Wind Tunnels and Experimental Fluid Dynamics Research Part 15 ppt

Wind Tunnels and Experimental Fluid Dynamics Research 548 For the present qualitative analysis two dimensional computations carried out over the model symmetry plane are taken under consideration; in particular the conditions H 0 =35 MJ/kg, P 0 =2 bar are analyzed (this condition corresponding to the lower freestream Knudsen number: 1.47*10 -3 ) by comparing the results obtained with a classical Navier- Stokes approach and DSMC method, in order to check possible local effects of rarefaction. Note that for this high enthalpy case it has been decided to not perform the CFD slip computation since more accurate DSMC calculations are not strongly CPU-time demanding due to the reduced number of needed particles. Specifically, this test case is characterized by the following flow properties M ∞ = 12.94, Re ∞ /m = 9.03 × 10 3 , T ∞ = 240 K and a model attitude of 12 deg. A grid-independence study for CFD simulations has been carried out as well as a study of DSMC solution sensitivity to the number of particles (not shown). A preliminary analysis has been carried out considering the wall at fixed temperature of 300 K, and the following Fig. 13 and Fig. 14 show the Mach number contours and the streamlines for the two performed computations. Figures show the strong bow shock wave ahead of the model, that is more inclined, as expected, in the case of DSMC simulation, the strong expansion on the bottom part of the model, and finally the shock wave boundary layer interaction around the corner and the subsequent recirculation bubble, that is in incipient conditions in the case of rarefied flow simulation. Fig. 13. CFD: Mach number contours and streamlines Evaluation of Local Effects of Transitional Knudsen Number on Shock Wave Boundary Layer Interactions 549 Fig. 14. DSMC: Mach number contours and streamlines The Fig. 15 exhibits the slip velocity wall distribution predicted by DSMC calculation showing a peak value of about 1,3% of freestream velocity in correspondence of the beginning of the flat plate downstream of the model nose. It can be underlined that these low values of slip velocity were expected since, differently from the validation test case (i.e. the hollow cylinder flare), no sharp leading edge is present in this PWT model, therefore continuum regime flow conditions are predicted around the nose. Looking also at Fig. 15 , it can be observed that the same qualitative cuspid-like distribution has been predicted in correspondence of the corner, where a separation (or incipient separation like in this case) occurs. Fig. 15. Slip velocity distribution Wind Tunnels and Experimental Fluid Dynamics Research 550 By carefully examining Fig. 16 and Fig. 17, and remembering the analysis performed for the validation test case, the same considerations apply to the present applicative case in high enthalpy conditions. In particular, a reduction of separation extent is observed with DSMC calculation (see Fig. 13 and Fig. 14), as well as a slight reduction of the mechanical load acting on the flap (see Fig. 16). Finally, also looking at Fig. 15, in correspondence of the section where the maximum of slip velocity occurs, i.e. X=0.1 m, the local Knudsen number is: 2 1005.4 − ×≈= δλ δ Kn and this value justifies the occurrence of local effects of rarefaction on the prediction of important aspects of shock wave boundary layer interaction as well as the extent of separation region. Fig. 16. Pressure coefficient distribution Evaluation of Local Effects of Transitional Knudsen Number on Shock Wave Boundary Layer Interactions 551 Fig. 17. Skin friction coefficient distribution As a conclusion, it must be stressed the fact that local rarefaction effects must be taken into account when designing plasma wind tunnel tests at limit conditions of the facility envelope, in particular for very low pressures and high enthalpies as in the present case. This is particularly true when plasma test requirements are represented by the reproduction on the test model (or on parts of it) of given values of mechanical and thermal loads, as well as of shock wave boundary layer interaction characteristics (i.e. separation length, peak of pressure, peak of heat flux, etc.). 