Instrumentation Technique for Mixed Hypersonic Flow in Short Duration Facility

Waveriderwhich with its geometry designed for a specific Mach number optimally exploits the lift generated due to shock wavesis a promising aerodynamic configuration for hypersonic flight. A conically derived waverider is designed for a nominal Mach number of 6. Detailed aerodynamic tests on the designed wave rider are initiated. Since hypersonic flows are essentially characterized by high flow total enthalpies, their simulations require impulse facilities like shock tunnels. Experiments are performed in IISc Hypersonic Shock tunnel HST-2 at the design Mach number as well as at an offdesign Mach number of 8, with a flow total enthalpy of 2 MJ/kg. Heat transfer rates are important concern at such high total enthalpies. The surface heat transfer distribution on the waverider at design and off-design conditions, measured using platinum thin film sensors, are presented in this paper. Despite the same flow total enthalpy, higher heat transfer rates are observed at Mach 6 than at Mach 8.

Recent advances in hypersonic research activities have led to the development in ground based test facilities for simulating near-realistic flight conditions in the laboratory. Basic performance data such as drag, lift, thrust etc., are usually obtained from ground-based experimentation for physical understanding of the flow fields and subsequent prototype developments. Hypersonic flows are produced for a very short time in impulse facilities such as shock tunnels, free-piston shock tunnels and expansion tubes. The test times are in the order of few milliseconds in the shock tunnel and for the expansion tubes, it is still less (~50 μs). The force measurements in these facilities are very challenging due the need of fast response sensors and instrumentations. In this review paper, two force measurement techniques are discussed that are best suited for short duration facilities namely; accelerometer balance and stress-wave force balance system.

A three-component accelerometer balance system is used to study the drag reduction effect of an aerodisc on large angle blunt cones flying at hypersonic Mach numbers. Measurements in a hypersonic shock tunnel at a freestream Mach number of 5.75 indicate more than 50% reduction in the drag coefficient for a 120degrees apex angle blunt cone with a forward facing aerospike having a flat faced aerodisc at moderate angles of attack. Enhancement of drag has been observed for higher angles of attack due to the impingement of the flow separation shock on the windward side of the cone. The flowfields around the large angle blunt cone with aerospike assembly flying at hypersonic Mach numbers are also simulated numerically using a commercial CFD code. The pressure and density levels on the model surface, which is under the aerodynamic shadow of the flat disc tipped spike, are found very low and a drag reduction of 64.34% has been deduced numerically.

A program of experimental research and analysis was conducted to examine the heat transfer and pressure distributions in regions of shock/shock interaction over smooth and transpiration-cooled hemispherical noseshapes. The objective of this investigation was to determine whether the large heat transfer generated in regions of shock/shock interaction can be reduced by transpiration cooling. The experimental program was conducted at Mach numbers of 12 to 16 in the Calspan 48-Inch Shock Tunnel. Type 3 and type 4 interaction regions were generated for a range of freestream unit Reynolds numbers to provide shear layer Reynolds numbers from 10 exp 4 to 10 exp 6 to enable laminar and turbulent interaction regions to be studied. Shock/shock interactions were investigated on a smooth hemispherical nosetip and a similar transpiration-cooled nosetip, with the latter configuration being examined for a range of surface blowing rates up to one-third of the freestream mass flux. While the heat transfer measurements on the smooth hemisphere without shock/shock interaction were in good agreement with Fay-Riddell predictions, those on the transpiration-cooled nosetip indicated that its intrinsic roughness caused heating-enhancement factors of over 1.5. In the shock/shock interaction studies on the smooth nosetip, detailed heat transfer and pressure measurements were obtained to map the variation of the distributions with shock-impingement position for a range of type 3 and type 4 interactions. Such sets of measurements were obtained for a range of unit Reynolds numbers and Mach numbers to obtain both laminar and turbulent interactions. The measurements indicated that shear layer transition has a significant influence on the heating rates for the type 4 interaction as well as the anticipated large effects on type 3 interaction heating. In the absence of blowing, the peak heating in the type 3 and type 4 interaction regions, over the transpiration-cooled model, did not appear to be influenced by the model's rough surface characteristics. The studies of the effects of the transpiration cooling on type 3 and type 4 shock/shock interaction regions demonstrated that large surface blowing rates had significant effect on the structure of the flowfield, enlarging the shock layer and moving the region of peak-heating interaction around the body.

