Lift enhancement and flow structure of airfoil with joint trailing-edge flap and Gurney flap

The impact of Gurney flaps (GF), of different heights and perforations, on the aerodynamic and wake characteristics of a NACA 0015 airfoil equipped with a trailing-edge flap (TEF) was investigated experimentally at Re = 2.54 × 10⁵. The addition of the Gurney flap to the TEF produced a further increase in the downward turning of the mean flow (increased aft camber), leading to a significant increase in the lift, drag, and pitching moment compared to that produced by independently deployed TEF or GF. The maximum lift increased with flap height, with the maximum lift-enhancement effectiveness exhibited at the smallest flap height. The near wake behind the joint TEF and GF became wider and had a larger velocity deficit and fluctuations compared to independent GF and TEF deployment. The Gurney flap perforation had only a minor impact on the wake and aerodynamics characteristics compared to TEF with a solid GF. The rapid rise in lift generation of the joint TEF and GF application, compared to conventional TEF deployment, could provide an improved off-design high-lift device during landing and takeoff.     

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

 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. Received: 25 July 1996/ Accepted: 14 December 1998     

Investigation of transonic and supersonic flow past a pair of bodies

Transonic and supersonic flows past a pair of bodies have been experimentally investigated. The leading bodies were spheres, cylinders, and cones, while the trailing bodies were flat-ended circular cylinders. The leading and trailing bodies were joined by cylindrical rods of various lengths, aligned with the axis of symmetry. For these models, the pattern of flow between the bodies and the Mach number dependence of the drag coefficientC _x were determined in the acceleration and deceleration flow regimes in a wind tunnel. The experimental results are used to analyze the properties of the flow between the bodies and the variation of the aerodynamic coefficients of the models. The reasons for the hysteresis in the behavior of the coefficients in the acceleration and deceleration stages are discussed. The influence of the shape and dimensions of the leading body on the modelC _x is evaluated. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 2, pp. 158–164, March–April, 1998.     

Hypersonic interference heating in the vicinity of surface protuberances

The understanding of the behaviour of the flow around surface protuberances in hypersonic vehicles is developed and an engineering approach to predict the location and magnitude of the highest heat transfer rates in their vicinity is presented. To this end, an experimental investigation was performed in a hypersonic facility at freestream Mach numbers of 8.2 and 12.3 and Reynolds numbers ranging from Re _∞/m = 3.35 × 10⁶ to Re _∞/m = 9.35 × 10⁶. The effects of protuberance geometry, boundary layer state, freestream Reynolds number and freestream Mach numbers were assessed based on thin-film heat transfer measurements. Further understanding of the flowfield was obtained through oil-dot visualizations and high-speed schlieren videos. The local interference interaction was shown to be strongly 3-D and to be dominated by the incipient separation angle induced by the protuberance. In interactions in which the incoming boundary layer remains unseparated upstream of the protuberance, the highest heating occurs adjacent to the device. In interactions in which the incoming boundary layer is fully separated ahead of the protuberance, the highest heating generally occurs on the surface just upstream of it except for low-deflection protuberances under low Reynolds freestream flow conditions in which case the heat flux to the side is greater.     

Experimental investigation of the influence of the flow structure on the aerodynamic coefficients of the IXV vehicle

In the framework of the ESA Future Launchers Preparatory Program (FLPP) an experimental study on the aerodynamic behavior during the re-entry phase of the Intermediate eXperimental Vehicle (IXV) configuration was conducted in the DLR hypersonic wind tunnel H2K in Cologne. Tests were carried out at Mach 6.0 and 8.7 with different flap deflection angles and the angle of attack varied continuously between 20° and 55° to investigate the flow topology as well as the aerodynamic forces and moments and the surface pressure distribution. The experimental data show that depending on the combination of the flap deflection angle (δ _L/R) and angle of attack (α) the complex flow structure in the vicinity of the flaps significantly influences the vehicle’s aerodynamic coefficients. An analysis of this shock/shock and shock/boundary layer interaction causing flow separation with reattachment is performed.     

