NUMERICAL AND EXPERIMENTAL INVESTIGATIONS OF TEST MODELS AERODYNAMICS

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1 NUMERICAL AND EXPERIMENTAL INVESTIGATIONS OF TEST MODELS AERODYNAMICS A.V. Vaganov, S.M. Drozdov, S.M. Zadonsky, V.I. Plyashechnic, M.A. Starodubtsev, S.V. Chernov, V.L. Yumashev TsAGI, Zhukovsky, Moscow Region, Russia For an inaccuracy analysis of balance experiment at stand conditions there are used experimental investigation results of test models aerodynamic characteristics. Statistical analysis of multiple investigation outcomes allow to obtain mean (nominal) values of aerodynamic coefficients and to estimate an accuracy of its definition. The results obtained may be considered as a quantitative estimation of wind tunnel (WT) that is assumed to be a complicated measuring complex. Using results mentioned above it is possible to draw a conclusion of necessary modifications of separate WT components in order to increase a responsibility and decrease a balance experiment inaccuracy. Problem solution of vehicle models aerodynamic characteristics definition using experimental methods has a number of peculiarities for different flight modes. One of the main points characterizing the experimental data validity is a completeness of taking into account of systematic inaccuracies. These inaccuracies can occur considerably due to differences between models experimental conditions at wind tunnels and natural flight conditions. At TsAGI it was developed and then put into experimental researches practice a test model possessing a shape near to winged air-space vehicle (ASV) form of "Space Shuttle and "Buran" class (Fig. 1). A model manufacturing without control surfaces simulation was aimed at wind tunnel testing for large angle of attack range (α мах 60 ). This range is rather importance for practical researches of aerodynamic characteristics of some vehicles. Experimental investigations of test model aerodynamic characteristics were performed at following TsAGI s wind tunnels: T-116, T-117, Shock UT-1M (SWT), Hot Shot IT-1 and Hot Shot IT-2M (HSWT) at common Mach number range М= and at Reynolds number range Re= [1]. At aerodynamic coefficients calculation the forces acting on a model were related to a product of dynamic pressure and wing planform square (S). At pitching moment coefficient definition a fuselage length (L) was assumed as a character size. Pitching moment coefficient is determined relatively to conditional mass center which is located at a distance of 0.65L from nose bluntness. Multiple investigation outcomes and statistical analysis of the results are shown a good correlation of aerodynamic characteristics obtained at TsAGI supersonic Figure 1. General test model view wind tunnels in accordance to Mach number and to M / Re parameter. An experimental data correlation of coefficient Сх o and lift-to-drag ratio K in accordance to M / Re parameter is shown that a longitudinal force coefficient demonstrates a minimal value for a M / Re parameter about This coefficient increasing at bigger and less correlation parameter values is caused by M and Re numbers influence correspondingly (Fig. 2). A.V. Vaganov, S.M. Drozdov, S.M. Zadonsky, V.I. Plyashechnik, M.A. Starodubtsev, S.V. Chernov, V.L. Yumashev, 2008

2 Section IV Figure 2. Changing of longitudinal force and maximal lift-to-drag ratio values of test model in accordance to M / Re parameter Statistical treatment results of multiple investigation data presented in Fig. 3 allows noticing that an inaccuracy in normal force coefficient measuring depends on angle of attack value. This inaccuracy increases about 3 6 times during angle of attack changing from от 0 to 50. An angle of attack changing influences considerably less on inaccuracies in definition of longitudinal force and pitching moment coefficients. This fact is caused, in general, by changing character of proper coefficients in dependence of angle of attack. Obtained numerous experimental data of test model aerodynamic characteristics have the meaning itself. These data may be considered as a basis of program packages testing both used now and program developing. The above mentioned packages are based on different numerical codes. It should be noticing that in order to extend the verification possibilities of numerical codes it is advisably to add the flowfield investigation results near a test model surface to these experimental data. With this purpose experimental researches of test model were continued at T-117 TsAGI wind tunnel at Mach numbers М=7.5; Re L =( ) 10 6 and М=8.3; Re L = For verification possibility extending it was made a model revision that included a wing manufacture with exchangeable elevons (δ el = 0, 17 and 20 ) and a flap (δ d = 0 and 17 ) located under a fuselage. The elevons and a flap occur a wing/fuselage stern along a length at a trim of it windward surface. The relative squares of elevons and a flap are S el /S = and S d /S = Experimental researches performed at these Mach numbers have had two aims. At first, it needs to get an estimation of stream quality of such nozzles with reference to integral aerodynamic characteristics investigation. At second, it should be necessary to get data of test model surface flowfield during simulation of control surface deflection. σ Cy M=4.05 M=6.95 M=9.77 M=10.5 M=13.6 σ Cx M=4.05 M=6.95 M=9.77 M=10.5 M= α ο α ο Figure 3. Root mean square inaccuracy of force coefficients definition 2

