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At least 811 records · Page 45
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Molecular flow velocity using Doppler shifted Raman spectroscopy

Measurements of molecular flow velocity, static pressure, and translational temperature in the free-stream of a supersonic wind tunnel and behind the shock of a simple model are reviewed. Based on the free-stream demonstration using inverse Raman spectroscopy, an experiment is outlined to investigate the lee-side flow field above a swept delta wing and simulated spectra expected for the leading-edge vortex are included. The extension of the technique to hypersonic wind tunnels is also explored through the use of simulated spectra.

Exton, Reginald J.↗
NASA NTRS1989

A Mach 6 external nozzle experiment with Argon-Freon exhaust simulation

A scramjet exhaust simulation technique for hypersonic wind tunnel testing has been developed. Mixtures of Argon and Freon correctly match the inviscid simulation parameters of Mach number, static-pressure ratio, and the ratio of specific heats at the combustor exit location; this simulation is accomplished at significantly reduced temperatures and without combustion. An investigation of nozzle parametrics in a Mach 6 freestream showed that the external nozzle ramp angle, the cowl trailing-edge angle, an external nozzle flow fence and the nozzle static-pressure ratio significantly affected the external nozzle thrust and pitching moment as measured by the integration of surface-pressure data. A comparison of Argon-Freon and air exhaust simulation showed that the external nozzle thrust and pitching moment were in error by roughly a factor of 2 using air due to the incorrect match of the ratio of specific heats. An assessment of two-dimensional Euler and Navier-Stokes codes for predicting external nozzle aerodynamic characteristics was made by comparing computed and experimental results.

Pittman, James L.↗
NASA NTRSDate not supplied

Aero-Propulsive Modeling for eVTOL Aircraft Using Wind Tunnel Testing with Multisine Inputs

A novel approach for modeling the aero-propulsive characteristics of an electric vertical takeoff and landing (eVTOL) aircraft was developed and demonstrated in wind tunnel testing. The approach was applied to the NASA LA-8 tandem tilt-wing eVTOL aircraft, using an efficient, hybrid experiment design composed of a static I-optimal response surface design for slowly-varying test variables, and dynamic orthogonal phase-optimized multisine excitations for the control surfaces and electric propulsors. Both the static and dynamic experiment designs were executed simultaneously to collect informative data for model identification. Statistically-weighted multivariate orthogonal function modeling was used to combine local modeling results computed in the frequency domain using data collected with dynamic excitations operating on the control effectors to form an aggregate aero-propulsive model. The final identified model exhibited good predictive capability when compared to validation data acquired separately from the data used to develop the model. The required test time using these new techniques was reduced by at least a factor of five compared to previous static wind tunnel testing for the LA-8 aircraft, while providing more informative data, greater parameterization flexibility, and high-quality models.

system identification↗
NASA NTRS1985

Experiments with solid particle seeding

Kaoline, a hydrated aluminum silicate clay, is investigated as a seeding material for laser velocimetry. It is inexpensive but is polydispersed with some of the fineparticles being too large to follow wind tunnel flow and is in the form of nonspherical platelets having an aspect ratio of approximately 4/1. Gravity sedimentation experiments as a means of narrowing the fineparticle sizes distribution are being conducted. The fineparticle size distribution of Engelhard ASP 200 kaolin suspended in ethanol (0.00792 grams kaolin/ml ethanol) as received, after 24 hours gravity sedimentation and after 48 hours sedimentation, respectively is shown. A shearing atomizer is used to inject the fineparticles. Gravity sedimentation is carried out in an 800 ml pyrex beaker. Following gravity sedimentation, the top 3.5 inches are siphoned from the liquid, which has a column height of 4.5 inches. In a like manner, longer settling times will serve to further narrow the fineparticle distribution range. As successive sedimentations are effected, the number of fineparticles per unit volume of ethanol decreased markedly.

Nichols, C. E., Jr.↗
NASA NTRS1981

Natural Laminar Flow Flight Experiment

A supercritical airfoil section was designed with favorable pressure gradients on both the upper and lower surfaces. Wind tunnel tests were conducted in the Langley 8 Foot Transonic Pressure Tunnel. The outer wing panels of the F-111 TACT airplane were modified to incorporate partial span test gloves having the natural laminar, flow profile. Instrumentation was installed to provide surface pressure data as well as to determine transition location and boundary layer characteristics. The flight experiment encompassed 19 flights conducted with and without transition fixed at several locations for wing leading edge sweep angles which varied from 10 to 26 at Mach numbers from 0.80 to 0.85 and altitudes of 7620 meters and 9144 meters. Preliminary results indicate that a large portion of the test chord experienced laminar flow.

