Showing papers on "Pitching moment published in 1999"

Robust nonlinear control of a hypersonic aircraft

Abstract: For the longitudinal motion of a hypersonic aircraft containing twenty-eight inertial and aerodynamic uncertain parameters, robust flight control systems with nonlinear dynamic inversion structure are synthesized. The system robustness is characterized by the probability of instability and probabilities of violations of thirty-eight performance criteria, subjected to the variations of the uncertain system parameters. The design cost function is defined as a weighted quadratic sum of these probabilities. The control system is designed using a genetic algorithm to search a design parameter space of the nonlinear dynamic inversion structure. During the search iteration, Monte Carlo evaluation is used to estimate the system robustness and cost function. This approach explicitly takes into account the design requirements and makes full use of engineering knowledge in the design process to produce practical and efficient control systems. A4 MY, m 4 Nomenclatm-e speed of sound, ftls drag coefficient lift coefficient moment coefficient due to pitch rate moment coefficient due to angle of attack moment coefficient due to elevator deflection thrust coefficient reference length, 80 ft drag, lbf altitude, ft moment of inertia, 7 X lo6 slug-ft2 lift, lbf Mach number pitching moment, lbf-ft mass, 9375 slugs pitch rate, radis radius of the Earth, 20,903,500 ft radial distance from Earth’s center, ft reference area, 3603 ft2 thrust, lbf velocity, ft/S angle of attack, rad throttle setting flight-path angle, rad elevator deflection, rad gravitational constant, 1.39 X 1Or6 ft3/s2~ density of air, slugsIft

High-Lift Optimization Design Using Neural Networks on a Multi-Element Airfoil

Abstract: Roxana M. GreenmanAerospace EngineerNASA Ames Research CenterMoffett Field, California 94035-1000, U. S. A.Tel: 650-604-3997, Fax: 650-604-2238E-mail: rgreenman@mail.arc.nasa.govKarlin R. RothAerospace EngineerNASA Ames Research CenterMoffett Field, California 94035-1000, U. S. A.Tel: 650-604-6678, Fax: 650-604-2238E-mail: kroth@mail.arc.nasa.govABSTRACTThe high-lift performance of a multi-element airfoil wasoptimized by using neural-net predictions that were trainedusing a computational data set. The numerical data was gener-ated using a two-dimensional, incompressible, Navier-Stokesalgorithm with the Spaiart-Allmaras turbulence model. Becauseit is difficult to predict maximum lift for high-lift systems, anempirically-based maximum lift criteria was used in this studyto detemaine both the maximum lift and the angle at which itoccurs. Multiple input, single output networks were trainedusing the NASA Ames variation of the Levenherg-Marquardtalgorithm for each of the aerodynamic coefficients (lift, drag,and moment). The artificial neural networks were integratedwith a gradient-based optimizer. Using independent numericalsimulations and experimental data for this high-lift configura-tion, it was shown that this design process successfully opti-mized flap deflection, gap, overlap, and angle of attack tomaximize lift. Once the neural networks were trained and inte-grated with the optimizer, minimal additional computerresources were required to perform optimization runs with dif-ferent initial conditions and parameters. Applying the neuralnetworks within the high-lift rigging optimization processreduced Me amount of computational time and resources by83% compared with traditional gradient-based optimization pro-cedures for multiple optimization runs.NOMENCLATURECa drag coefficient, C a - D/(q_c)Ct lift coefficient, C l - L/(q=c)C m moment coefficient, C,. --M/(q=c 2)C e.,is

Aircraft surface contamination sensing system using control surface hinge moment measurements

Abstract: A method and system for sensing surface contamination on an aircraft having a control surface. A control element including the control surface is connected to the aircraft by a hinge. The method and system sense a control surface hinge moment about a line on the hinge of the aircraft. A control surface steady hinge moment coefficient is calculated from control surface hinge moment data representing the hinge moment over a period of time. An unsteady hinge moment is calculated which is dependent on the control surface steady hinge moment coefficient. By using a value which is relatively constant during uncontaminated surface conditions, this value can be compared against calculated values to check for variations. If the calculated unsteady value varies from the known uncontaminated values, a warning may be given, alerting an operator of unsafe conditions. This warning is given before a critical value is reached, allowing the operator a period of time to react to the warning. Alternatively, flight control systems may be notified so they may take corrective action.