3 Figure 4. Computational grid blocks on forward and lateral boundaries of base calculation domain Along with experimental investigations the numerical simulation of model streamlining was performed. Such simulation was carried out using two program packages, the former of that have been an ARGOLA-2 package developed at TsAGI [2]. In this package it is realized a multi-zone technology for flow calculation of non-viscose and non-heat-transfer gas with complex geometry and topology. The surface subdividing of forward boundary of calculation domain and lateral boundary (symmetry plane) of test model is shown in Fig. 4. Inviscid streamlining problem in ARGOLA-2 package was set up for unsteady gasdynamic equations. Three-dimensional steady flowfield was calculated by time-approximation technique using Godunov-Colgan method and its ascertained modifications. These calculation techniques were started by introducing some initial flow gas parameter distribution and then these data were evolutionary calculated to its limit state. The second calculation was performed using numerical solution of Reynolds averaged completed Navier-Stokes equations for perfect heat-transfer gas. These numerical calculations were made utilizing industry standard software "ANSYS CFX". Numerical integration of Navier-Stokes differential equations is carried out in frameworks of "ANSYS CFX" industry standard software using limited volume technique on hexahedral grid of mesh points. Three-dimensional grid developing is performed in frameworks of ANSYS ICEM CFD package. Some fragments of this grid are presented in Fig. 5. A grid model has some thickenings at boundary layer region and near a vehicle nose where shock layer is formed. During equations solution a first order time-approximation technique is used. The obtained solution possesses a first order approximation type. 3

4 Section IV Cya, K 2.0 α=30 ο m z 0.04 α=30 ο 1.5 K The first series of tests (WT T-116 и Т-117 TsAGI) The second series of tests (WT Т-117 TsAGI) Cya M M At a first investigation stage let us analyze a convergence of balance test results obtained during the second series (WT T-117, M=7.3 and 8.5) with model aerodynamic characteristics of first series test (WT T-116 (M=2-9.8) and T-117 TsAGI (М=10.5, 16.8)) [1] The first series of tests (WT T-116 и Т-117 TsAGI) The second series of tests (WT Т-117 TsAGI) Figure 6. A comparison of test model aerodynamic characteristics of first and second experimental series An accordance of aerodynamic coefficients of lift, pitching moment and lift-to-drag ratio is illustrated in plots presented in Fig. 6. A feature of these curves is the point that plotted data of first series test are the mean values obtained as a result of statistical treatment of tenfold tests. Let us note a wholly satisfactory correlation of a second series outcomes of experimental investigation of test model aerodynamic characteristics (M=7.5 and 8.3) and first series data. One of the main investigation stage of test model aerodynamics is an experiment data application to numerical codes testing. As a first step it can be considered an accordance of integral aerodynamic characteristics of test model obtained both by numerical streamlining simulation and by experimental methods. Comparisons of experimental data and numerical simulation results are shown in Fig. 7. Let us note a satisfactory accordance of calculated and experimental values of aerodynamic characteristics including an agreement of drag coefficient Сх obtained with an aid of CFX package. A marked divergence in Сх coefficient value during calculation using ARGOLA-2 package is caused, in general, by viscous effects which don t simulate in Euler equations framework. Lift coefficient is defined, in general, by non-viscous component. Well developed system of shock waves is realized near a model surface with undeflected control surfaces. This shock system is caused by supersonic overflow of configuration elements. At surface breaking places, for example, at a transition from nose to Figure 7. A comparison of experimental and calculated values of aerodynamic characteristics 4

