Low-density Wind Tunnel Research Articles

High-Sensitivity Atomic Layer Thermopile Heat- Flux Sensor and Its Application in Hypersonic Low-Density Wind Tunnel Tests

High-Sensitivity Atomic Layer Thermopile Heat- Flux Sensor and Its Application in Hypersonic Low-Density Wind Tunnel Tests

Hypersonic aerodynamic force balance using temperature compensated semiconductor strain gauges

Metal foil strain gauges remain the state-of-the-art transducers for wind tunnel balances. While strain gauge technology is very mature, piezoresistive semiconductor sensors offer alternatives that are worth exploring to assess their unique benefits, such as better strain resolution and accuracy, which would enable balances to be designed with higher factors to safety and hence longer fatigue lifetimes. A new three-component balance, based on temperature compensated semiconductor strain gauges, is designed, calibrated and tested in a hypersonic low density wind tunnel. The static accuracy of the semiconductor balance is calibrated better than 0.3% FS, and the dynamic accuracy of the balance is established using a HB-2 standard model in a Mach 12 hypersonic flow. Good experimental repeatability is confirmed to be better than 2.5% FS, and the effectiveness of the balance is demonstrated by comparing the forces and moments of measured data with computational fluid dynamics simulations, as well as reference wind tunnel results under similar conditions.

Experimental measurement of the pressure drop and heat transfer characteristics of a crossflow heat exchanger in low-density flows

NASA and other space agencies are currently planning the first human missions to Mars in the next 20 years. These missions will require much more power than current robotic missions, and so will require significant waste heat rejection capabilities to maintain steady state operations. Forced convection heat exchangers, particularly crossflow tube array compact heat exchangers, may offer significant volume- and mass-efficiency benefits over radiators for these applications. However, no experimentally validated correlations exist in the literature for the pressure drop and heat transfer characteristics of a forced-convection heat exchanger under the relevant dimensionless conditions which correspond to low (7-200) Reynolds number and moderate (0.0009-0.009) Knudsen number under axial and vortical flows. A staggered-tube crossflow heat exchanger has been constructed and its performance experimentally measured in a low-density wind tunnel. Data has been gathered in low-Reynolds, moderate-Knudsen conditions for both axial and vortical flows over a range of pressures and wall temperatures that is relevant to future Mars or high-altitude heat exchanger operation using both CO2 and air. The measured pressure drop exceeds what is predicted by correlations in the literature at very low (<75) Reynolds number while the Nusselt numbers were fairly well-predicted by existing correlations. A new correlation for pressure drop is presented for low-Reynolds, axial and vortical flow through a staggered circular tube array in crossflow.

Study on Numerical Algorithm of the N-S Equation for Multi-Body Flows around Irregular Disintegrations in Near Space

There has been a concern that the accurate numerical simulation of multi-body flow, which is caused by the multiple disintegrations of expired spacecraft re-entering into the near space, has a critical bottleneck impact on the falling area of the disintegrated debris. To solve this problem, an O-type grid topology method has been designed for the multi-body flow field of irregular debris formed by multiple disintegrations in near space, and a finite-volume implicit numerical scheme has been constructed for the Navier-Stokes equations to solve the aerodynamic interference characteristics of irregular multi-body flow, and further the N-S equation numerical algorithm has been established for the irregular multi-body flows in near space. The reliability of the method has been verified by the comparison of the present computation and the experiment of the low-density wind tunnel for the two-body flow of sphere, cylinder and square scripts. The objects of this study are from the multiple disintegrations of the Tiangong-1 spacecraft during uncontrolled re-entry into the atmosphere, including propelling cylinders and low-temperature lock cabinets. A series of simulations of multi-body flow mechanisms around different combinations have been carried out with varied shapes and spacing. As a result, it is found that when the distance of irregular debris (e.g., two propelling cylinders) in the near space is in the range of Δy &lt; 3D or Δx &lt; D, there is an obvious multi-body interference between debris, and the flow characteristics are obviously changed. When the distance between the debris in near space reaches a certain level, the influence of mutual interference can be ignored. For example, when the y-direction distance between multiple bodies is greater than 3D, the flow interference tends to be small and can be ignored, and we can regard them as two separate pieces to be carried out by the numerical prediction of flight track and falling area in engineering application. The results provide a practical design criterion for the integrated simulation platform which is used to simulate the multi-physics complex aerodynamics of space vehicles from the free-molecule flow of the outer space to the near-ground continuum flow.