Steers, L. L.↗
NASA NTRS2016

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet Flows with Shock Interactions

NASA and the U.S. aerospace industry are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The computational analyses of modern aircraft designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty remains in the aft signatures due to boundary layer and nozzle exhaust jet effects. Wind tunnel testing without inlet and nozzle exhaust jet effects at lower Reynolds numbers than in-flight make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel is planned for February 2016 to address the nozzle jet effects on sonic boom. The experiment will provide pressure signatures of test articles that replicate waveforms from aircraft wings, tails, and aft fuselage (deck) components after passing through cold nozzle jet plumes. The data will provide a variety of nozzle plume and shock interactions for comparison with computational results. A large number of high-fidelity numerical simulations of a variety of shock generators were evaluated to define a reduced collection of suitable test models. The computational results of the candidate wind tunnel test models as they evolved are summarized, and pre-test computations of the final designs are provided.

Wind Tunnel↗
NASA NTRS1994

Development of a multicomponent force and moment balance for water tunnel applications, volume 1

The principal objective of this research effort was to develop a multicomponent strain gauge balance to measure forces and moments on models tested in flow visualization water tunnels. An internal balance was designed that allows measuring normal and side forces, and pitching, yawing and rolling moments (no axial force). The five-components to applied loads, low interactions between the sections and no hysteresis. Static experiments (which are discussed in this Volume) were conducted in the Eidetics water tunnel with delta wings and a model of the F/A-18. Experiments with the F/A-18 model included a thorough baseline study and investigations of the effect of control surface deflections and of several Forebody Vortex Control (FVC) techniques. Results were compared to wind tunnel data and, in general, the agreement is very satisfactory. The results of the static tests provide confidence that loads can be measured accurately in the water tunnel with a relatively simple multicomponent internal balance. Dynamic experiments were also performed using the balance, and the results are discussed in detail in Volume 2 of this report.

Suarez, Carlos J.↗
NASA NTRS2006

Innovation in Aerodynamic Design Features of Soviet Missiles

Wind tunnel investigations of some tactical and strategic missile systems developed by the former Soviet Union have been included in the basic missile research programs of the NACA/NASA. Studies of the Soviet missiles sometimes revealed innovative design features that resulted in unusual or unexpected aerodynamic characteristics. In some cases these characteristics have been such that the measured performance of the missile exceeds what might have been predicted. In other cases some unusual design features have been found that would alleviate what might otherwise have been a serious aerodynamic problem. In some designs, what has appeared to be a lack of refinement has proven to be a matter of expediency. It is a purpose of this paper to describe some examples of unusual design features of some Soviet missiles and to illustrate the effectiveness of the design features on the aerodynamic behavior of the missile. The paper draws on the experience of the author who for over 60 years was involved in the aerodynamic wind tunnel testing of aircraft and missiles with the NACA/NASA.

Spearman, M. Leroy↗
NASA NTRS1985

Flight flutter testing at Ames-Dryden

Over the past several years the NASA Ames-Dryden Flight Research Facility at Edwards Air Force Base has developed a variety of flight flutter and ground test techniques which have been applied to an assortment of new or modified aerospace research vehicles. This paper presents a summary of these techniques and the experiences gained from these applications. Topics discussed include the roles of ground vibration testing, flight flutter testing, wind tunnel flutter model testing, predictive analyses, and aeroservoelastic considerations. Data are presented for a wide variety of aircraft, including remotely piloted vehicles, modern fighters with relaxed static stability and highly augmented flight control systems, aircraft modified for laminar flow control experiments, a glider modified for deep stall tests, and aircraft with skewed or forward swept wings. The conclusions include a brief discussion of future directions in flight flutter testing at Ames-Dryden.

Felt, L. R.↗
NASA NTRS1975

Control of shock wave-boundary layer interactions by bleed in supersonic mixed compression inlets

An experimental investigation was conducted to determine the effect of bleed on a shock wave-boundary layer interaction in an axisymmetric mixed-compression supersonic inlet. The inlet was designed for a free-stream Mach number of 2.50 with 60-percent supersonic internal area contraction. The experiment was conducted in the NASA Lewis Research Center 10-Foot Supersonic Wind Tunnel. The effects of bleed amount and bleed geometry on the boundary layer after a shock wave-boundary layer interaction were studied. The effect of bleed on the transformed form factor is such that the full realizable reduction is obtained by bleeding of a mass flow equal to about one-half of the incident boundary layer mass flow. More bleeding does not yield further reduction. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise.

