Hypersonic model testing in a shock tunnel

Time-averaged Stanton number and surface-pressure distributions are reported for the first-stage vane row and the first-stage blade row of the Rocketdyne Space Shuttle Main Engine two-stage fuel-side turbine. These measurements were made at 10, 50, and 90 percent span on both the pressure and suction surfaces of the component. Stanton-number distributions are also reported for the second-stage vane at 50 percent span. A shock tube is used as a short-duration source of heated and pressurized air to which the turbine is subjected. Platinum thin-film gages are used to obtain the heat-flux measurements and miniature silicone-diaphragm pressure transducers are used to obtain the surface pressure measurements. The first-stage vane Stanton number distributions are compared with predictions obtained using a quasi-three dimensional Navier–Stokes solution and a version of STAN5. This same N–S technique was also used to obtain predictions for the first blade and the second vane.

Double cone is a commonly used axial symmetric configuration in numerical and experimental studies to simulate complex flow with shock/shock and shock/boundary layer interactions [1–4]. For hypersonic vehicle thermal protection system design, interactions between the bow shock with the shock waves generated by the wing or control surface and between shock wave and boundary layer must be considered of primary concern because it may cause severe local heating load. Double cones with different cone corners become a good choice for aerothermodynamic researchers to investigate the complex heating distribution caused by different kinds of shock interactions. It is critical to accurately determine the spatially continuous heat flux distributions on double cone model surface for understanding of complex flow mechanisms and for comparison with computational fluid dynamics (CFD) predictions, but it is difficult to get the detailed aerodynamic heating data with traditional point measurement techniques, which use separately located thin film gauges or coaxial thermocouples to get the heat flux value on the point. Temperature-sensitive paint (TSP) technique, a global heating measurement technique, makes acquiring such data possible. In this work, a new developed fast-responding TSP system [5] is used to measure the global heat flux distributions on two double cone models in shock tunnel, and the results are compared with the data from thin film gauges and CFD predictions. The main objective of this experiment is to develop experience in the application of TSP in the shock tunnel testing environment and to evaluate the effectiveness of the different measurement and computation techniques in the complex flow field.

Chapter 11 - Shock Tunnel Studies of High-Enthalpy Ionized Airflows

Transient numerical simulations of fluid and heat flow are performed for a number of heat exchanger segments with cylindrical and ellipsoidal form of tubes in a staggered arrangement. Based on the recorded time distributions of velocity and temperature, time-average values of Reynolds number, drag coefficient, and Stanton number are calculated. The drag coefficient and the Stanton number are smaller for the ellipsoidal tubes than for the cylindrical tubes. With an increasing hydraulic diameter, the difference between the two forms of tubes diminishes. To validate the selected numerical approach, the calculated time-average values are compared with experimental data. The time-average values are further used to construct the drag coefficient and the Stanton number as polynomial functions of Reynolds number and hydraulic diameter. The polynomial functions obtained are to be used as input correlations for a heat exchanger integral model.

Drag and heat flux reduction induced by the pulsed counterflowing jet with different waveforms on a blunt body in supersonic flows

Limited available data suggest a substantial impact of Mach number on the heat transfer from an array of jets impinging on a surface at fixed Reynolds number. Many jet array heat transfer correlations currently in use are based on tests in which the jet Reynolds number was varied by varying the jet Mach number. Hence, this data may be inaccurate for high Mach numbers. Results from the present study are new and innovative because they separate the effects of jet Reynolds number and jet Mach number for the purposes of validating and improving correlations that are currently in use. The present study provides new data on the separate effects of Reynolds number and Mach number for an array of impinging jets in the form of discharge coefficients, local and spatially averaged Nusselt numbers, and local and spatially averaged recovery factors. The data are unique because data are given for impingement jet Mach numbers as high as 0.60 and impingement jet Reynolds numbers as high as 60,000, and because the effects of Reynolds number and Mach number are separated by providing data at constant Reynolds number because the Mach number is varied, and data at constant Mach number because the Reynolds number is varied. As such, the present data are given for experimental conditions not previously examined, which are outside the range of applicability of current correlations.

The unsteady effects in the shock tunnel flow establishment can aid hypersonic inlet starting. An unself-starting inlet may be started in the shock tunnel initial operation. To confirm the hypersonic inlet self-starting ability in shock tunnel, a new test method has been developed. It adopts a light obstacle to choke the inlet flow for unstarting the inlet at the initial stage of the shock tunnel operation. The obstacle can then be fast blown out to determine whether the inlet will be restarted under steady test flow conditions. Shutterlike fast response is required for the flow blockage generated by the obstacle because of the short steady flow duration of the shock tunnel. With the help of the shutter-like obstacle, the inlet could be ”pre-set” in either started or unstarted state prior to the steady test flow. Generic hypersonic inlets with different internal contraction ratios have been developed for inlet starting experiments. Various flow characteristics in the inlet starting process were achieved by applying a high speed schlieren system. The internal contraction ratio limits for inlets self-starting and unstarting were obtained and compared with literature. The effect of leading edge bluntness on the inlet starting characteristics were also examined. It was found that the leading edge bluntness play quite important roles in the inlet starting process.

