part (the throat) of the venturi passage, where the kinetic energy is highest and the pressure is lowest. Then by use of Bernoulli s equation for the relationship between pressure and kinetic energy, the volumetric flow speed in the duct can be calculated from the pressure difference and the inlet and throat widths. The design of this system represents a compromise among length, pressure recovery, uniformity of flow, and complexity of assembly. Traditionally, venturis are used to measure faster flows in narrower cross sections, with longer upstream and downstream passages to maintain accuracy. The dimensions of the passages of the present venturi system are sized to provide a readily measurable pressure drop. Multiple throats are used to minimize the length needed to recover internal energy and enable the velocity profile to recover to near flatness.

Miller, Robert A.↗
NASA NTRS1981

Computation of unsteady turbulent boundary layers with flow reversal and evaluation of two separate turbulence models

A procedure which solves the governing boundary layer equations within Keller's box method was developed for calculating unsteady laminar flows with flow reversal. This method is extended to turbulent boundary layers with flow reversal. Test cases are used to investigate the proposition that unsteady turbulent boundary layers also remain free of singularities. Turbulent flow calculations are performed. The governing equations for both models are solved. As in laminar flows, the unsteady turbulent boundary layers are free from singularities, but there is a clear indication of rapid thickening of the boundary layer with increasing flow reversal. Predictions of both turbulence models are the same for all practical purposes.

Cebeci, T.↗
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.↗