Fukuda, M. K.↗
NASA NTRS1975

On the computation of two- and three-dimensional steady transonic flows by relaxation methods

The paper is concerned with the application of the Murman and Cole (1971) relaxation scheme to steady, inviscid transonic flow problems in two and three dimensions. This scheme, which automatically accounts for weak shock waves, uses separate difference operators in elliptic and hyperbolic regions. The details of the scheme are described in terms of the original small disturbance formulation of Murman and Cole. In particular, Murman's recent (1973) introduction of fully conservative difference operators to obtain the correct shock jumps is examined. The extension to treating the exact isentropic equation is then covered with special attention given to Jameson's (to appear) rotated difference scheme for supersonic flow regions. The bulk of the discussion is related to two-dimensional procedures, and some comparisons with experiment are made, with emphasis on the effects of viscosity and wind-tunnel walls. Application of the Murman-Cole scheme is then discussed for small disturbances in three dimensions.

Bailey, F. R.↗
NASA NTRS1986

Failure analysis of a large wind tunnel compressor blade

A failure analysis was performed to establish the cause of a 2014-T6 aluminum compressor rotor blade failure in a large NASA-Ames wind tunnel. Metallurgical failure analysis by light photography, fractographic SEM examinations, and fatigue experiments showed that a 0.13-mm-deep scratch in the shank of the blade had acted as an initiation site for a fatigue crack, which subsequently grew by Stage II fatigue before occurrence of the final fracture by unstable crack growth. Studies of modal deformation plots and blade FEM analysis indicated that Mode-3 was the most likely mode responsible for the resonance condition which produced the vibration loading that effected crack growth. Tunnel measurements data showed that the crack growth was controlled by both the tunnel resonance and the dampening from the crack itself.

Hampton, Roy W.↗
NASA NTRS1987

Aerodynamic design modification of a hypersonic wind tunnel nozzle by CSCM with high order accuracy

An improved version of the CSCM implicit Navier-Stokes solver with a flexible data structure has been applied, in conjunction with a sophisticated patched grid system, to the aerodynamic analysis of hypersonic axisymmetric contoured nozzles of the NASA-Ames 3.5-foot hypersonic wind tunnel. The results obtained are found to be in good agreement with experiment. Based on the analysis, a newly designed throat contour for the Mach 14 nozzle has been recommended.

Yang, J. Y.↗
NASA NTRS1993

Selected computations of transonic cavity flows

An efficient diagonal scheme implemented in an overset mesh framework has permitted the analysis of geometrically complex cavity flows via the Reynolds averaged Navier-Stokes equations. Use of rapid hyperbolic and algebraic grid methods has allowed simple specification of critical turbulent regions with an algebraic turbulence model. Comparisons between numerical and experimental results are made in two dimensions for the following problems: a backward-facing step; a resonating cavity; and two quieted cavity configurations. In three-dimensions the flow about three early concepts of the stratospheric Observatory For Infrared Astronomy (SOFIA) are compared to wind-tunnel data. Shedding frequencies of resolved shear layer structures are compared against experiment for the quieted cavities. The results demonstrate the progress of computational assessment of configuration safety and performance.

Atwood, Christopher A.↗
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 NTRS1994

Receptivity of flat-plate boundary layer in a non-uniform free stream (vorticity normal to the plate)

Recent progress in both the linear and nonlinear aspects of stability theory has highlighted the importance of the receptivity problem. One of the most unclear aspects of receptivity study is the receptivity of boundary-layer flow normal to vortical disturbances. Some experimental and theoretical results permit the proposition that quasi-steady outer-flow vortical disturbances may trigger by-pass transition. In present work such interaction is investigated for vorticity normal to a leading edge. The interest in these types of vortical disturbances arise from theoretical work, where it was shown that small sinusoidal variations of upstream velocity along the spanwise direction can produce significant variations in the boundary-layer profile. In the experimental part of this work, such non-uniform flow was created and the laminar-turbulent transition in this flow was investigated. The experiment was carried out in a low-turbulence direct-flow wind tunnel T-361 at the Central Aerohydrodynamic Institute (TsAGI). The non-uniform flow was produced by laminar or turbulent wakes behind a wire placed normal to the plate upstream of the leading edge. The theoretical part of the work is devoted to studying the unstable disturbance evolution in a boundary layer with strongly non-uniform velocity profiles similar to that produced by outer-flow vorticity. Specifically, the Tollmien-Schlichting wave development in the boundary layer flow with spanwise variations of velocity is investigated.

Kogan, M. N.↗
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↗