4. Conclusion Local effects of rarefaction on Shock-Wave-Boundary-Layer-Interaction have been studied by using both the continuum approach with the slip flow boundary conditions and the kinetic one by means of a DSMC code. The hollow cylinder flare test case for ONERA R5Ch wind tunnel conditions was numerically rebuilt in order to validate the methodologies. The free stream Knudsen number for the selected test case implies that much of the flow is in continuum conditions, even though local effects of rarefaction have been checked. In particular, the comparison with experimental data has shown that rarefactions effects are not negligible in prediction of the separation length. The CFD code with slip flow boundary conditions has shown good predicting capabilities of the size of the recirculation bubble, and the analysis of the density profiles inside boundary layer has shown a good agreement between DSMC and CFD with slip conditions in different sections along the body. Definitively, the present wind tunnel test case, simulated with the three different methodologies (classics CFD, CFD with slip flow boundary conditions and DSMC), has shown that local rarefaction effects are significant for the prediction of important aspects of shock wave boundary layer interaction as the sizing of recirculation bubble and it has been also shown that CFD with slip flow boundary conditions is, in this case, a good compromise between computational cost and accuracy. Wind Tunnels and Experimental Fluid Dynamics Research 552 The same considerations apply to a CIRA Plasma Wind Tunnel test case, where significant rarefactions effects were found on the SWBLI phenomenon; therefore they must be taken into account when designing plasma wind tunnel tests at limit conditions of the facility envelope, in particular for very low pressures and high enthalpies as in the present case. 5. References Bird, G. A., Molecular Gas Dynamics and the Direct Simulation of Gas Flows, Clarendon, Oxford, 1994. Bird, G. A., “The DS2V/3V Program Suite for DSMC Calculations” Rarefied Gas Dynamics, 24th International Symposium, Vol. 762 edited by M. Capitelli, American Inst. Of Physics, NY, 2005, pp. 541-546, February, 1995. Borrelli S., Pandolfi M., “An Upwind Formulation for the Numerical Prediction of Non Equilibrium Hypersonic Flows”, 12th International Conference on Numerical Methods in Fluid Dynamics, Oxford, United Kingdom, 1990. Chanetz, B., Benay, R., Bousquet, J., M.,Bur, R., Pot, T., Grasso, F., Moss, J., Experimental and Numerical Study of the Laminar Separation in Hypersonic Flow", Aerospace Science and Technology, No. 3, pp. 205-218, 1998. Di Clemente M., Marini M., Schettino A., “Shock Wave Boundary Layer Interaction in EXPERT Flight Conditions and Scirocco PWT”, 13th AIAA/CIRA International Space Planes and Hypersonics Systems and Technologies Conference, Capua, Italy, 2005. Kogan N. M., Rarefied Gas Dynamics, Plenum, New York, 1969. Markelov, G., N., Kudryavtsev A. N., Ivanov, M., S., “Continuum and Kinetic Simulation of Laminar Separated Flow at Hypersonic Speeds”, The Journal of Spacecraft and Rockets, Vol. 37 No. 4, July-August 2000. Marini, M., “H09 Viscous Interaction at a Cylinder/Flare Junction”, Third FLOWNET Workshop, , Marseille, 2002. Millikan R.C., White D.R., “Systematic of Vibrational Relaxation”, The Journal of Chemical Physics, Vol. 39 No.12, pp. 3209-3213, 1963. Park C., “A Review of Reaction Rates in High Temperature Air”, AIAA paper 89-1740, June 1989. Park C., Lee S.H., “Validation of Multi-Temperature Nozzle Flow Code NOZNT”, AIAA Paper 93-2862, 1993. Ranuzzi, G., Borreca, S., “CLAE Project. H3NS: Code Development Verification and Validation”, CIRA-CF-06-1017, 2006. Yun K.S., Mason E. A., “Collision Integrals for the Transport Properties of Dissociating Air at High Temperatures”, The Physics of Fluids, Vol. 39 No.12, pp. 3209-3213, 1962. 27 Investigation on Oblique Shock Wave Control by Surface Arc Discharge in a Mach 2.2 Supersonic Wind Tunnel Yinghong Li 1 and Jian Wang 2 1 Engineering College, Air Force Engineering University 2 Army Aviation Institute China 1. Introduction A shock wave is a typical aerodynamic phenomenon in a supersonic flow, and if controlled effectively, a series of potential applications can be achieved in aerospace fields, such as reducing wave drag and sonic boom of the supersonic vehicle, optimizing shock waves of the supersonic inlet in off-design operation states, decreasing pressure loss induced by shock waves in the supersonic wind tunnel or aeroengine internal duct, controlling shock waves of the wave rider, changing shock wave symmetry to achieve flight control and inducing shock waves in the aeroengine nozzle to achieve thrust vector control. Shock wave control can be achieved by many mechanical or gas dynamic methods, such as the ramp angle control in supersonic inlet and the holl/cavum control in self-adapted transonic wing. Because