Unitary Plan Wind Tunnel Research Articles

Visualization Of Freestream Vortices In The NASA Langley 4-Foot Supersonic Unitary Plan Wind Tunnel

Streamwise vortices present in the freestream of the NASA Langley 4-Foot Supersonic Unitary Plan Wind Tunnel have been observed experimentally for the first time. A high-energy, burst-mode laser system was used to conduct time-resolved planar laser Mie scattering in the freestream of this facility. The freestream of the facility was observed over Mach numbers ranging from 2.4 to 4.6 and (unit) Reynolds number between 3.3 × 106 to 14.8 × 106 1/m. Apparent vortical structures were observed in the tunnel freestream at Mach numbers ≤ 3, with little difference observed for varying the Reynolds number. With further increasing Mach number, protrusive and oscillatory vertical ‘sprite’ structures were observed emanating from the tunnel floor boundary layer. Neither of these structural phenomena has been observed before experimentally in the facility and acts to confirm and inform predictions made in computational fluid dynamic simulations of the facility.

Hybrid Pulse-Burst / Cross-Correlation DGV For High-Speed Flow Measurements At 100 KHz

Hybrid pulse-burst Doppler global velocimetry (PB-DGV)/pulse-burst, cross-correlation DGV (PB-CC-DGV) was demonstrated for the first time in the NASA Langley 4-foot Supersonic Unitary Plan Wind Tunnel, allowing simultaneous acquisition of mean and instantaneous high-repetition-rate velocity fields. The technique was used to make planar velocity measurements across an oblique shockwave generated by a large splitter plate set at a -2° angle of attack in a Mach 2.4 flow. Sequences of >100 consecutive instantaneous velocity images were obtained at a rate of 100 kHz. Velocity fields from both aspects of the hybrid measurement indicated errors less than 1-percent of anticipated velocities. Likewise, all assessments of measurement precision were approximately 1.6-percent of the local velocities. Additional demonstrations in different flowfields including a subsonic axi-symmetric jet and a supersonic Mars reentry vehicle model further illustrate the utility of the new hybrid method.

New Flow-Through Wind-Tunnel Balance for Retropropulsion Testing

Supersonic retropropulsion, or the use of retrorockets starting at supersonic conditions, is an enabling technology for large-scale human Mars lander vehicle concepts. A new flow-through balance was designed and fabricated to support an upcoming retropropulsion test campaign at the NASA Langley Research Center Unitary Plan Wind Tunnel. The balance employs membranes in place of bellows, which have historically been used, and passes the flow-through load-bearing parts of the balance. Flow through the balance is optimized via the addition of flow guides, which are manufactured as separate parts, and attached to the balance. Calibration of this new flow-through balance demonstrates that it can suitably characterize the load and pressure ranges it was designed for, which provides aerodynamicists a new tool to directly measure supersonic retropropulsion performance of powered descent vehicles.

Estimation of Uncertainties for a Model Validation Experiment in a Wind Tunnel

A high-quality model validation experiment was performed in the NASA Langley Research Center Unitary Plan Wind Tunnel to assess the predictive accuracy of computational-fluid-dynamics models for a blunt-body supersonic retropropulsion configuration at freestream Mach numbers from 2.4 to 4.6. Static and fluctuating surface pressure data were acquired on a 5-in.-diam (127-mm-diam) test article with a forebody composed of a spherically blunted, 70 deg half-angle cone and a cylindrical aft body. One unpowered configuration with a smooth outer mold line was tested as well as three powered, forward-firing nozzle configurations: a centerline nozzle, three nozzles equally spaced around the forebody, and a combination with all four nozzles. A key objective of the experiment was the determination of experimental uncertainties from a range of sources such as random measurement variation, flowfield nonuniformity, and model/instrumentation uncertainties. This paper discusses 1) the design of the experiment to best capture these uncertainties for the baseline unpowered configuration, 2) the methodology used in quantifying the various sources of uncertainty, and 3) examples of the uncertainties applied to unpowered and powered experimental results. The uncertainty analysis, which concentrates on the unpowered configuration, showed that flowfield nonuniformity was the dominant contributor to the overall uncertainty. This finding is in agreement with other wind-tunnel experiments that have quantified various sources of uncertainty.