Numerical analysis on flutter of Busemann-type supersonic biplane airfoil

The Busemann-type supersonic biplane can effectively reduce the wave drag through shock interference effect between airfoils. However, considering the elastic property of the wing structure, the vibration of the wings can cause the shock oscillation between the biplane, which may result in relative aeroelastic problems of the wing. In this research, fluid–structure interaction characteristics of the Busemann-type supersonic biplane at its design condition have been studied. A theoretical two-dimensional structure model has been established to consider the main elastic characteristics of the wing structure. Coupled with unsteady Navier–Stokes equations, the fluid–structure dynamic system of the supersonic biplane is studied through the two-way computational fluid dynamics/computational structural dynamics (CFD/CSD) coupling method. The biplane system has been simulated at its design Mach number with different nondimensional velocities. Different initial disturbance has been applied to excite the system and the effects of the position of the mass center on the system’s aeroelastic stability is also discussed. The results reveal that the stability of the airfoil in supersonic biplane system is decreased compared with that of the airfoil isolated in supersonic flow and such stability reduction effect should be given due attention in practical design.     

Development of an Intermittency Equation for the Modeling of the Supersonic/Hypersonic Boundary Layer Flow Transition

An intermittency transport equation is developed in this study to model the laminar-turbulence boundary layer transition at supersonic and hypersonic conditions. The model takes into account the effects of different instability modes associated with the variations in Mach numbers. The model equation is based on the intermittency factor γ concept and couples with the well-known SST k–ω eddy-viscosity model in the solution procedures. The particular features of the present model approach are that: (1) the fluctuating kinetic energy k includes the non-turbulent, as well as turbulent fluctuations; (2) the proposed transport equation for the intermittency factor γ triggers the transition onset through a source term; (3) through the introduction of a new length scale normal to wall, the present model employs the local variables only avoiding the use of the integral parameters, like the boundary layer thickness δ, which are often cost-ineffective with the modern CFD methods; (4) in the fully turbulent region, the model retreats to SST model. This model is validated with a number of available experiments on boundary layer transition including the incompressible, supersonic and hypersonic flows past flat plates, straight/flared cones at zero incidences, etc. It is demonstrated that the present model can be successfully applied to the engineering calculations of a variety of aerodynamic flow transition with a reasonably wide range of Mach numbers.     

Hysteresis in transverse galloping: The role of the inflection points

Transverse galloping is here considered as a one-degree-of-freedom oscillator subjected to aerodynamic forces, which are described by using the quasi-steady hypothesis. The hysteresis of transverse galloping is also analyzed. Approximate solutions of the model are obtained by assuming that the aerodynamic and damping forces are much smaller than the inertial and stiffness ones. The analysis of the approximate solution, which is obtained by means of the method of Krylov–Bogoliubov, reveals the existing link between the hysteresis phenomenon and the number of inflection points at the aerodynamic force coefficient curve, C_y(α); C_y and α being, respectively, the force coefficient normal to the incident flow and the angle of attack. The influence of the position of these inflection points on the range of flow velocities in which hysteresis takes place is also analyzed.     

Freestream effects on base pressure development of an annular plug nozzle

An experimental investigation has been carried out to study the effect of freestream flow and cowl-length variation on (i) upstream flow interference effects and (ii) the base wake-closure nozzle pressure ratio. It is observed that for supersonic freestream Mach numbers the nozzle exhaust seems to only slightly influence the upstream interference effects for M = 1.2 but shows significant influence for M = 1.6. Increasing the cowl-length further reduces the upstream flow interference effects significantly. Further, the reduced momentum thrust from the inner nozzle in the presence of freestream for similar nozzle pressure ratio (relative to static tests) delays the downstream movement of the system of shocks on the plug surface. In the case of the plug truncated at 40% length, this delays the onset of base-wake closure and hence, increases the base-wake closure nozzle pressure ratio with increasing freestream Mach number. Increasing the cowl-length also helps to increase the base pressure thrust contribution at all operating conditions.     