Enhancement of signal-to-noise ratio of schlieren visualization measurements in low-density wind tunnel tests using modal decomposition

Enhancement of signal-to-noise ratio of schlieren visualization measurements in low-density wind tunnel tests using modal decomposition

Investigation of Mach number and Reynolds number effects on flow around a near-critical geometry prism at Reynolds number &lt;i&gt;O&lt;/i&gt;(10&lt;sup&gt;3&lt;/sup&gt;) using a low-density wind tunnel

In this study, steady drag force measurements and time-resolved schlieren visualization were conducted for subsonic compressible flows around a rectangular cylinder with a side length ratio from 0.5 to 1.2. The effect of the compressibility on the critical geometry was evaluated based on the drag coefficient using a low-density wind tunnel in the Mach number range from 0.1 to 0.6 at the Reynolds number of O(103). In addition, the relationship between the change in the drag coefficient and the flow field was discussed. The side length ratio that the drag coefficient takes the maximum was confirmed to change with the Mach numbers. Remarkably, the peak drag coefficient at the Mach number of 0.6 appears at the side length ratio of 1.0, even though the side length ratio of the critical geometry at the incompressible flow is approximately 0.65. However, the sharp peak in the drag coefficient which is observed in the previous incompressible studies could not be reproduced even at the Mach number of 0.2. This is considered to be due to the lack of the spanwise length of the model or the Reynolds number effects. The schlieren images showed the influence of the Mach number on the location of the vortex formation, and the location moves downstream as the Mach number increases. The change in the flow fields behind rectangular cylinders causes a change in the back pressure, and it might result in a sudden change in the drag coefficient.

High frequency AC electric glow discharge visualization technology and application in big diameter hypersonic low-density wind tunnel

In order to visualize the rarefied flow field in a Φ1m hypersonic low-density wind tunnel, an electric glow discharge technique based on high frequency excited power has been developed. Firstly, finite element simulation analysis has been carried out, and it is concluded that the breakdown voltage can be reduced by using high frequency power supply; then an electric glow discharge apparatus has been fabricated, and the discharge images were compared with simulations; besides, a clear flow field around HB-1 standard model has been observed using this technology; finally, the influence of glow discharge on the flow field, as well as the differences between direct current and high frequency electric glow discharges is discussed at the end of this paper.

Experimental investigation on compressible flow over a circular cylinder at Reynolds number of between 1000 and 5000

In the present study, a compressible low-Reynolds-number flow over a circular cylinder was investigated using a low-density wind tunnel with time-resolved schlieren visualizations and pressure and force measurements. The Reynolds number ( ) based on freestream quantities and the diameter of a circular cylinder was set to be between 1000 and 5000, and the freestream Mach number ( ) between 0.1 and 0.5. As a result, we have clarified the effect of on the aerodynamic characteristics of flow over a circular cylinder at . The results of the schlieren visualization showed that the trend of effect on the flow field, that are the release location of the Karman vortices, the Strouhal number of vortex shedding and the maximum width of the recirculation, is changed at approximately . In addition, the spanwise phase difference of the surface pressure fluctuation was captured by the measurement using pressure-sensitive paint at approximately of higher- cases. The observed spanwise phase difference is considered to relate to the spanwise phase difference of the vortex shedding due to the oblique instability wave on the separated shear layer caused by the compressibility effects. The Strouhal number of the vortex shedding is influenced by and , and those effects are nonlinear. However, the effects of and can approximately be characterized by the maximum width of the recirculation. In addition, the effect on the drag coefficient can be characterized by the maximum width of the recirculation region and the Prandtl–Glauert transformation.

DSMC modeling of rarefied ionization reactions and applications to hypervelocity spacecraft reentry flows

The DSMC modeling is developed to simulate three-dimensional (3D) rarefied ionization flows and numerically forecast the communication blackout around spacecraft during hypervelocity reentry. A new weighting factor scheme for rare species is introduced, whose key point is to modify the corresponding chemical reaction coefficients involving electrons, meanwhile reproduce the rare species in resultants and preserve/delete common species in reactants according to the weighting factors. The resulting DSMC method is highly efficient in simulating weakly inhomogeneous flows including the Couette shear flow and controlling statistical fluctuation with high resolution. The accurate reliability of the present DSMC modeling is also validated by the comparison with a series of experimental measurements of the Shenzhou reentry capsule tested in a low-density wind tunnel from the HAI of CARDC. The obtained electron number density distribution for the RAM-C II vehicle agrees well with the flight experiment data, while the electron density contours for the Stardust hypervelocity reentry match the reference data completely. In addition, the present 3D DSMC algorithm can capture distribution of the electron, N+ and O+ number densities better than the axis-symmetric DSMC model. The introduction of rare species weighting factor scheme can significantly improve the smoothness of the number density contours of rare species, especially for that of electron in weak ionization case, while it has negligible effect on the macroscopic flow parameters. The ionization characteristics of the Chinese lunar capsule reentry process are numerically analyzed and forecasted in the rarefied transitional flow regime at the flying altitudes between 80 and 97 km, and the simulations predict communication blackout altitudes which are in good agreement with the actual reentry flight data. For the spacecraft reentry with hypervelocity larger than the second cosmic speed, it is forecasted and verified by the present DSMC modeling that ionization reactions will cover the windward capsule surface, leading to reentry communication blackout, and the communication interruption must be considered in the communication design during reentry in rarefied flow regimes.