Three-dimensional vortex interactions with a stationary blade

Abstract: Surface pressure measurements were obtained during a three-dimensional vortex interaction with a NACA 0015 aerofoil. The upper and lower surfaces of the blade experienced different aerodynamic loads which appear to be controlled by the impact of the vortex axial core flow on the blade surface. On the upper surface of the blade, where the vortex core flow was away from the aerofoil, the interaction was characterised by the generation of a suction peak. On the lower surface, where the axial component was towards the blade, a pressure pulse developed and seemed to be influenced by the vortex approach angle. These features resulted in rapid changes in normal force and quarter chord pitching moment during the interaction. This impulsive loading of the blade may provide some explanation for sound generation and control degradation problems associated with the tail rotor of helicopters.

Freestream Disturbance Effects on an Airfoil Pitching at Constant Rate

Abstract: Effects of freestream disturbances on the dynamic stall process of an NACA 0012 airfoil undergoing ramp-type pitching motion from 0-30 deg were studied using measurements of unsteady surface pressures. A thin circular cylinder was located upstream of the test airfoil and was offset vertically between -0.133 and 0.133 chord lengths with respect to the airfoil pitching axis to provide the freestream disturbances. The measurements were carried out at a constant chord Reynolds number of 8 x 10 4 for a relatively low reduced pitch-rate range k = 0.01-0.04. As compared with the undisturbed flow case, the imposed freestream disturbances were able to produce a significant increase in absolute magnitude of the peak suction pressure near the leading edge, resulting in a noticeable phase delay in stall process. The growth of the leading-edge suction peak caused by insertion of the upstream rod was found to be sensitive to its vertical offset position. Nevertheless, the pressure variation caused by the freestream disturbances was appreciable only in the forward 20% chord from the leading edge, bringing a lower effect on the airfoil lift and momentum coefficients

DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

Abstract: To quantify the benefits of smart materials and structures adaptive wing technology, Northrop Grumman Corp. (NGC) built and tested two 16 percent scale wind tunnel models (a conventional and a smart model) of a fighter/attack aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Pahase 1. Performance gains quantified included increased pitching moment (CM), increased rolling moment (Cl) and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist effected by SMA torque tube mechanisms, compared to conventional hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center''s (LaRC) 16ft Transonic Dynamic Tunnel (TDT) in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12 percent increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10 percent increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

Sensing aircraft icing effects by flap hinge moment measurement

Abstract: Simulated la&-droplet ice accretions were tested on a NACA 23012 airfoilwith simple flap in the UlUC subsonic wind tunnel: Reduced ‘lift, increased drag, chsnges in pitching moment and hinge moment were measumdduetotheicesimulation Theseresultedfioma separation bubble thatdimmed behind the simulated ice, upward force imposed on the tlap that acts to deflect the flap in that direction. , The flap is essentially sucked upward by the lower pressure. Due to the increased lift force on the control surface a negative hinge moment also occurs. This abmpt~ice-induceo flow separation can lead to a sudden signiticant change in hinge moment leaving insufficient time for the pilot to react correctly. Such occmences have led to aircmft accidents in the East. Thq, it is desirable to sense impending problems before they occur, in order to’warn the pilot and potentially alter ,-the control system.’ ” A better understanding’ of this ’ !’ ihenomenon is re@idm@I& fluctuation.of the flap hinge nmmeiltwacpanm*al&ggyg!~ an& ximmdkd ,by a RMS a’maximumvalue ator near msximumlift. Thecurrentresearchrelatesthisunsteady parameter to the steady-st@e aerodynamic coefficients in addition to the flow characteristics associated with the separation bubble. As opposed to the steady-state value, the chsnge in the RMS hinge moment was observed duringtheline.arphaseoftheliftcurveseveraldegrees before stall.