滑动蒙皮变后掠气动力非定常滞回与线性建模

针对低速不可压条件下滑动蒙皮方式变后掠过程中非定常动态气动特性开展了3方面的研究工作: (1)飞行器变形过程中非定常动态气动特性风洞试验技术研究;(2)变形过程中滞回效应研究和机理分析; (3)基于风洞试验结果开展变形过程中非定常动态气动力线性建模. 初步研究表明: (1)采用强迫振荡法可以有效地获取变形过程中非定常动态气动力滞回效应; (2)造成变形过程中气动滞回效应的机理有两个即``动边界效应'和``流场结构滞回效应', 造成滞回效应的机理可能主要在于后者; (3)引入升力系数和俯仰力矩系数随后掠角变化率的动导数概念, 可以建立变形过程中非定常动态气动力线性模型.      

Aerodynamic characteristics of a body of revolution with gas-permeable surface areas

Results of experimental studies and numerical calculations of aerodynamic characteristics of a supersonic flow around a body of revolution with a gas-permeable porous nose cone and an internal duct are presented. At a flow velocity corresponding to the Mach number M = 3, the body considered is found to have a lower drag coefficient (approximately by 9%) than a similar body impermeable for the gas and a lower streamwise static stability.     

Three-dimensional steady supersonic duct flow using Lagrangian formulation

In this paper, numerical simulation of three-dimensional supersonic flow in a duct is presented. The flow field in the duct is complex and can find its applications in the inlet of air-breathing engines. A unique streamwise marching Lagrangian method is employed for solving the steady Euler equations. The method was first initiated by Loh and Hui (1990) for 2-D steady supersonic flow computations and then extended to 3-D computation by the present authors Loh and Liou (1992). The new scheme is shown to be capable of accurately resolving complicated shock or contact discontinuities and their interactions. In all the computations, a free stream of Mach numberM=4 is considered.This article was processed using Springer-Verlag T_EX Shock Waves macro package 1.0 and the AMS fonts, developed by the American Mathematical Society.     

Existence and Stability of the Inner Boundaries of the Domain of Multiple Hysteresis of Static Aerodynamic Forces and Moments

The results of experimental investigations of the multiple static hysteresis of the aerodynamic characteristics of a rectangular high-aspect-ratio wing are presented. Schematic wing-flow structure patterns, time dependences of the coefficients c _y(t), m _z(t), and m _x(t) and their frequency spectra obtained for a fixed model are given for different boundaries of the hysteresis domain. The time dependences of aerodynamic forces and moments are analyzed at angles of incidence at which sharp changes are observed. It is shown that the static hysteresis can be described by a mathematical model used in catastrophe theory.     

The hypersonic Mach number independence principle in the case of viscous flow

The hypersonic Mach number independence principle of Oswatitsch is important for hypersonic vehicle design. It explains why, above a certain flight Mach number (M _∞ ≈ 4−6, depending on the body shape), some aerodynamic properties become independent of the flight Mach number. For ground test facilities this means that it is sufficient for the Mach number in the test section to be high enough, that Mach number independence exists. However, the principle was derived for calorically perfect gas and inviscid flow only. In this paper a theoretical study for blunt bodies in the case of viscous flow is presented. We provide numerical results which give insight into how attached viscous flow behaves at high Mach numbers. The flow past an axisymmetric configuration is analysed by applying a coupled Euler/second-order boundary-layer method. Wall boundaries are treated by assuming an adiabatic or radiation-adiabatic wall for laminar flow. Calorically perfect or equilibrium air is accounted for. Lift, drag, and moment coefficients, and lift-to-drag ratios are given for several combinations of flight Mach number and altitude, i.e. Reynolds number. For blunt bodies considered here, which are pressure dominated, Mach number independence occurs for the adiabatic wall, but not for the radiation-adiabatic wall assumption.     