This paper presents the results of an experimental study of the unsteady nature of a hypersonic separated turbulent flow. The nomimal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5×107/m. The separated flow was generated using finite span forward facing steps. An array of flush mounted high spatial resolution and fast response platinum film resistance thermometers was used to make multi-channel measurements of the fluctuating surface heat trtansfer within the separated flow. Conditional sampling analysis of the signals shows that the root of separation shock wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer thickness. The compression waves converge into a single leading shock beyond the boundary layer. The shock structure is unsteady and undergoes large-scale motion in the streamwise direction. The length scale of the motion is about 22 percent of the upstream influence length of the separation shock wave. There exists a wide band of frequency of oscillations of the shock system. Most of the frequencies are in the range of 1–3 kHz. The heat transfer fluctuates intermittently between the undisturbed level and the disturbed level within the range of motion of the separation shock wave. This intermittent phenomenon is considered as the consequence of the large-scale shock system oscillations. Downstream of the range of shock wave motion there is a separated region where the flow experiences continuous compression and no intermittency phenomenon is observed.

In this study, the performances of methods of flow separation into the surface flow and base flow of the total stream flow in rivers with different hydraulic features have been examined. For this purpose, daily mean stream flow data of Büyük Melen and Aksu Rivers which are in the same watershed but with different features (average flow value, catchment area, mean elevation) have been separated as surface flow and base flow with the use of Digital Filtering Method (DFM) and United Kingdom Institute of Hydrology Method (UKIH). The recession coefficient in DFM has been used as α = 0.830; and the number of elements for the groups formed to determine the turning points of flow data in UKIH has been used as N = 5. The study has revealed that the results of surface and base flows obtained by both flow separation methods show similarity in all rivers in an acceptable level; and the surface flow values agree better (R2 > 0,76) compared to base flow values (R2 > 0.63). However, it has been seen that as long as the total flow values decrease, the surface flow results for both methods come closer to each other; but the results of base flow get differentiated. This situation has been clearly seen in the results of surface (R2 > 0.89) and base (R2 > 0.63) flows belonging to Aksu River which has lower values of flow when compared to the ones of Büyük Melen River.

An experiment is described that combines in situ measurements made with very fast response sensors on a 300 m tower at the Boulder (Colorado) Atmospheric Observatory (Kaimal and Gaynor, 1983), with remote sensing of atmospheric layer structure by acoustic and FM-CW radar sounders (e.g., Little, 1969; Richter, 1969). A finestructure of laminae is often present within elevated inversion layers that is reminiscent of the sheet and layer structure found in the ocean thermocline and in lakes and estuaries (e.g., Gregg, 1982). Figure 1 shows an acoustic record of such a stepped profile event through which a moving carriage on the tower carried temperature, humidity and wind sensors along time-height trajectories shown by the superimposed staight lines. The temperature measurements were made with platinum wires, humidity was measured with a Lyman-α humidiometer, and wind was measured by three-component sonic anemometers. The_data were then digitized at_a 10 Hz rate, and the carriage rose at 0.56 ms-1 and descended at 0.58 ms-1. Figure 2 shows the resulting profiles of temperature, humidity and wind. The transition through the sheets and layers within the inversion is often extremely abrupt with very large gradients. Near_the center of the lower sheet the local gradient of temperature was 6°C m-1. An experiment is described that combines in situ measurements made with very fast response sensors on a 300 m tower at the Boulder (Colorado) Atmospheric Observatory (Kaimal and Gaynor, 1983), with remote sensing of atmospheric layer structure by acoustic and FM-CW radar sounders (e.g., Little, 1969; Richter, 1969). A finestructure of laminae is often present within elevated inversion layers that is reminiscent of the sheet and layer structure found in the ocean thermocline and in lakes and estuaries (e.g., Gregg, 1982). Figure 1 shows an acoustic record of such a stepped profile event through which a moving carriage on the tower carried temperature, humidity and wind sensors along time-height trajectories shown by the superimposed staight lines. The temperature measurements were made with platinum wires, humidity was measured with a Lyman-α humidiometer, and wind was measured by three-component sonic anemometers. The_data were then digitized at_a 10 Hz rate, and the carriage rose at 0.56 ms-1 and descended at 0.58 ms-1. Figure 2 shows the resulting profiles of temperature, humidity and wind. The transition through the sheets and layers within the inversion is often extremely abrupt with very large gradients. Near_the center of the lower sheet the local gradient of temperature was 6°C m-1.