Correlation Between Geometric Similarity of Ice Shapes and the Resulting Aerodynamic Performance Degradation: A Preliminary Investigation Using WIND

Abstract: Aerodynamic performance calculations were performed using WIND on ten experimental ice shapes and the corresponding ten ice shapes predicted by LEWICE 2.0. The resulting data for lift coefficient and drag coefficient are presented. The difference in aerodynamic results between the experimental ice shapes and the LEWICE ice shapes were compared to the quantitative difference in ice shape geometry presented in an earlier report. Correlations were generated to determine the geometric features which have the most effect on performance degradation. Results show that maximum lift and stall angle can be correlated to the upper horn angle and the leading edge minimum thickness. Drag coefficient can be correlated to the upper horn angle and the frequency-weighted average of the Fourier coefficients. Pitching moment correlated with the upper horn angle and to a much lesser extent to the upper and lower horn thicknesses.

Stall Resistance Features of Lifting-Body Airplane Configurations

Abstract: A generic lifting-body airplane model was tested in a lowspeed wind tunnel. The experimental data indicate that at lower angles of attack the lift-over-drag ratio is comparable to other high-efficiency designs. The high angle of attack aerodynamics of this configuration is influenced by the side-edge vortex system observed above the aft section of the lifting fuselage. Consequently, the total lift of the airplane model continues to increase beyond the angle of wing stall, accompanied by increasing nose-down pitching moments. In principle, such characteristics allow the tailoring of the configuration lift and pitching moment in a manner that lift-loss effects beyond wing stall are minimal. Furthermore, the sharp increase in the nosedown moment of the fuselage can be positioned within the angle-of-attack performance curve such that airplane stall can become unreachable (and the configuration becomes stall resistant). The present study investigates some of the geometrical parameters influencing these aerodynamic effects so that such inherent stall-resistant characteristics can be developed early in the design stage of a lifting-body airplane configuration.

Hypersonic flow control using upstream focused energy deposition

Abstract: A numerical study of centerline and off-centerline power deposition at a point upstream of a two-dimensional blunt body at Mach 6.5 at 30 km altitude are presented. The full Navier-Stokes equations are used. Wave drag, lift, and pitching moment are presented as a function of amount of power absorbed in the flow and absorption point location. It is shown that wave drag is considerably reduced. Modifications to the pressure distribution in the flow field due to the injected energy create lift and a pitching moment when the injection is off-centerline. This flow control concept may lead to effective ways to improve the performance and to stabilize and control hypersonic vehicles.

Analysis of the scaling effects for missile configuration with lateral thruster

Abstract: Steady-state Computational Fluid Dynamic (CFD) analyses using GASP version 3 flow solver has be431 pezformed for a missile configuration with side thruster at given full scale f&e flight~conditions and quarter scale wind tunnel model conditic~ls with similar computaticmal grids. The given jet exit conditions for both cases were required to have the same thrust az&icients, which is an empirical scaling law used in the jet interaction (JI) cmmuui~. The purpose of this analysis is to examine aud compare the integrated axial and normal forces and pitching moment obtained fi-om the two computations to determine if they are indeed identical. It was found that the axial aud ncmnal force coefficients differ by 10 and 3 percent, respectively, and that the pitching moment coefficients vary by 28 percent. The differences in axial force between the two cases is attributed to the difference of Reynolds numbs between them and thecause of pitchiig moment diffezence is due the Closeness of the pitching and the pressure centers. The scaling law of constant thrust coefficient is satisfactory in present study. However, in a separated study, it was found that in addition to keep the thrust coefficients the same, one must also keep the jet Machnumbers the same for both cases in or& to obtain correct andbest sin&&y.