1025 蛍光油膜法による低レイノルズ数領域の剥離泡長さの計測

This paper describes measurement results of leading-edge separation-bubble length on a flat plate at low Reynolds numbers applying a luminescent oil film technique (GLOF), which allows us to calculate a quantitative reattachment point based on the skin friction vector. The present results are compared with previous experimental data obtained by a water channel and a low-density wind tunnel. A common trend that the separation-bubble length increases and gradually decreases with Reynolds number increasing is observed. However, the results have differences in the peak Reynolds number and separation bubble length even between wind tunnel data.

Pressure-sensitive paint technique for surface pressure measurements in a low-density wind tunnel.

A low-density wind tunnel called the Mars wind tunnel, has been developed at Tohoku University that can produce a high subsonic flow at low pressures for aerodynamic measurements of low-Reynolds-number aircraft wings aimed at developing aircraft applicable to the atmosphere on the planet Mars. Accurate surface pressure measurements on the wing are essential for analysis of not only aerodynamic performance, including lift and drag, but also the flow fields around the wing. This paper presents a surface pressure measurement technique using pressure-sensitive paint (PSP) applicable for Mars wind tunnel tests under low-pressure conditions. The results show that a PSP composed of palladium tetra(pentafluorophenyl) porphyrin (PdTFPP) and poly[1-(trimethylsilyl)-propyne] [poly(TMSP)] exhibits a high-pressure sensitivity at pressures as low as 1 kPa, and the absolute values of the static pressures measured by the PSP accorded well with the values derived from static pressure sensors used as a reference. A calibration methodology for the non-uniform pressure sensitivity on the test model, including a temperature calibration, is also established. The PSP technique clearly demonstrated pressure sensitivity over a distinctive low-pressure region inside a leading edge separation bubble on a flat plate at low Reynolds numbers.Graphical

Coupled Navier–Stokes/Direct Simulation Monte Carlo Simulation of Multicomponent Mixture Plume Flows

A coupled particle-continuum simulation method is presented to compute the multicomponent mixture plume flow from the attitude-control engines of a satellite through the use of the Navier–Stokes solver and direct simulation Monte Carlo method. In this study, the implicit hybrid flux iteration scheme for the axisymmetric compressible Navier–Stokes equations is constructed to solve the inner flowfield of the nozzle. The Navier–Stokes computing technique is implemented considering slip boundary theory for the near-continuum slip flow near the nozzle exit. Direct simulation Monte Carlo methods for the axisymmetric core plume and three-dimensional far-field plume flows of multispecies mixture gas are developed by using a radial weighted factor, by tracking the molecular motion trajectory, and with secondary Cartesian and unstructured cell processing techniques. A multiregion decomposition and hybrid Navier–Stokes/direct simulation Monte Carlo algorithm with two-way and one-way coupling is established to solve the internal and external flow of the thruster, including the near-field, far-field, backflow, and gas-surface contaminated regions. After constructing the coupled Navier–Stokes/DSMC simulation scheme, the present method is applied to solve the plume flowfield and impinging contamination effects from the satellite attitude-control engine and solar array panel; the simulation results correspond well with the experimental measurements from the low-density wind tunnel and the theoretical predictions. To study the contamination effects produced by the multicomponent mixture plume, the current method is employed to simulate a five-component mixture plume flowfield of tens of meters from two representative attitude-control engines installed in different locations of a satellite in orbit. The results of the nonreacting multicomponent flow with equivalent mole fractions can be achieved with a one-component gas simulation if the species is assigned the properties of the mixture. The deposition rate of the plume flows produced by the two teamwork engines can be superimposed on each other. This method can be used to efficiently predict the impingement contamination effects of the gas-fired mixture plume flow on the satellite and solar array panels.

Research on vacuum plume and its effects

In vacuum environment, the exhaust flow of attitude control thrusters would expand freely and produce the plume, which possibly causes undesirable contamination, aerodynamic force and heating effects to the spacecraft. Plume work station (PWS) is developed by Beihang University (BUAA) for numerically simulating the vacuum plume and its effects. An approach which combines the direct simulation Monte Carlo (DSMC) method and difference solution of Navier–Stokes (N–S) equations is applied. The internal flows in nozzles are simulated by solving the NS equations. The flow parameters at nozzle exit are used as the inlet boundary condition for the DSMC calculation. Experimental studies are carried out in a supersonic low density wind tunnel which could simulate the 60–80km altitude environment to investigate the plume and its effects. To demonstrate the capability of PWS, numerical simulations are performed for the vacuum plume of several typical attitude control thrusters. The research results are of great help for the engineering design.