Supersonic Retropropulsion Dynamic Data Analysis from NASA Langley Unitary Plan Wind Tunnel

A supersonic retropropulsion experiment was recently conducted at NASA Langley Research Center’s Unitary Plan Wind Tunnel, test section 2, for a range of Mach numbers from 2.4 to 4.6. A 5-in.-diam 70 deg sphere-cone forebody with a 9.55 in. cylindrical aftbody section was the experimental model used, which is capable of multiple retrorocket configurations. These configurations include a single central nozzle on the center point of the forebody, three nozzles at the forebody half-radius, and a combination of the first two configurations. A series of measurements were achieved through various instrumentation, including forebody and aftbody surface pressures, internal pressures and temperatures, and high-speed schlieren visualization. Several high-speed pressure transducers on the forebody and in the plenum were implemented to look at unsteady flow effects. The following work focuses on analyzing frequency traits due to the unsteady flow for a range of thrust coefficients for single-, tri-, and quadnozzle test cases at a freestream Mach number of 4.6 and an angle of attack ranging from 0 to . This analysis uses MATLAB®’s fast Fourier transform, Welch’s method (modified average of a periodogram), to create a power spectral density and analyze any high-speed pressure transducer frequency traits due to the unsteady flow.

Supersonic Retropropulsion Validation Experiment in the NASA Langley Unitary Plan Wind Tunnel

The development of supersonic retropropulsion, an enabling technology for heavy payload exploration missions to Mars, is the focus of the present paper. A new experimental model, intended to provide computational fluid dynamics model validation data, was recently designed for the Langley Research Center Unitary Plan Wind Tunnel test section 2. Pretest analyses using modern computational fluid and thermal analysis tools were instrumental for sizing and refining the model, over the Mach number range of 2.4–4.6, such that tunnel blockage and internal flow separation issues would be minimized. A 5-in.-diam 70 deg sphere-cone forebody, which accommodates up to four area ratio nozzles, followed by a 9.55-in.-long cylindrical aftbody, was developed for this study based on the computational results. The model was designed to allow for a large number of surface pressure measurements on the forebody and aftbody. Supplemental data included high-speed schlieren video and internal pressures and temperatures. The run matrix was developed to allow for the quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and bias errors due to flowfield or model misalignments. Observations and trends from this initial test in the Unitary Plan Wind Tunnel are presented.

Analysis of Navier–Stokes Codes Applied to Supersonic Retropropulsion Wind-Tunnel Test

Advancement of supersonic retropropulsion as a technology will rely heavily on the ability of computational methods to accurately predict vehicle aerodynamics during atmospheric descent, where supersonic retropropulsion will be employed. A wind-tunnel test at the NASA Langley Unitary Plan Wind Tunnel was specifically designed to aid in the support of Navier–Stokes codes for supersonic retropropulsion applications. Three computational fluid dynamics codes [data parallel line relaxation, fully unstructured Navier–Stokes three-dimensional, and overset grid flow solver] were exercised for multiple nozzle configurations for a range of freestream Mach numbers and nozzle thrust coefficients. The computational fluid dynamics pretest analysis of this wind-tunnel test aided in the test model design process by identifying the potential for tunnel blockage or unstart, of liquefaction within the plume, and of separation occurring at the internal fingers of the nozzles. This analysis led to a reduced model diameter, heating of the plenum, and reducing the nozzle area ratio, and the requirement to radius the corners at the fingers, to counter these potentials, respectively. Comparisons to test data were used to determine the existing capability of the codes to accurately model this complex flow, identify modeling shortcomings, and gain insight into the computational requirements necessary for correctly computing these flows. All three codes predict similar surface pressure coefficients and flowfield structures, such as jet termination shock, interface, bow shocks, and recirculation regions. However, the codes differ on the level of unsteadiness predicted.

Supersonic Retropropulsion Computational Fluid Dynamics Validation with Langley Foot Test Data

Validation of computational fluid dynamics for supersonic retropropulsion is shown through the comparison of three Navier–Stokes solvers and wind-tunnel test results. The test was designed specifically for computational fluid dynamics validation and was conducted in the NASA Langley Research Center supersonic foot Unitary Plan Wind Tunnel. The test includes variations in the number of nozzles, Mach and Reynolds numbers, thrust coefficient, and angles of orientation. Code-to-code and code-to-test comparisons are encouraging, and possible error sources are discussed.