Wind Tunnel Investigation of Joined Wing Configurations

Abstract: : The Air Force Research Laboratory's Munitions Directorate is looking to extend the range of its small smart bomb. One proposed idea is to retrofit the bombs with a wing kit, particularly a joined wing configuration. A typical joined wing configuration is one where the wings are positioned in such a way that they form a diamond in both plan and front views. The purpose of this study is to conduct low speed wind tunnel testing of the joined wing configuration to help determine if the joined wing is more beneficial than a single wing configuration. Configurations with differing sweep angles and tip interconnects will be tested in the AFIT 5' wind tunnel. The lift, drag, and pitching moment coefficients will be ascertained. All researched literature indicates that certain joined wing configurations outperform its single wing counterpart.

A review of European code-validation studies in high-enthalpy flow

Abstract: The paper focuses on the experiments performed in the European high–enthalpy facilities during the crewed space flight hypersonic ground testing technology programme. Emphasis is placed on simple configuration testing enabling validation of the physical modelling within non–equilibrium Navier–Stokes codes. The high–enthalpy facilities used are the ONERA hot shot F4 and the DLR Stalker tube HEG. The simple configurations are the Nozzle, the Electre blunt code, the hyperboloid flare, and the 70 degree blunt cone. In addition to these axisymmetric configurations, the Halis/Orbiter configuration was extensively tested and numerically computed to validate the ground to flight extrapolation methodology. Real gas effects on aerodynamic forces and in particular on pitching moment for the Halis have been reproduced in the F4 and compared with flight. The paper concludes with a series of recommendations for additional improvements in nozzle calibrations using non–intrusive methods and for an increased use of CFD for the design of experiments, the definition of the test environment, the interpretation of the results and finally the extrapolation to flight.

An Inviscid Computational Study of an X-33 Configuration at Hypersonic Speeds

Abstract: This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of a simplified X-33 configuration. The major components of the X-33 vehicle, namely the body, the canted fin, the vertical fin, and the bodyflap, were simulated in the CFD model. The rear-ward facing surfaces at the base including the aerospike engine surfaces were not simulated. The FELISA software package consisting of an unstructured surface and volume grid generator and two inviscid flow solvers was used for this study. Computations were made for Mach 4.96, 6.0, and 10.0 with perfect gas air option, and for Mach 10 with equlibrium air option with flow condition of a typical point on the X-33 flight trajectory. Computations were also made with CF4 gas option at Mach 6.0 to simulate the CF4 tunnel flow condition. An angle of attack range of 23 to 48 degrees was covered. The CFD results were compared with available wind tunnel data. Comparison was good at low angles of attack; at higher angles of attack (beyond 25 degrees) some differences were found in the pitching moment. These differences progressively increased with increase in angle of attack, and are attributed to the viscous effects. However, the computed results showed the trends exhibited by the wind tunnel data.

Experimental study on dynamic characteristics of delta wing in pitching motion

Abstract: The Crst purpose of the research is to make a simple, comiact and inexpensive force balance, which will be improved and be applied to measure aerodynamic force in undesirable or severe condition like in very high temperature flow. The second is to investigate dynamic characteristics of aerodynamic forces and moment for a delta wing in pitching motion experimentally with the developed force balance, which is compared with static results. The third puqiose is to study aerodynamic characteristics of the pitching delta wing near the gmund and compared with those out of ground effect. A force balance in the shape of a ring to measure lift, drag .and pitching moment was developed, and was found to get these forces and moment witbin the error of about 7.5%. Existence of the ground fasten a stall and increases a maximum lift coefficient by 24% compared with tho& out of ground effect. Derivative of the measured pitching moment coefficient with respect to attack angle indi+tes that longitudinal static stability increases owing to the ground effect on the delta wing in pitching motion around at a stall angle of attack.