This paper presents the results of a numerical study of hypersonic interacting flows at flow conditions that include those for which experiments have been conducted in the Calspan-University of Buffalo Research Center (CUBRC) Large Energy National Shock (LENS) tunnel and the ONERA R5Ch low-density wind tunnel. The computations are made with the direct simulation Monte Carlo (DSMC) method of Bird. The focus is on Mach 9.3 to 11.4 flows about flared axisymmetric configurations, both hollow cylinder flares and double cones. The results presented highlight the sensitivity of the calculations to grid resolution, provide results concerning the conditions for incipient separation, and provide information concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Although important flow parameters as Mach number, Reynolds number and total enthalpy can be reproduced in most hypersonic experiments quite well, due to different surface temperature effects in wind tunnel and flight, scaling as well as specific flow properties of shock wave/boundary layer interactions are different. This especially holds for short-duration facilities like, e.g. shock tunnels where due to short running times the models remain more or less at ambient temperature. To overcome this shortcoming, an experimental study has been conducted using a preheatable ramp model with 15° ramp angle. This allowed us to adjust the surfaces to an arbitrary temperature just before the experiment started. Pressure and heat flux measurements clearly showed the effect of varying surface and free stream temperatures. These results are supported by schlieren pictures and infrared measurements. The comparison of the measurements with theoretical and numerical results shows a good agreement. Separation bubble scaling laws proposed by Katzer and Davis have been applied and partially confirmed using the local conditions of the boundary layer at separation.

Particle coagulation in steady turbulent flows: Application to smoke aging

We examined 11 cases when the interplanetary magnetic field (IMF) was intensely northward (> 10 nT) for long durations of time (> 3 hours), to quantitatively determine an upper limit on the efficiency of solar wind energy injection into the magnetosphere. We have specifically selected these large BN events to minimize the effects of magnetic reconnection. Many of these cases occurred during intervals of high‐speed streams associated with coronal mass ejections when viscous interaction effects might be at a maximum. It is found that the typical efficiency of solar wind energy injection into the magnetosphere is 1.0 × 10−3 to 4.0 × 10−3, 100 to 30 times less efficient than during periods of intense southward IMFs. Other energy sinks not included in these numbers are discussed. Estimates of their magnitudes are provided.

The purpose of this investigation was to find the relation between measured static pressure and true static pressure as a function of the viscous interaction parameter χ. The results obtained make possible accurate interpretation of measured static pressure data.

Direct numerical simulations of incompressible turbulent forced convection over irregular, anisotropic surface roughness in a pressure-driven plane channel flow have been performed. Heat transfer was simulated by solving the passive scalar transport equation with Prandtl number Pr = 0.7. The roughness topographies under investigation here are based on an X-ray computed tomography scan of an additively manufactured internal cooling passage, which had an irregular, multi-scale and mildly non-Gaussian height distribution. Three different roughness topographies and three different friction Reynolds numbers (Reτ = 395, 590, 720) were considered, along with reference smooth-wall simulations at matched Reτ. By systematically varying the roughness topography and flow conditions, a direct computational assessment of aero-thermal performance (pressure losses and heat transfer) and the Reynolds analogy factor, i.e. 2Ch/Cf, where Ch is the heat-transfer coefficient (Stanton number) and Cf is the skin-friction coefficient, was conducted. The results highlight the profound impact that the roughness orientation (relative to the flow direction) has upon the aero-thermal performance of additively manufactured internal passages, with transverse-aligned roughness augmenting heat transfer by as much as 33%, relative to its streamwise-aligned counterpart. An interrogation of velocity and temperature statistics in the near-wall region was also performed, which underlined the growing dissimilarity between heat transfer and drag as fully-rough conditions are approached.

Direct numerical simulations of incompressible turbulent forced convection over irregular, anisotropic surface roughness in a pressure-driven plane channel flow have been performed. Heat transfer was simulated by solving the passive scalar transport equation with Prandtl number Pr = 0.7. The roughness topographies under investigation here are based on an X-ray computed tomography scan of an additively manufactured internal cooling passage, which had an irregular, multiscale and mildly non-Gaussian height distribution. Three different roughness topographies and three different friction Reynolds numbers (Reτ = 395, 590, 720) were considered, along with reference smooth-wall simulations at matched Reτ. By systematically varying the roughness topography and flow conditions, a direct computational assessment of aero-thermal performance (pressure losses and heat transfer) and the Reynolds analogy factor, i.e., 2Ch/Cf, where Ch is the heat-transfer coefficient (Stanton number) and Cf is the skin-friction coefficient, was conducted. The results highlight the profound impact that the roughness orientation (relative to the flow direction) has upon the aero-thermal performance of additively manufactured internal passages, with transverse-aligned roughness augmenting heat transfer by as much as 33%, relative to its streamwise-aligned counterpart. An interrogation of velocity and temperature statistics in the near-wall region was also performed, which underlined the growing dissimilarity between heat transfer and drag as fully rough conditions are approached.

Near wakes and cross-flow separation in hypersonic flows have important implications for the effective design and engineering of next-generation flight vehicles, reentry capsules, and propulsion systems. The present work utilizes numerical simulations of a Mach 10 slab delta at a moderate angle of attack in adiabatic, cold wall, and slip wall flow in order to reconcile the cross-flow vortex formation with free-interaction processes observed on primitive shapes and capsules. Validation of the simulation methodology is performed by comparing with experimental data from the literature for pressure and Stanton number distributions. First, the work shows that a boundary layer separation is not required to form a cross-flow recirculation flow pattern. This indicates that the flow reversal leading to the rotating flow is related to gasdynamic phenomena, rather than being a result of boundary layer separation alone. Subsequently, simulation results are used to examine the mechanisms of viscous–inviscid interaction leading to separation. The work demonstrates that despite substantially changing pressure and local skin friction along the delta wing, the separation process obeys a nearly self-similar free-interaction process over a large portion of the body. A free-interaction parameter modified to account for the momentum balance in the direction of the leading edge is presented and demonstrates the predictability of this interaction at various stations along the delta. These results and observations demonstrate that past work on near-wake free interactions is more general than previously shown and confirms and builds on early work in the literature.