INVESTIGATIONS OF AEROGASDYNAMICS OF RE-ENTRY BALLISTIC VEHICLE EXPERT

Transcription

1 International Conference on Methods of Aerophysical Research, ICMAR 28 INVESTIGATIONS OF AEROGASDYNAMICS OF RE-ENTRY BALLISTIC VEHICLE EXPERT N.P. Adamov, A.M. Kharitonov, I.I. Mazhul, L.G. Vasenyov, V.I. Zapryagaev, V.I. Zvegintsev Khristianovich Institute of Theoretical and Applied Mechanics SB RAS, 639, Novosibirsk, Russia J.M. Muylaert ESTEC, Noordwijk, Netherlands Introduction The EXPERT (European EXPErimental Re-entry Testbed) ballistic re-entry capsule was proposed by the European Space Agency (ESA) [1, 2]. The vehicle, designed for obtaining in-flight experimental data on aerothermodynamics at high supersonic velocities, is supposed to be launched by the Volna launcher with subsequent ballistic descent [3]. The EXPERT vehicle will be used for in-flight research of various critical aerothermodynamic phenomena, such as the boundary-layer transition on blunt bodies, real gas effects during shock wave/boundary layer interaction, effect of surface catalyticity, etc. The effective solution of the problem posed implies comprehensive pre-flight research, including numerical studies by advanced numerical methods and experimental studies of models in ground-based facilities for possible extrapolation of wind-tunnel data to flight conditions. The general view of the EXPERT model manufactured at a scale of 1:8 for tests in T-313 and AT-33 wind tunnels based at ITAM SB RAS is shown in Fig. 1. The results of earlier studies in ITAM wind tunnels can be found in [4-7]. The present paper describes new data obtained for the EXPERT model, namely: Fig.1. General view of the EXPERT model. - results of experimental studies of the total aerodynamic characteristics of the model in the T- 313 blowdown wind tunnel at a Mach number М = 4 and in the adiabatic-compression AT-33 wind tunnel with shaped nozzle at М = 14; - results of studying the boundary-layer transition on the model at М = 4 in the T-313 wind tunnel; - experimental data on the pressure distribution over the surface at М = 1 in the AT-33 wind tunnel; - results of numerical calculations of an axisymmetric model problem by the Euler and Navier- Stokes equations and comparison of calculated with experimental data on the pressure distribution over the model surface. 1. Experimental condition 1.1. Т-313 blowdown wind tunnel The experimental data in the Т-313 blowdown wind tunnel of ITAM SB RAS with the test section 6х6 mm were obtained at free-stream Mach number М = 4.4, unit Reynolds number Re = /m, and total pressure q = Pa, and in the range of incidences α = N.P. Adamov, L.G. Vasenyov, V.I. Zapryagaev, V.I. Zvegintsev, I.I. Mazhul, A.M. Kharitonov, J.M. Muylaert, 28 1

2 Section I The aerodynamic forces were measured with an AB-313 rider-type mechanical balance presenting a standard measuring tool for Т-313 tests. The width of measurement ranges over the drag force, lifting force, and pitch moment were respectively 11.8 kg, 61.7 kg, and 16.4 kgm. The instrumental error is estimated as %,.25-.8%, and.15-.5% of the indicated range, respectively. Besides, from time to time repeated tests of the HB-2 reference model were performed in the wind tunnel aimed at revealing possible random errors. According to the data obtained in a series of 15 multiple tests of the HB-2 model, the square-mean deviations at attack angles α = and 16 were respectively: - for longitudinal force coefficient.287 and.297, - for normal force coefficient.117 and.61, - for pitching moment coefficient.15 and.762. These deviations give an idea of possible random spread of experimental data; as an approximation, these values can also be used for the EXPERT model. Recalculation of these values to characteristic EXPERT sizes (area and length) yields, with a confidence probability of Р =.95, the following values of possible random errors: - for longitudinal force coefficient.923 and.925, - for normal force coefficient.356 and.196, - for pitching moment coefficient.135 and.735. Note that at determination angles of attack α measured with respect to the longitudinal axis of the model it was take into account the sting deformations owing to the acting loads. Apart from the total aerodynamic characteristics, schlieren pictures of the flow around the model, taken with the help of an IAB-451 Tepler divies, and soot-oil pictures of model surface streamlines were obtained, giving some presentations into the flow structure in the vicinity of the flaps АТ-33 adiabatic-compression wind tunnel The force measurement program in АТ-33 included measurements of aerodynamic characteristics of the EXPERT model at nominal Mach number М = 14, angles of attack α =, 3, and 6, and roll angles of the model with respect to the longitudinal axis γ =, 45, and 9. The study was carried out with the use of a profile nozzle with radius of the exit section D n = 4 mm at unit Reynolds number Re = /m. This value is close to natural in-flight values. Preliminarily, measurements of the flow velocity field with a Pitot comb were carried out. According to experimental data, at the indicated Reynolds number the Mach number at the nozzle plane was М = 13.8, the root-mean-square deviation being within ~.1. Force measurements in the АТ-33 facility were carried out with the help of six component strain-gage balances. Note that, previously, in execution of ISTC project No. 219, multiple tests at М = 1, Re ~ /m, and α = 6 were performed aimed at determination of the mean-square deviation of measured data directly for the EXPERT model. As a result, the following values of mean-square deviations were obtained [4, 7]: - for longitudinal force - σ C A =. 41, - for normal force - σ C Z =. 24, - for pitch moment - σ m z =. 22. Thus, measured values of aerodynamic coefficients include random errors, equal to ± 2σ, which corresponds to a confidence probability of P ~.95 in the Student distribution. These errors were then used to analyze the test data obtained at М = 14. In addition, to check the reliability of data obtained, doubled tests at α = 6 were conducted. The tests in the АТ-33 facility comprised two stages, a main test series (runs ) and a 2