Gurney flap aerodynamic unsteadiness

Abstract: A Gurney flap is a thin strip of material attached at the trailing edge of the upper surface of a racing car wing in order to increase the downforce of a wing of limited size. The flow around a single element racing car wing with and without a 4.7% Gurney flap has been investigated experimentally and computationally. Attention has been concentrated on the unsteady wake which is characterized by alternate vortex shedding as is observed behind circular cylinders, flat plates and other two-dimensional bluff bodies. Using novel experimental and postprocessing techniques it has been possible to determine the instantaneous velocity and pressure fields in the intermediate and far wake. The fluctuating pressures on the surface of the aerofoil have been measured and integrated to determine the fluctuating lift and pitching moment on the wing. These were found to be small compared with the time-averaged values despite the strong fluctuations in the wake. Steady-state computational fluid dynamic simulations were performed and the aerofoil surface pressure distribution was predicted accurately. The relationship between wake unsteadiness and the pressure acting on the rear of bluff bodies (the base pressure) is discussed and its importance to Gurney flaps is assessed.

The decelerator pitch-dependent performances prediction based on 3D aeroelastic analysis

Abstract: The results of numerical study of decelerators in flow are presented for varying incident (pitch) angles. The problem was solved in 3D aeroelastic approach coupling structural and CFD codes. The capabilities of integrated PC software MONSTR are demonstrated for different types of decelerators: ring parachute, disc + conic skirt parachute, cross parachute and ballute in wide range of pitch angles including subcritical and critical angles causing the canopy collapse. The main aerodynamic performances predicted (the canopy shape, pressure distribution, drag and moment coefftcients Cx, Cy, Cm) are compared with data from experiments in wind tunnels. Nomenclature A incidence (pitch) angle [degrees], At, ; static stability angle A: Cm(Ab)=O and dCmldA~0, Cp pressure coefficient, ACp pressure difference coefficient, M Mach number, Cx -coeff. of parachute drag in downstream direction, Cy appropriate coefficient for normal force Cm pitching moment coeff. ref. to confluence poin& Cxa coefficient of drag along canopy axis, Cya appropriate coefficient for normal force, CD full drag coeff. = (Cx’+C#’ = (Cxaz+Cyaz)‘n D decelerator size (diameter of axisymmetric canopy or ballute, span of cross canopy blades) [ml, F parachute reference area [sq.m], H conic skirt width [m], Kp canopy porosity coefficient = (vents area) / F, L line length [ml, N number of suspension lines, S undimensioned coord. along canopy span [ -1; +l], W fabric permeability [cub.m/sq.m/s] in standard test. Introduction The inflatable type of decelerator/stabilizer designed to produce high drag force can be usually Copyright 8 1999 by Y.Mosseev. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission * Ph.D., Head of R&D Department considered as a bluff body from aerodynamic point of view. The flow separation from canopy edges and outer surface, recirculation zone with large scale vortices in near wake behind the canopy, high movability of its surface and related “attached” fluid mass all these factors can cause not only flow instability as a whole, but also pitch instability of decelerator, its oscillation and swinging of payload. It is known the parachute stability is affected by canopy shape, fabric permeability and canopy porosity (vents, slots), line length. A decelerator is statically stable when Cm(A)=0 (necessary condition) and dCmldA ~0 (sufficient condition). Evidently, all decelerators with at least 2-fold symmetry have Cm(O)=0 but only some of them fit the sufficient static stability condition. Many parachutes with solid canopy has nonzero balancing pitch angle Ab #O and this in turn causes gliding effect even for axisymmetric parachutes. So, circular parachutes (]A&30”...43”) glides at about 30”-40” in unpredictable directions. As it has very low pitching moment near balancing angles, dynamic disturbances can easily force parachute to change the sign of At, and swinging begins with large amplitude. Sometimes the parachute coning oscillation are noticed. Moreover, if disturbance increases ]A] above ]AtJ#O by 5” or more the dip on lateral canopy surface arises, growths and can cause a collapse. As a result drag force drops and some time needed to reinflate the canopy. In order to provide full featured flight dynamics analysis for parachute+payload system not only basic aerodynamic performances Cx, Cy, Cm must be defined, but also their derivatives with respect to time. The coefficients are responsible for static stability (balancing angle, gliding effect, full drag Cn value) and derivatives for dynamic one (oscillation amplitudes, dampening characteristics, consequences of external disturbances). As the flight test has the dynamic nature and still not too informative an identification of static values Cx, Cy, Cm and appropriate derivatives with respect to time depending on “A” with sufficient accuracy seems to be questionable. For this reason the basic aerodynamic performances responsible for static stability are used to defined in