the structural configurations of these methods are somewhat complex and the flow control response is also slow, plasma flow control based on gas discharge physics and electromagnetohydrodynamics (EMHD) theory has been developed recently in the shock wave control field. Using this method, substantial thermal energy can be added in the shock wave adjacent areas, then the angle and intensity of shock wave change subsequently. Meyer et al investigated whether shock wave control by plasma aerodynamic actuation is a thermal mechanism or an ionization mechanism, and the experimental results demonstrated that the thermal mechanism dominates the shock wave control process [1, 2]. Miles et al investigated the shock wave control by laser energy addition experimentally and numerically, and the research results showed that when the oblique shock wave passed by the thermal spot induced by laser ionization, the shock wave shape distorted and the shock wave intensity reduced [3]. Macheret et al proposed a new method of virtual cowl induced by plasma flow control which can optimize the shock waves of supersonic inlet when its operation Mach number is lower than the design Mach number [4]. Meanwhile, they used the combination method of e-beam ionization and magnetohydrodynamic (MHD) flow control to optimize the shock waves of supersonic inlet when operating in off-design states, and the research results demonstrated that the shock waves can reintersect in the cowl adjacent area in different off-design operation states with the MHD acceleration method and the MHD power generation method, respectively [5]. Leonov et al used a quasi-dc Wind Tunnels and Experimental Fluid Dynamics Research 554 filamentary electrical discharge, and the experimental results showed that shock wave induction, shock wave angle transformation and shock wave intensity reduction, etc could all be achieved by plasma flow control [6-8]. Other than oblique shock wave control, the bow shock wave control by plasma aerodynamic actuation was also studied by Kolesnichenko et al [9], Ganiev et al [10], and Shang et al [11] for the purpose of reducing peak thermal load and wave drag. This paper used the arc discharge plasma aerodynamic actuation, and the wedge oblique shock wave control by this plasma aerodynamic actuation method was investigated in a small-scale short-duration supersonic wind tunnel. The change laws of shock wave control by plasma aerodynamic actuation were obtained in the experiments. Moreover, a magnetic field was applied to enhance the plasma actuation effects on a shock wave. Finally, a qualitative physical model was proposed to explain the mechanism of shock wave control by plasma aerodynamic actuation in a cold supersonic flow. 2. Experimental setup The design Mach number of the small-scale short-duration supersonic wind tunnel is 2.2 and its steady operation time is about 30-60 s. The test section is rectangular with a width of 80mm and a height of 30 mm. The gas static pressure and static temperature in the test section are 0.5 atm and 152 K, respectively. The groove in the test section lower wall is designed for the plasma aerodynamic actuator fabrication. The power supply consists of a high-voltage pulse circuit and a high-voltage dc circuit. The output voltage of the pulse circuit can reach 90 kV, which is used for electrical breakdown of the gas. The dc circuit is the 3 kV-4 kW power source, which is used to ignite the arc discharge. The plasma aerodynamic actuator consists of graphite electrodes and boron-nitride (BN) ceramic dielectric material. Three pairs of graphite electrodes are designed with the cathode- anode interval of 5mm and the individual electrode is designed as a cylindrical structure which is embedded in the BN ceramic. The upper gas flow surface of electrodes and ceramic must be a plate to ensure no unintentional shock wave generation in the test section. The controlled oblique shock wave is generated by a wedge with an angle of 20 ◦ . As shown in figure 1, the plasma aerodynamic actuator is embedded in poly-methyl-methacrylate (PMMA) and then inserted into the groove of the test section lower wall. There are 10 pressure dots with a diameter of 0.5mm along the flow direction for the gas pressure measurement. As shown in figure 2, the static magnetic field is generated by a rubidium-iron-boron magnet which consists of four pieces. Two pieces construct the N pole and the other two pieces construct the S pole. The magnetic field strength in the zone of interaction is about 0.4 T. Based on the MHD theory, the main purpose of adding