3 International Conference on Methods of Aerophysical Research, ICMAR 28 preliminary series (runs ). Note that in the main test series the model nose was located in the nozzle exit section, and in the preliminary series, at a distance of 35 mm behind the nozzle exit. 2. Total aerodynamic characteristics All values of the aerodynamic-force coefficients presented below were normalized to characteristic area F ref =.187 m 2 ; the pitching moment coefficient was additionally normalized to length L ref =.2 m. The pitching moment was determined with respect to the model nose tip. The initial position of the model in the wind tunnels corresponded to the state with open flaps located in the vertical symmetry plane of the model; in the present test series, the deflection angle of all the four flaps was δ = T-313 wind tunnel In processing the measured data for aerodynamic loads on the EXPERT model and in determining the flow conditions, the routine procedure for the Т-313 facility was used. The obtained values of aerodynamic coefficients in the coordinate system of the model sting are shown in Fig. 2 versus the angle of attack. Here, we consider two positions of the model, the initial position with the roll angle γ =, and a second position with the model rotated with respect to the longitudinal axis through roll angle γ = 45. First of all, it should be noted that a nonzero angle of the model had almost no effect on the aerodynamic characteristics throughout the whole examined range of attack angles. From the theoretical point of view, it is quite clear that the characteristics at α = should be independent of model rotation with respect to the longitudinal axis. That is why a comparison of data obtained at α = and at roll angles γ = and 45 allows one to reveal possible errors induced by flow asymmetry in the experiments. For instance, these may be wash-induced errors, errors owing to a slip angle of the model misaligned on the support or on the α-gear of the wind tunnel, etc. The obtained almost perfect data coincidence for γ = and 45 at α = points to no errors of the indicated Fig. 2. Aerodynamic characteristics of the EXPERT model at М = 4. Fig. 3. Soot-oil visualization data obtained in Т-313 tests. 3

4 Section I type and good repeatability of the tests in Т-313 wind tunnel. It should be noted that the longitudinal force coefficient C A somewhat increases with the angle of attack. For α = -16 this coefficient ranges in the interval C A = Apart from the total aerodynamic characteristics, some data were obtained concerning the flow structure near the model at the regimes under consideration. For instance, Fig. 2 exemplifies the schlieren picture of the flow around the model at α = and γ =. In the images, the boundarylayer region on the model and the boundary-layer interaction with flap-induced compression shocks are distinctly seen. From these data it can be concluded that the separation of the boundary layer, if any, takes place over rather a small region, especially at low angles of attack. This conclusion is corroborated by soot-oil visualization pictures of the flow over the model surface obtained at α =. An example is given in Fig. 3; clearly visible here are oil spreading lines interpreted as lines of boundary-layer flow separation АТ-33 wind tunnel The tests in the adiabatic-compression 8 16 wind tunnel were carried out at nominal freestream Mach number M = 14 with the use of a Po/1 2, To, KPa K profil nozzle with outlet diameter D n 4 = 4 1 mm. 2 8 The АТ-33 facility is an impuls-type wind tunnel; that is why a critical point here is 1.6 realization of a working regime with.28 q/1 parameters of the incident flow weakly 2, Re 1 /1 7, KPa.24 1/m 1.4 changing in time. Graphs showing how the.2 regime of the flow at the nozzle exit varies in time during a wind-tunnel run are shown in.8 1. Fig. 4; presented here are the stagnation pressure Р, the stagnation temperature T, the velocity head q, and the unit Reynolds number Fig. 4. АТ-33 wind tunnel free-stream parameters. Re versus time. Noteworthy is rather a weak variation of the flow quantities over a time interval lasting for t ~ 1 ms; these data prove realization of a working regime necessary for measurements. The experimental data obtained are shown in Fig. 5 for aerodynamicforce coefficients in the balance frame. Plotted in the same graph is the probable level of random errors as predicted from previous data obtained in multiple tests at М = 1 [4, 7]. It should be noted that the longitudinal force coefficient, varying in the interval C A ~ , displays an insignificant (within 1%) increase over the angle of attack. Repeated tests at α = 6 show that, here, rather a good convergence of data, especially for the longitudinal force coefficient C A, takes place. The spread in the measured values of the normal force and pitching moment γ = γ = γ = 45 o γ = 9 o - preliminary series C z ± 2σ ± 2σ ± 2σ C A Fig. 5. Aerodynamic characteristics at М = 13.8, Re 1 ~ /m α, deg m z