Development of a Low-Density Wind Tunnel for Simulating Martian Atmospheric Flight

A new type of low-density wind tunnel called the “Mars Wind Tunnel (MWT)” has been developed to simulate airfoil flow in the Mars atmosphere. This paper presents the design features of the MWT and its construction. The results of operational performance tests including the driving performance of a multiple supersonic ejector and calibration tests for evaluating the flow characteristics in the test section are shown. These tests were conducted using air as the working gas. It was found that the ejector is an effective means to induce high subsonic flow even under the condition of a low Reynolds number. The pressure ratio achievable with the ejector is inversely proportional to the total pressure. The MWT can simulate flow in the Reynolds number range from 103 to 105 and Mach number up to 0.74. It is confirmed that the MWT has a flow quality and operational characteristics that are suitable for aerodynamic testing of Mars airplanes.

A wind tunnel for the calibration of Mars wind sensors

A major limitation in the development of wind sensors for use on Mars is the lack of suitable testing and calibration facilities. A low-density wind tunnel has been developed at Oxford University for calibration of wind sensors for Mars landers, capable of providing stable or dynamically varying winds, of air or carbon dioxide, at Martian pressures (5–10 mbar) and speeds (0.5–30 m/s), and temperatures of 200–300 K. The flow field in the test section was calculated using analytical and computational modelling techniques, and validated experimentally using a pitot probe. This facility's stability and accuracy offer significant advantages with respect to previous calibration facilities.

Interaction between Shock Wave and Boundary Layer in Nonequilibrium Hypersonic Rarefied Flow

An experimental study of the interaction between a shock wave and a boundary layer over a flat plate with a sharp leading edge in hypersonic rarefied gas flow is presented. Experiments in a low-density wind tunnel using an electron beam probe were conducted at the Shock Wave Laboratory, RWTH Aachen, Germany. Rotational temperatures for stagnation temperatures of T0=670∼1000 K and Kn=0.024∼0.028 based on a reference length of 0.05m were calculated using Muntz’s method and Robben and Talbot’s method. The domain of quasi two-dimensional flow over the plate was determined from three-dimensional rotational temperature measurements. Nonequilibrium between translational and rotational temperatures was observed near the leading edge, and the experimental rotational relaxation length explains the rotational collision number of 2∼4.

Experimental and Numerical Study of Hypersonic Rarefied Gas Flow over Flat Plates

An experimental study of the hypersonic rarefied gas flow over a flat plate with a sharp leading edge is presented. Experiments in a low-density wind tunnel using an electron beam probe were conducted at the Shock Wave Laboratory, RWTH Aachen, Germany. Rotational temperatures for stagnation temperatures of To = 670 ∼ 1000 K and Kn = 0.024 ∼ 0.028 based on a reference length of 0.05 m were calculated by Muntz's method and by Robben and Talbot's method. The domain of quasi-two-dimensional flow over the plate was determined from three-dimensional rotational temperature measurements. A direct simulation Monte Carlo (DSMC) method is applied to obtain a nonequilibrium state for the rarefied gas flow over the flat plate.

Interaction between a Shock Wave and a Boundary Layer in a Nonequilibrium Hypersonic Rarefied Flow (5th Report, Nonequilibrium of Rotational Temperature in a Flow over a Plate)

An experimental study of the interaction between shock wave and boundary layer over a flat plate with a sharp leading edge in hypersonic rarefied gas flow is presented. Experiments in a low density wind tunnel using an electron beam probe were conducted at the Shock Wave Laboratory, RWTH Aachen, Germany. Rotational temperatures for stagnation temperatures of T0=670-1000K and Kn=0.024-0.028 based on a reference length of 0.05m were calculated by Muntz's method and by Robben and Talbot's method. The domain of quasi two-dimensional flow over the plate was determined from three-dimensional rotational temperature measurements. Nonequilibrium between translational and rotational temperatures was observed near the leading edge, and the experimental rotational relaxation length explains the rotational collision number of 2-4.

Interaction between a Shock Wave and a Boundary Layer in a Nonequilibrium Hypersonic Rarefied Flow (4th Report, Nozzle Design for Low Density Wind Tunnel)

Interaction between a Shock Wave and a Boundary Layer in a Nonequilibrium Hypersonic Rarefied Flow (4th Report, Nozzle Design for Low Density Wind Tunnel)

Determination of accommodation coefficients of translational and internal energy using a thin wire in a free-molecular flow

A new experimental–computational method for determination of translational and internal energy accommodation coefficients on a thin wire in a free-molecular flow is elaborated. The method is based on numerical solution of the heat balance equation of the wire. Measurements were performed in a low-density wind tunnel with Ar, He, N2, CH4, and CO2. In addition, the proposed method provides determination of thermophysical properties of the wire.