Flow control effects of three types of nanosecond (NS) dielectric barrier discharge (DBD) plasma actuations in suppressing the flow separation of a compressor airfoil at low Reynolds number are investigated using large-Eddy simulation. It has been found that the NS DBD plasma actuation can effectively suppress the flow separation and two mechanisms behind the flow control effect have been uncovered. First, the NS DBD plasma actuation can induce distorted flow structure (DFS) within the flowfield, and the induced DFS of plasma actuations located upstream of the flow separation zone can bring the Kelvin–Helmholtz instability of the shear layer between mainstream flow and separated flow froward to the separation point. Then, a large-scale spanwise vortex, which promotes the mixing of the main flow and the separated flow, is induced within the flow separation zone, thus resulting in suppression of the flow separation. Second, the plasma actuation located at the blade leading edge can improve the momentum of the laminar boundary layer, owing to the propagations of compressive waves as well as DFS along the blade surface. The plasma actuation located within the separation zone fails to suppress the boundary-layer flow separation effectively because the induced DFS cannot influence the shear layer between mainstream flow and separated flow.

The present experiment is conducted for the hypersonic flow of low Reynolds number and shows the possibility of low density hypersonic flow by using a shock tunnel and electron-beam densitometers. Electron-beam densitometer has been developed for the measurement of flow at the test section of shock tunnel by the author. The induced surface pressure on a flat plate with a sharp leading edge in the low density hypersonic flow is determined from the density measured by the described densitometer in previous papers. In the present paper, the density distribution ahead of a sphere is obtained by appying the scanning of the electron-beam.

Simulation of two-dimensional turbulent forced convection flow adjacent to inclined backward-facing step is presented to examine the effects of step length and inclination angle on flow and heat transfer distributions. The flow is considered to be steady, turbulent and incompressible. The low temperature fluid flows over a horizontal surface with an inclined backward step such that the bottom wall and step are heated with uniform temperature. The continuity, Navier-Stokes and energy equations are solved numerically and the k- model is used for turbulence flow modeling. Because of the complex flow geometry, the governing equations are transformed into the computational domain and the discretized forms of these equations are obtained with finite volume scheme. The relation between physical and computational domains is obtained by conformal mapping technique using the Schwarz-Christoffel transformation. Numerical results provided conclusive evidence that the step length and its inclined angle have considerable effects on the heat transfer and flow distributions. The fluid flow and heat transfer predictions are compared with theoretical results of others and satisfactory agreement is observed.

The results of an experimental investigation of high-enthalpy, hypersonic flow over sharp leading-edge compression corners are presented and discussed. In particular, the possible effects of real gas behavior are examined. Measurements have been made of the heat transfer and pressure distributions for flat plate and compression corner flow. Some flow visualization data have also been obtained. Test flows were generated using a free-piston shock tunnel operating in the reflected mode. The reservoir enthalpy ranged from 3 to 19 MJ kg -1 , giving freestream speeds of 2.3-5.5 km s -1 . For these conditions, the flow remains laminar throughout. The flat plate data for both high- and low-enthalpy flows are in agreement with the reference enthalpy method for heat transfer and the weak interaction theory for pressure. Also, the measured flat plate boundary-layer thickness compares well with an expression strictly valid for perfect gas flows only. The high- and low-enthalpy compression corner flows have upstream influence and plateau pressure behavior similar to perfect gas flow. That is, real gas effects for the present flows appear to be negligible. This is consistent with the essentially chemically frozen viscous and inviscid flow upstream of the interaction.

In this study the centre manifold is applied for reduction and limit cycle calculation of a highly nonlinear structural aeroelastic wing. The limit cycle is arisen from structural nonlinearity due to the large deflection of the wing. Results obtained by different orders of centre manifolds are compared with those obtained by time marching method (fourth-order Runge-Kutta method). These comparisons show zero, third and fifth order manifolds are very good approximation of this system. The aeroelastic model is a low aspect ratio rectangular cantilevered wing in a low subsonic flow which is structurally modeled by the Von Karman plate theory. A continuous time reduced order modified vortex lattice aerodynamics model is utilized in aerodynamics modeling.