5 International Conference on Methods of Aerophysical Research, ICMAR 28 coefficients lies within the limits of the square-mean inaccuracies. Note that the data presented in Fig. 5 were obtained in two test series, the data convergence being rather good. Very probably, some deviation of data at α = 6 was a consequence of different locations of the model with respect to the nozzle exit. At α = 6, we investigated into the effect due to the roll angle upon model rotation with respect to the balance longitudinal axis. The data were obtained for roll angles γ =, 45, and 9. Recall that the angle γ = refers to the open state of one of the flap pairs in the vertical symmetry plane, and the angle γ = 9, in the horizontal symmetry plane. Like in the Т-313 tests at М = 4, the aerodynamic coefficients here proved to be independent of the model roll angle. The spread of С А can be argued to be within ± 2σ C A. The normal force and pitching moment coefficients display a similar behavior, although the spread of experimental values here is more pronounced. 3. Boundary-layer transition The experimental data for the boundary-layer transition on the surface of the EXPERT model were obtained in the Т-313 wind tunnel at free-stream Mach number М = 4.4 for the angle of attack α =. For the purposes pursued in the present study, a method was used in which the boundarylayer transition was measured with the help of Pitot probes located directly on the surface of the model at a fixed point х along the longitudinal axis, employing variation of approach-flow quantities, in particular, variation of the unit Reynolds number Re 1 1/m. In this method, at the values of approach-flow quantities corresponding to the start and end points of the boundary-layer transition the Pitot probes are expected to register respectively a minimum and maximum pressure. Thus, analyzing the gained dependences for the measured values of the relative pressures p / p, we were able to determine the critical Reynolds numbers for the boundary-layer transition start and end points. Here and below, р is the total pressure in the forward flow, assumed to be equal to the wind-tunnel setling chamber pressure, and р is the pressure measured by a Pitot probe installed on the model. As the Pitot tubes, syringe needles with outer diameter d =.4 mm and inner diameter d 1 =.15 mm were used. The initial state of the model in the wind tunnel was with open flaps located in the vertical symmetry plane. Two Pitot probes were installed on the model, one in the mentioned vertical plane of symmetry at a longitudinal distance х = 26.7 mm from the nose tip (hereafter to be referred to as point 1), and the other, in a plane at a rolling angle of 45 with х = 24.9 mm (point 2). The position of the Pitot probes and a photograph of the model forebody are shown in Fig. 6. These Pitots were mounted on the surface of the model so that the distance between the nose tip and the Pitot location was greater than (15-2) d. The free-stream unit Reynolds number Re 1 was varied in each test by measuring the total pressure in the wind-tunnel stilling chamber in the interval р (1.7-13) 1 5 Pa. In this way, the range Re 1 (8-64) 1 6 1/m was covered. The total pressures here were measured with the help of TDM2-A pressure sensors whose measurement range was up to Pa with measurement accuracy.5%. The actual pressure level was (.1-.9) 1 5 Pa. Two series of tests were performed. The first series was performed on the model with the natural non-controlled surface roughness that took place after the force measurements in АТ-33 and Т-313 wind tunnels as a result of model surface exposure to the stream, especially on the blunt forebody. In the second series, the roughness was decreased appreciably by polishing the model surface. The measured data were used to plot curves of p / p versus the Reynolds number Rex = Re 1 x (here, х is the longitudinal coordinate of the Pitot probe at point 1 or 2). These curves are shown in Fig. 6 as obtained in two test series, one with a rough, and the other with 5