magnetic field is applying a Lorentz body force to the charged particles in the arc plasma, which can influence the plasma actuation effects on shock wave. The test systems consist of a gas pressure measurement system, a schlieren photography system and an arc discharge voltage-current measurement system. The gas pressure measurement system is used to measure and compute the oblique shock wave intensity with the data-acquisition frequency of 1 kHz and the acquisition time of 3-10 s. The schlieren photography system is used to photograph the configuration of the oblique shock wave. It uses the Optronis® high-speed CCD camera with the maximum framing rate of 200 000 Hz. Investigation on Oblique Shock Wave Control by Surface Arc Discharge in a Mach 2.2 Supersonic Wind Tunnel 555 For the purpose of acquiring the pulsed arc discharge process in the flow, the framing rate in this paper is selected as 8000 Hz with an exposure time of 0.0001 s and a resolution of 512 × 218 pixels. The arc discharge voltage and current are monitored by a voltage probe (P6015A, Tektronix Inc.) and a current probe with a signal amplifier (TCP312+TCPA300, Tektronix Inc.), respectively. The two signals are measured by a four-channel digital oscilloscope (TDS4104, Tektronix Inc.). Fig. 1. Sketch of arc discharge plasma aerodynamic actuator. Fig. 2. Sketch of magnet fabrication on the wind tunnel test section. Wind Tunnels and Experimental Fluid Dynamics Research 556 3. Test results and discussion 3.1 Electrical characteristics Under the test conditions of Mach 2.2, the arc discharge is a pulsed periodical process with a period of 2-3 ms, and the discharge time only occupies 1/20 approximately in a period. The discharge voltage-current curves including several discharge periods are shown in figure 4(a). It can be seen that the discharge intensity is unsteady with some periods strong but some other periods weak. The discharge voltage-current-power curves in a single period are shown in figure 4(b). The discharge process in a single period can be divided into three steps. The first step is the pulse breakdown process. When the gas breakdown takes place, the discharge voltage and the current can reach as high as 13 kV and 18 A, respectively, and the discharge power reaches hundreds of kilowatts. However, this step lasts for an extremely short time of about 1μs, which indicates that it is a typical strong pulse breakdown process. The second step is the dc hold-up process. After the pulse breakdown process, arc discharge starts immediately. The discharge voltage decreases from 3 kV to 300- 500V and the discharge current increases to 3-3.5A correspondingly. The discharge power is maintained at 1-1.5 kW. This step lasts for a long time of about 80μs. The third step is the discharge attenuation process. Because the supersonic flow blows the plasma channel of the arc discharge downstream strongly, the Joule heating energy provided by the power supply dissipates in the surrounding gas flow intensively. As a result, the discharge voltage increases gradually. Both the discharge current and power decrease. When the power supply cannot provide the discharge voltage, the discharge extinguishes. After some time, the next period of discharge will start again. This attenuation step lasts for about 20μs. The time-averaged discharge power of the above three steps within 100μs is about 1.3kW. From figure 3 we can see that the arc discharge plasma is strongly bounded near the wall surface and blown downstream by the supersonic flow. The arc discharge is transformed from a large-volume discharge under static atmospheric conditions to a large-surface discharge under supersonic flow conditions. Fig. 3. Arc discharge picture in the supersonic flow. 3.2 The wedge oblique shock wave control by typical plasma aerodynamic actuation Three pairs of electrodes discharge simultaneously in the experiments. Under the conditions of an input voltage of 3 kV and an upwind-direction magnetic control, the wedge oblique shock wave control by this plasma aerodynamic actuation was investigated in detail. Because of the fabrication error and actuator surface roughness, there are some unintentional shock waves in the test section before the wedge. The wedge in the supersonic flow generates a strong oblique shock wave, which can be seen from figure 5(a). Because the boundary layer in the test section lower wall before the wedge is somewhat thick with a thickness of about 3-4 mm, the start segment of the oblique shock wave is composed of many weak compression waves, which intersect in the main flow to form the strong oblique shock wave. [...]