AeroForce - Thrust/Drag Bookkeeping and Aerodynamic Force Breakdown over Components

Abstract: Aeroforce is an aerodynamic data postprocessing tool that permits a more detailed analysis of the location of aerodynamic forces. For complex configurations the force evaluation of an arbitrary split-up of the entire aerodynamic surface into components like body, wing, pylon or engine is supported. Aeroforce calculates forces and aerodynamic coefficients for these components directly from the pressure and friction distributions.

Unsteady Characteristics of Lift and Drag Acting on a Pitching Airfoil.

Abstract: In the present study the dynamic lift and drag acting on the pitching airfoils such as a flat plate, NACA 0010, NACA 0020, NACA 65-0910 and BTE have been measured by using a six-axes sensor in a water tunnel at low Reynolds number region under the condition of various pitching frequencies. The unsteady characteristics of the dynamic lift and drag have been compared with the quasi-steady ones which are measured under the stationary condition. The pitching motion is available for keeping the lift higher after the separation occurs. The characteristics of the dynamic lift is quite different from the quasi-steady one at high pitching frequency region such as 3.0 Hz. As the pitching frequency decreases, the amplitude of the dynamic lift becomes closer to the quasi-steady one. However, the phase remains different between the steady and unsteady conditions even at low pitching frequency such as 0.05 Hz. On the other hand, the dynamic drag is governed strongly by the angle of attack.

Modeling the High Speed Research Cycle 2B Longitudinal Aerodynamic Database Using Multivariate Orthogonal Functions

Abstract: The data for longitudinal non-dimensional, aerodynamic coefficients in the High Speed Research Cycle 2B aerodynamic database were modeled using polynomial expressions identified with an orthogonal function modeling technique. The discrepancy between the tabular aerodynamic data and the polynomial models was tested and shown to be less than 15 percent for drag, lift, and pitching moment coefficients over the entire flight envelope. Most of this discrepancy was traced to smoothing local measurement noise and to the omission of mass case 5 data in the modeling process. A simulation check case showed that the polynomial models provided a compact and accurate representation of the nonlinear aerodynamic dependencies contained in the HSR Cycle 2B tabular aerodynamic database.

A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3

Abstract: A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).

Aerodynamic experimentation with ducted models as applied to hypersonic air-breathing vehicles

Abstract: A methodology of experimentation in high supersonic wind tunnels for studying aerodynamic characteristics of hypersonic flying vehicles powered by air-breathing engines is discussed. Investigations of such total aerodynamic forces as drag, lift and pitching moment at testing the models are implicit when the air flow through the model ducts is accomplished so that to provide the simulation of the external flow around the airplane and flow over the inlets, but the operating engines and, hence, the exhaust jets are not modeled. The methods used for testing such models are based on the measurement of duct stream parameters alongside with the balance measurement of aerodynamic forces acting on the models. In the tests, aerometric tools are used such as narrow metering nozzles (plugs), pitot and static pressure probes, stagnation temperature probes and pressure orifices in walls of the model duct. The aerometric data serve to determine the flow rate and momentum of the stream at the duct exit. The internal non-simulated forces of the model ducts are also determined using the conservation equations for energy, mass flow and momentum, and these forces are eliminated from the aerodynamic test results. The techniques of the said model testing have been well developed as applied to supersonic aircraft, however their application for hypersonic vehicles whose models are tested at high supersonic speeds, Mach number M∞>4, implies some specific features. In the present paper, the results of experimental and theoretical study of these features are discussed. Some experimental data on aerodynamics of hypersonic aircraft models received in methodological tests are also presented. The tunnel experiments have been carried out in the Mach number range M∞=2–6.