6 Section I smooth, model surface. Note that both Pitot probes registered close values of p / p since these probes were located at roughly same distances from the nose tip, in the matching region of the spherical bluntness with the conical part of the model. Just one maximum displayed by the curves p / p = f (Rex) points to the fact that, in our experiments, only the end point of the boundary-layer transition region was registered. Identification of the start point of the boundary-layer transition necessitates reaching lower numbers Re1, which was impossible because of violation of the wind-flow regime at low total pressures р < Pa. Additionally, it follows from the data of Fig. 6 that surface roughness largely Fig. 6. Determination of the boundary-layer transition. affects the magnitude of the critical Reynolds number Re et. For instance, for the rough surface a value Re et ~ was obtained, whereas for the smooth surface the boundary-layer transition was found to occur farther downstream, the critical Reynolds number in the latter case being increased to Re et ~ On the whole, it can be concluded that over the most part of the surface, including the flap location, a turbulent boundary layer took place. This conclusion complies with the obtained soot-oil pictures of surface streamlines (М = 4.5, Re /m) indicative of only an insignificant separation zone resulting from the boundary-layer interaction with flap-induced compression shocks. The obtained values of critical Reynolds numbers enable estimation of the probable position of the boundary-layer transition end point in natural flight conditions. For instance, for fight trajectories at a re-entry velocity V = 5 m/s the value of the Reynolds number calculated from the vehicle length is Re L , the unit Reynolds number being Re /m. Then, the position of the transition end point, calculated as x et = Re et / Re1, is.83 m and.52 m respectively for the smooth and rough surfaces. In other words, the laminar-turbulent transition takes place near the nose tip, on the clotoid portion connecting the spherical blunting with the conical part of the surface. 4. Pressure distribution over the surface Measurements of pressure distributions over the surface of the EXPERT model were carried out at Mach number М = 1 (conical nozzle with outlet diameter D n = 4 mm) at the following values of free-stream flow quantities: flow stagnation pressure Р = 985 kpa, stagnation temperature Т =124 K, unit Reynolds number Re 1 = /m. The incidences were α =, 3, and 6 o. For the whole pressure pattern on the windward and lee sides of the model to be grasped, at each angle of attack pressure measurements were carried out in two wind-tunnel runs with model rotation through 18 degrees with respect to the longitudinal axis. 6

7 International Conference on Methods of Aerophysical Research, ICMAR 28 For making the pressure measurements possible, the EXPERT model was provided with 2 pressure taps located in the vertical symmetry plane. Of these pressure taps, five taps were located on a flap with δ = 2. The diameter of all pressure holes was 1 mm. The pressure measurements were performed with TDM2-A absolute pressure sensors installed inside the model as close as possible to its surface. With the pressure taps considered from the model nose on, the pressure sensors rated measurement ranges were: to 3 atm. at the first two taps; to.6 atm. at tap 3; and to.3 atm. at taps 4-9. The rated measurement range of the pressure sensors for the remaining pressure taps was to.1 atm. The measurements and the registration of measured signals were carried out on an Eckelmann Steuerungstechnik GmbH SCP 32 highprecision multi-channel measuring system. The registration lasted for 3.3 seconds, including the time required for the hypersonic flow to pass the model, so that to additionally cover zero sensor indications prior to and after the air discharge from the nozzle. P /1 2, kpa P U U, V Fig. 7. The total pressure in the wind-tunnel setling chamber and the sensor signal versus time. A comparative picture of recorded signals for the pressure sensor installed in the windtunnel setling chamber and for the pressure sensor installed at the location of the first pressure tap versus time is shown in Fig. 7. At the first stage of the data treatment procedure, time intervals were identified in which working flow conditions were realized in the wind tunnel. Here, the duration of a) γ = α = = 3 o = 6 o b) γ = 18 o α = = 3 o = 6 o Fig. 8. Distribution of the relative pressure over the windward (a) and lee (b) surface of the model. the working regime was ~ 15 ms in the time interval from 14 to 155 ms. The pressure sensor produced a uniform signal suitable for subsequent treatment. Further treatment of registered signals was carried out using the routinely procedure accepted for АТ-33 [8]. The obtained pattern of the relative pressure over the generatrixes of the windward and lee surfaces in the symmetry plane of the model is shown in Fig. 8 for various angles of attack α. The same figure shows, on an enlarged scale, the pressure distribution outside the forebody, in the flap region, with the origin located at х = mm. As could be expected, an increase in the angle of 7