5 cylindrical central fuselage part and at windward surface stern trim there are observed rarefaction wave regions (Fig. 8). It should be especially noted a good agreement of experimental and calculated values of pitching moment. A divergence in these values is characterized by difference in balancing angle of attack values which equals about 1.5. A flow pattern near a test model surface was investigated using shadow method and with the aid of oil dye drop spread. A flowfield near a model surface have a complex, three-dimensional character. This fact is caused by model geometry and by large range of angles of attack. At zero angle of attack a spreading line is located on an upper surface of a nose part having a negative inclination angle and on a wing leading edge blunted spherically. A spreading line is observed in a symmetry plane of lower surface also. Thus a flow in a shock layer near a configuration lower surface after a turn near a wing panel leading edge separates and then reattaches to a vehicle symmetry plane. Flow structure near a windward surface changes during angle of attack increasing. At angle of attack α=35 it is observed a flow spreading on a whole vehicle windward surface. This is characteristically for large angle of attack mode. At end sections of a wing and a flap deflected on an angle of 17 о it is observed a more intensive flow spreading and a curvature of streamlines increases. In front of elevons deflected on an angle of 20 the flow separation takes place and then it reattaches to flaps sections surface. Separation and reattachment lines have a curvilinear form. An interference of overflow about elevons sections deflected on an angle of 20 is minimal and it value is defined by interaction at fuselage region only where gas outflow from separation region occurs (Fig. 9). Figure 8. Experimental investigation of flowfield near a test model surface in a mode of undeflected control surfaces (M=7.5) 5

6 Section IV Figure 9. Flowfield experimental investigation near a test model surface at δ el =20 о и δ d =17 о. A shadow pattern analysis shows that at a range of large angle of attack a flow inside an interaction region between a head shock wave and a compression shock caused by deflected elevons has a complex three-dimensional structure. An interaction region shape at an angle of attack α = 45 allows to assume about transonic flow mode in front of deflected elevons. For calculation results verification during overflow gasdynamics features investigation the shadow patterns and oil spectra were used. These data were obtained during test model investigation in TsAGI s wind tunnel T-117. At first let us consider that shock wave calculation shape is in a good agreement to a pattern experimentally observed (Fig. 10). Figure 10. A comparison of calculated and experimental locations of shock waves. Dark points experiment; M=7.5; Re=10 6 6

7 a) b) Figure 11. A comparison of flow patterns near a model surface obtained by using CFX package flow simulation and during experiment with the aid of oil points. a)- upper (leeward) surface; b) lower (windward) surface As concerns to a flow in disturbed region on the whole that flow calculation simulates not only gasdynamics structures forming inside a volume between model surface and shock wave but it simulates the phenomena near test model surface caused by boundary layer flow. In particular, streamline calculation patterns on a model surface are in a good agreement with experimental results (Fig. 11). Conclusions. In this activity there are presented results of numerical/experimental investigation of aerodynamic characteristics and flowfields near a test model surface. This model shape is similar to winged air-space vehicle form of "Space Shuttle and "Buran" class. The experimental investigation of aerodynamic characteristics and flowfields near a test model surface were performed at TsAGI s wind tunnel T-117 at supersonic freestream velocities: M=7.5; Re L =( ) 10 6 and М=8.3; Re L = Numerical simulation of test model three-dimensional streamlining and model aerodynamic characteristics calculation was performed using two program packages. At first, it is an ARGOLA-2 package developed at TsAGI. In this package it is realized a multi-zone technology for flow calculation of non-viscose and non-heat-transfer gas. At second, it is industry standard software "ANSYS CFX". In frameworks of this package numerical integration of Navier-Stokes differential equations is carried out using limited volume technique. Comparison of aerodynamic characteristics experimental values and analogous data of first series investigation of test model [1] demonstrated a good correlation of two series test data. Numerical methods verification has shown that integral aerodynamic characteristics calculation results have been in a good agreement with experimental data. The data obtained are useful for estimation of reliability and inaccuracy in aerodynamic coefficients measuring in supersonic wind tunnels. These data are useful for numerical program testing also (both for numerical codes existing now and for program packages witch could be future developed to calculate the aerodynamic characteristics of hypersonic vehicle). The work was performed under Financial Support of Department aimed program in frameworks of National Project of High Education School Development DAP VP HESD (ВЦП НПР ВШ ). REFERENCES 1 A.V. Vaganov, S.M. Zadonsky, V.I. Plyashechnik. Results of aerodynamic characteristics researches of test models in TsAGI wind tunnels// Intern. Conf. on the Methods of Aerophysical Research: Proc. Part V, Novosibirsk, P Kosyh A.P., Nercesov G.G., Chelysheva I.F., Yumashev V.L. Numerical simulation of three-dimensional overflow about supersonic vehicles and their elements on a basis of multi-zone technology. TsAGI sci. notes, v. XXXV, 1-2,