... numerical and the experimental results for both cases 3 2 Y 1 0 7 8 X 9 a) Numerical 10 b) Experimental Fig 2 Experimental and numerical flow field over the model with large belt (15 ) at   0 574 Wind Tunnels and Experimental Fluid Dynamics Research Figure 3 shows the shock wave formed in front of the model and ahead and behind the belt used for varying the model cross section at 6 degrees angle of attack... field, and the pressure and density contours for flow over the small belt model at zero angle of attack, Mach number of 1.6, and Reynolds a)   0 b)   4 and 6 Fig 4 Comparison of experimental, theoretical, and numerical longitudinal pressure distribution a) x/d=3 b) x/d=11.25 Fig 5 Comparison of experimental and theoretical circumferential pressure distribution at   4 and 6 576 Wind Tunnels and. .. circumferential pressure distribution at   4 and 6 576 Wind Tunnels and Experimental Fluid Dynamics Research 15 Y Y 15 10 10 5 5 0 0 0 10 0 20 X a) Grid with five blocks used 10 15 X 20 b) Velocity vector field 15 15 10 Y Y 5 5 10 Belts 5 0 0 0 5 10 X 15 c) Pressure contours 20 0 5 10 X 15 20 d) density contours Fig 6 Experimental Flow around the model with small 5 belt number of 8×106 (described... wedge angle is designated as q * The angle and intensity of the induced oblique shock wave are designated as b * and ps* , respectively As q * < q and on the condition of constant Mach number, the relationships of b * < b and ps* < ps can be concluded based on the oblique shock wave relations of ( Ma ~ q ~ b ) 566 Wind Tunnels and Experimental Fluid Dynamics Research Therefore, based on the above thermal... Discharge in a Mach 2.2 Supersonic Wind Tunnel 565   where Aa and Ab are the cross section area of region a and b respectively, M a and Mb are the mass flux of region a and b respectively Supposing the height of region a and b are 30mm and 2mm, respectively, so Aa Ab = 15 In our experiments, the Mach number and gas stagnation temperature of the cold supersonic flow are 2.2 and 300 K, respectively From... actuation without magnetic control, the intensity of the wedge oblique shock wave 560 Wind Tunnels and Experimental Fluid Dynamics Research decreases only by 1.5%, but when applying the upwind-direction magnetic control, it decreases by 8.8% Moreover, when applying the downwind-direction magnetic control, it decreases by 11.6% The experimental results showed that the maximum shock wave intensity decrease is... the grid lines, especially near the walls 570 Wind Tunnels and Experimental Fluid Dynamics Research There have been many research performed on the areas of generation and use of multi-block grids, grid generation techniques, data management methods in different blocks, production of grid generation softwares which optimally require less trained users, and quicker grid generation, especially for complex... results obtained are at most 15% However, it is noted that such difference is much less (less than 7%) in other figures 580 Wind Tunnels and Experimental Fluid Dynamics Research a) x/d=5.75 b) x/d=11.25 Fig 10 Effects of angle of attack on boundary layer profiles Investigations of Supersonic Flow around a Long Axisymmetric Body a) x/d=5.75 b) x/d=11.25 Fig 11 Experimental and numerical boundary layer... area between arc discharge plasma and upper duct wall, and region b that corresponds to the high-temperature area of arc discharge plasma The sketch is shown in figure 13 As the gas pressure of cold supersonic flow is about the high level of 104 Pa, arc discharge plasma often reaches the Local Thermal Equilibrium (LTE) 564 Wind Tunnels and Experimental Fluid Dynamics Research state approximately which... and Experimental Fluid Dynamics Research The computational and the experimental boundary layer profiles for the simple model (uniform cross section) were compared at three different longitudinal stations and at zero angle of attack in Fig 8 From this figure, the agreements of data are relatively close, particularly for z . case) occurs. Fig. 15. Slip velocity distribution Wind Tunnels and Experimental Fluid Dynamics Research 550 By carefully examining Fig. 16 and Fig. 17, and remembering the analysis. method and the MHD power generation method, respectively [5]. Leonov et al used a quasi-dc Wind Tunnels and Experimental Fluid Dynamics Research 554 filamentary electrical discharge, and. Fig. 2. Sketch of magnet fabrication on the wind tunnel test section. Wind Tunnels and Experimental Fluid Dynamics Research 556 3. Test results and discussion 3.1 Electrical characteristics