8 Section I Euler - k-ε NRG experiment - run run 2663 shock wave ~ 4.8 ~ Fig. 9. Calculated and experimental distributions of pressure at zero angle of attack. attack is accompanied with a rise of pressure on the windward side and with a reduction of pressure on the lee side. The average level of the relative pressure near the stagnation point is ~ 125 irrespective of incidence. A most pronounced increase of the pressure on the flaps is observed on increasing the angle of attack to six degrees. Unfortunately, the small number of pressure taps on the flap provides no clear picture of the flow around flaps, including the boundarylayer separation during the interaction of the boundary layer with the compression shock. With the aim of possible refinement of the picture, we attempted numerical calculation of an axisymmetric body with the EXPERT longitudinal contour. The calculations were performed using the FLUENT application suite based on solving Euler and Navier-Stokes equations on the assumption of the k-ε RNG turbulence. The obtained numerical data on the distribution of pressure at zero angle of attack are shown in Fig. 9. In particular, the Euler calculations allowed us to evaluate the flow quantities in front of the flap and the pressure levels on the flap. For the Mach number in front of the flap and for the pressure drop across the shock, values M 1 = 2.8 and p 2 / p1 ~ 3.5 were obtained. Comparison of the latter pressure level with the simplest estimate of the critical pressure drop p kr = M 1 = 2.4 for turbulent boundary layer points to possible separation of the boundary layer occurring during the boundary layer/compression shock interaction. Subsequent Navier-Stokes calculations confirmed the presence of a separation zone in the flap region. Note also that outside the separation zone the pressure distributions yielded by viscous and inviscid calculations are almost coincidental. This work was sponsored by ISTC in frame of the projects 219, 3151 and 355. REFERENCES 1. Walpot L., Ottens H. FESART/EXPERT aerodynamic and aerothermodynamic analysis of the RAV and KHEOPS configuration. Technical report, TOS-MPA/2718/LW, ESTEC, Ottens H., Walpot L. EXPERT model 4.2, Model description and trajectory analysis. Technical report, TOS- MPA/2749/HO, ESTEC, Danilkyn V. Pre-contractual studies of the possibility to launch re-entry vehicles by Volna LV for aerothermodynamic investigations. Technical report, / , Makeyev design bureau, Kharitonov A.M., Adamov N.P., Brodetsky M.D., Vasenyov L.G., Mazhul I.I., Zvegintsev V.I., Paulat J.C., Muylaert J.M., Kordulla W. Investigation of Aerogasdynamics of Re-Entry Vehicles in the New Hypersonic Wind Tunnel at ITAM // Proceeding of the 44th AIAA Aerospace Sciences Meeting and Exhibit, Jan. 9-12, 26, Reno, Nevada, USA (AIAA Paper No ). 5. Kharitonov A.M., Zvegintsev V.I., Brodetsky M.D., Mazhul I.I, Muylaert J.M., Kordulla W., Paulat J.C. Aerodynamic Investigation of Aerospace Vehicles in the New Hypersonic Wind Tunnel AT-33 in ITAM // 4th Intern. Symp. on Atmospheric Re-Entry Vehicles and Systems, March 21-23, 25 Arcachon, France. 8

9 International Conference on Methods of Aerophysical Research, ICMAR Kharitonov A.M., Adamov N.P., Brodetsky M.D., Vasenyov L.G., Mazhul I.I., Zvegintsev V.I. Investigation of Aerogasdynamics of Re-Entry Vehicles in the New Hypersonic Wind Tunnel at ITAM // Proceeding of the 6 th Cina-Russia high-speed flow conference, June 2-26, 26, Mianyang Sichuan. 7. Adamov N.P., Brodetsky M.D., Vasenyov L.G., Mazhul I.I., Zvegintsev V.I., Kharitonov A.M., Paulat J.C., Muylaert J.M., Kordulla W. Aerodynamics of re-entry vehicles at natural Reynolds numbers // Thermophysics and Aeromechanics. 26. V.13. -No.3. P Kharitonov A.M., Zvegintsev V.I., Vasenyov L.G., Kuraeva A.D., Nalivajchenko D.G., Novikov A.V., Paikova M.A., Chirkashenko V.F., Shakhmatova N.V., Shpak S.I. Characteristics of the AT-33 hypersonic wind tunnel. Part 1. Velocity fields //. Thermophysics and Aeromechanics. 26. V.13. No1. P