Aerothermal behavior and performance of turbine blade squealer tip has been extensively studied, but previously mainly confined to very low speed flows. Some recent research efforts have revealed some distinctive tip heat transfer as well as aerodynamic characteristics in a transonic flow, many of which are confined to a linear cascade setting. There is a need to identify driving mechanisms for such configurations and to quantify corresponding performance characteristics in general, and in the process address how those drivers would relate to some specific issues of interest such as: 1) the differences between a cascade and a rotor tip, thus the impact of relative casing movement? 2) The difference between a transonic and a high subsonic flow condition? 3) the difference between a squealer tip and flat tip? In this study, the vortical structures and the loss mechanism of the Over Tip Leakage (OTL) flow in a high-subsonic stage and a transonic stage are analyzed using a proposed vortex model and Denton’s loss model. The two stages are specifically designed to work in the same Reynolds number with the efficiency comparable to the one predicted in the Smith’s chart. Two counter-rotating vortices, the Pressure-side Separation Vortex (PSV) and the Casing-driven Cavity Vortex (CCV), are identified as key drivers as inside a squealer cavity strongly impacting aerothermal performance (OTL mass flow and the heat transfer distribution). The cavity flow pattern results from the balancing between the CCV primarily driven by the relative casing movement, and the PSV originating from the flow separation around the pressure side rim. Consequently a detoured flow path is created within the squealer cavity when the leakage flow separating on the pressure side is entrained by the two vortices to the squealer bath and finally leaves the cavity along the suction side of the squealer wall. The entrainment of the leakage flow induces a signature of a high HTC strip on the squealer floor. Without the relative casing movement, the CCV is absent in the majority of the cavity so that the leakage flow can directly cross the tip which increases the leakage mass flow rate and reduce the HTC value on the cavity floor. Due to the distinct difference in vortical structures, the change of the HTC with and without relative casing movement can be interpreted by the vortex model. The vortex model is also able to explain the variation of the leakage mass flow rate and the HTC contours with the increasing squealer height. From the best knowledge of the authors, this is the first-of-its kind effort in using in the cavity vortices to systematically and quantitatively study the effect of the relative casing movement and the aerothermal performance of the tip with varying squealer height. In counting the aero losses for the two stages, three sources are attributed to the total loss generation: the baseline loss which assumes there is no leakage, the in-tip loss which is generated within the tip gap and the out-tip loss which accounts for the mixing loss between the OTL flow and the main passage flow. The out-tip loss is found to be the major source of the loss caused by the tip leakage flow. Denton’s loss model is used to calculate the out-tip loss quantitatively. The model predicted loss agrees in trend with the one calculated from the CFD results. According to the model, the out-tip loss is related to the main flow Mach number, the OTL mass flow rate and the velocity difference between the main flow and the OTL flow. From the distribution of these terms along the tip, the loss is mainly generated at the rear portion of the tip. The reduction of the leakage mass flow rate by the squealer is mainly at the zones where the CCV and the PSV are leaving the squealer cavity. The present analysis and results indicate that the squealer in the transonic stage tends to be less effective in reducing the out-tip loss compared to the subsonic one. The region of the tip where the CCV has already left the cavity with the PSV remaining shows the highest entropy generation rate, due to locally both the high mainstream velocity and the high OTL mass flow given the PSV being left as the only sealing vortex there.

A generalized reference enthalpy formulation for the skin friction, heat transfer and radiation-equilibrium temperature distributions over aerodynamic surfaces in attached hypersonic / hyperenthalpic flow is proposed. The formulation, which has been extensively employed in various forms by numerous investigators in the perfect gas regime, has also been recently demonstrated to provide adequate estimates of the heat transfer distribution in thermochemically active high enthalpy flow conditions when coupled to thermochemically active Euler solutions. It is now used to reveal the relevant similitude parameters for viscous effects in hypersonic flow, and the importance of the temperature distribution across the boundary layer and of the temperature-viscosity relation. It is shown that, although reproduction of the flight total flow enthalpy as well as surface temperature is the obvious solution for full viscous simulation in (perfect gas) hypersonic flow, the hot surface testing requirement and, in a number of practical applications, also the hot flow requirement may be relaxed with reasonably small error that can be of the same order as the measurement accuracy in present-day hypersonic testing. This similitude error, however, may increase significantly in cases exhibiting strong viscous/inviscid interaction or when the laminar-turbulent transition process becomes important. In this respect, alternative full simulation solutions, which are less demanding in terms of reproduction of the high levels of flight freestream and surface temperature or even Reynolds number, are discussed.

Asymptotic approximation for sound propagation in ducts carrying low subsonic mean flow