Supersonic Retropropulsion Computational-Fluid-Dynamics Validation with Ames Foot Test Data

A validation study of computational fluid dynamics for supersonic retropropulsion was conducted using three Navier–Stokes flow solvers. The study compared results from the computational-fluid-dynamics codes to each other and to wind-tunnel test data obtained in the NASA Ames Research Center Unitary Plan Wind Tunnel. Comparisons include surface pressure coefficient as well as unsteady plume effects and cover a range of Mach numbers, levels of thrust, and angles of orientation for zero-, one-, three-, and four-nozzle configurations. Flow-structure behavior changed with thrust and angle of orientation for all nozzle configurations. In general, the solvers compared best with the test data for the steadier cases of the one-nozzle and high-thrust three-nozzle configurations. Deviation in surface pressure was noted for the more unsteady cases and near transitions in behavioral modes. Strengths and weaknesses of the solvers are identified, and possible error sources are discussed.

Supersonic Retropropulsion Experimental Results from NASA Ames Foot Supersonic Wind Tunnel

Supersonic retropropulsion was experimentally examined in the Ames Research Center Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5, and 4.6, was a 5-in.-diam 70 deg sphere-cone forebody with a 9.55-in.-long cylindrical aftbody. The forebody was designed to accommodate up to four area ratio nozzles, one on the model centerline and the other three on the half-radius spaced 120 deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20 deg. Results and observations from the test are provided.

Application of a Reynolds-Averaged Navier–Stokes Approach to Supersonic Retropropulsion Flowfields

Systems analysis efforts have identified supersonic retropropulsion as a candidate decelerator technology for human exploration at Mars. These efforts are presently challenged by a lack of available models, including computational fluid dynamics analyses for databases representing the aerodynamic–propulsive interactions inherent to supersonic retropropulsion. This work presents an unsteady Reynolds-averaged Navier–Stokes approach to predict the flowfield structure, surface pressure distributions, and integrated aerodynamic force coefficients for four configurations recently tested in the NASA Langley Research Center Unitary Plan Wind Tunnel. These configurations have zero, one, three, and four nozzles on the forebody. Comparisons were made with experimental data for pressure distributions on the forebody and aftbody, and computational schlieren images illustrating the resulting flowfields were also generated. The results of this work illustrate the applicability of the Reynolds-averaged Navier–Stokes equations to this problem through comparison with data from a test series designed explicitly for the validation of computational fluid dynamics tools in simulating supersonic retropropulsion flowfields. The approach performed well in predicting the surface pressure distribution and flowfield structure for supersonic retropropulsion configurations with different numbers of nozzles at angles of attack of 0 and 8 deg and a thrust coefficient of approximately 2.0.

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

Throughoutthreefulldesignanalysiscycles, theAres Iproject withintheConstellationprogram has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley’s Unitary Plan WindTunnel (UPWT)at the overlapping conditionof Mach=1.6. Muchinsight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach-number range of 2.3-4.5, at a freestream Reynolds number of 1 x 10 6 based on body diameter. All configurations were run at angles of attack from -5 to 20 deg and angles of sideslip of -5 to 5 deg. These results were complemented with computational fluid dynamic predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model computational solutions compared well with experimental results for the baseline and three shelf/tab configurations. Over the range tested, Mach-number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from three different shelf/tab configurations, a fixed-control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

Measured and theoretical supersonic dynamic stability characteristics of a National Aero-Space Plane configuration

Wind-tunnel tests of a National Aero-Space Plane configuration were conducted in the NASA Langley Unitary Plan Wind Tunnel (UPWT). The model used is a Langley designed blended body configuration. Dynamic stability characteristics were measured on this configuration at Mach numbers of 2.0, 2.5, 3.5, and 4.5. In addition to tests of the baseline configuration, component breakdown tests were conducted. The test results show that the baseline configuration generally has stable damping about all three axes, with only isolated exceptions. In addition, there was generally good agreement between the in-phase dynamic parameters and the corresponding static data which were measured during another series of tests in UPWT. Also included are comparisons of the experimental damping parameters with results from the engineering predictive code aerodynamic preliminary analysis system. These comparisons show good agreement at low angles of attack; however, the comparisons are generally not as good at the higher angles of attack.

Design and experimental verification of an equivalent forebody representation of flowing inlets

A method by which a simple equivalent-faired body can be designed to replace a more complex body with flowing inlets has been demonstrated for supersonic flow. An analytically defined, geometrically simple faired inlet forebody has been designed using a linear potential code to generate flow perturbations equivalent to those produced by a much more complex forebody with inlets. An equivalent forebody wind-tunnel model was fabricated, and a test was conducted in NASA Langley Research Center's Unitary Plan Wind Tunnel. The test Mach number range was 1.6-2.16 for angles of attack of -4-16 deg. Test results show that, for the purposes considered here, the equivalent forebody simulates the original flowfield disturbances to an acceptable degree of accuracy.

Cavity door effects on aerodynamic loads of stores separating from cavities

A wind-tunnel investigation was conducted to determine the effects of cavity doors on the aerodynamic characteristics of compressed-ca rriage store configurations during separation from a shallow box cavity. The tests were conducted in the NASA Langley Unitary Plan Wind Tunnel at freestream Mach numbers of 1.70,2.00, and 2.65 for a constant Reynolds number/ft of 2.00X106. Results are summarized to show the effects of door opening angles and vertical height, folded and unfolded tail fins, and Mach number on the near-field aerodynamic separation characteristics of a single missile-type store with in-line cruciform wings and tail fins.

Computational and experimental investigation of cavity flowfields

This paper presents a computational and experimental investigation of supersonic flow past a cavity in a flat plate. The source of the particular interest in this problem is the ongoing study of the aerodynamic interference effects between a separating store and its bay in the parent body. An upwind relaxation scheme, utilizing flux vector splitting and line-Gauss-Seidel iterations, is used to solve Reynolds-averaged Navier-Stokes equations. Spatial discretizations of this two-dimensional analysis are based on implicit and finite-volume methods. Turbulence is modeled and shocks are captured. The flowfield of the symmetry plane at the half-width is computationally visualized and all flow properties are computed. Experimental tests are conducted in the Langley Unitary Plan Wind Tunnel to measure wall pressures and to capture schlieren photographs. Qualitative as well as quantitative data of computations and experiments agree very well. These two vehicles of investigation are merged to show open, closed and transitional cavity flow behaviors.

Planform effects for low-fineness-ratio multibody configurations at supersonic speeds

An experimental and theoretical investigation of planform effects on a low-fineness ratio multibody configuration has been conducted in NASA Langley Research Center's Unitary Plan Wind Tunnel at Mach numbers 1.6, 1.8, 2.0, and 2.16. Experimental and theoretical values of lift, drag, and pitching moment were obtained on three configurations which varied in outboard panel planform only. The three variations were at 65 deg delta, a 70/66 deg cranked arrow, and a 20 deg trapezoidal planform . The purpose of the study was to determine the effect of wing planform on the supersonic aerodynamics and to evaluate the ability of two existing linearized-theory aerodynamic methods to predict these effects. Experimental data showed that the planforms produced the lift, drag-due-to-lift, and pitching-moment characteristics typically found on single-body configurations. However, the data also showed that planform has a minimal influence on zero-lift drag, which is not typical of single-body configurations. Theoretical aerodynamic analysis indicated that codes based on linearized theory adequately predicted the effect of planform on the supersonic aerodynamics.

Aerodynamic Heating to Corrugation Stiffened Structures in Thick Turbulent Boundary Layers

The results of an experimental program to evaluate heat transfer and pressure distributions on corrugation roughened flat plates in thick turbulent boundary layers are presented. The experimental program consisted of tests in the tunnel wall boundary layers of the Langley Unitary Plan Wind Tunnel (UPWT) and Continuous Flow Hypersonic Tunnel (CFHT) at freestream Mach numbers of 2.5, 3.5, 4.5, and 10.3. Tests in the UPWT were conducted at a freestream Reynolds number/m of 10.8 x 106 and in the CFHT, at Reynolds numbers/m of 1.3 to 5.8 x 106. The test configurations consisted of 50.8 cm x 50.8 cm panels with corrugated beads of two different peak amplitudes, 0.61 cm and 0.29 cm. The angle of the corrugated beads relative to the flow direction was varied between 0° (aligned) and 90° (normal). The measured peak and average heat transfer are analyzed and correlated in terms of the bulk boundary layer, internal boundary layer, and geometric parameters. The data are also compared with similar data for thinner boundary layers, and with previously published correlation techniques.