Outflow of an unexpanded jet into counter supersonic and subsonic flows

The article gives the results of an experimental investigation of the geometric structure of an opposing unexpanded jet. It discusses flow conditions with interaction between the jet and sub- and supersonic flows. It is shown that, with the outflow of an unexpanded jet counter to a supersonic flow, there are unstable flow conditions. For stable flow conditions with one roll, dependences are proposed determining the form of a jet in a supersonic opposing flow. A generalized dependence is obtained for the distribution of the pressure at the surface of a body with a jet, flowing out counter to a subsonic flow. The range of change in the determining parameters are the following: Mach numbers at outlet cross section of nozzle, M _a = 1 and 3; Mach numbers of opposing flow, M = 0.6–0.9 and 2.9; degree of effectiveness of jet, n = p _a /p = 0.5–800 (p _a and p are the static pressures at the outlet cross section of the nozzle and in the opposing flow); the ratios of the specific heat capacities, _a = = 1.4; the drag temperatures of the jet and the flow, T_o = T_oa = 290°K.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 1, pp. 89–96, January–February, 1977.     

Aerodynamic forces and flow fields of a two-dimensional hovering wing

This paper reports the results of an experimental investigation on a two-dimensional (2-D) wing undergoing symmetric simple harmonic flapping motion. The purpose of this investigation is to study how flapping frequency (or Reynolds number) and angular amplitude affect aerodynamic force generation and the associated flow field during flapping for Reynolds number (Re) ranging from 663 to 2652, and angular amplitudes (α _A) of 30°, 45° and 60°. Our results support the findings of earlier studies that fluid inertia and leading edge vortices play dominant roles in the generation of aerodynamic forces. More importantly, time-resolved force coefficients during flapping are found to be more sensitive to changes in α _A than in Re. In fact, a subtle change in α _A may lead to considerable changes in the lift and drag coefficients, and there appears to be an optimal mean lift coefficient around α _A = 45°, at least for the range of flow parameters considered here. This optimal condition coincides with the development a reverse Karman Vortex street in the wake, which has a higher jet stream than a vortex dipole at α _A = 30° and a neutral wake structure at α _A = 60°. Although Re has less effect on temporal force coefficients and the associated wake structures, increasing Re tends to equalize mean lift coefficients (and also mean drag coefficients) during downstroke and upstroke, thus suggesting an increasing symmetry in the mean force generation between these strokes. Although the current study deals with a 2-D hovering motion only, the unique force characteristics observed here, particularly their strong dependence on α _A, may also occur in a three-dimensional hovering motion, and flying insects may well have taken advantage of these characteristics to help them to stay aloft and maneuver. An erratum to this article can be found at     

Shock wave triple-point morphology

The existence and characteristics of shock wave triple points are examined. The analysis utilizes a single flow plane for the three shocks and is local to the triple point. It applies when the flow is unsteady, three-dimensional, and the upstream flow is nonuniform. Under more restrictive conditions, a relation is also derived for the ratio of the curvature of the Mach stem to that of the reflected shock. For given values of the ratio of specific heats, γ, and the upstream Mach number, M ₁, a solution window is established. A parametric set of solutions is generated within the window for γ = 1, 1.4, and 5/3 and for 16 values of M ₁ ranging from solution onset to M ₁ = 6.A solution can be one of three types, these stem from the velocity tangency condition along the slip stream. Topics are addressed such as solution multiplicity, shock wave and slip stream orientation, the nature of the reflected wave (weak, strong, inverted, normal), the nature of the Mach stem (weak, strong, normal), and differences due to changes in γ and M ₁.     

Experimental study of the reverse flow in the forward separation region in a pulsating flow around a spiked body

The flow with a free-stream Mach number M _∞ = 6 around a cylindrical body with a sharp spike is studied. The existence of a supersonic reverse flow for one of the phases of the pulsating flow regime is experimentally validated. A range of spike lengths is determined, which ensures a region of a supersonic reverse flow near the side surface of the spike. The time of existence of the supersonic reverse flow region is shown to be 0.15 of the period of pulsations if the spike length equals the model diameter. __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 48, No. 4, pp. 30–39, July–August, 2007.     

Contouring the nozzles producing a uniform supersonic flow or a thrust maximum in the presence of a curvilinear sonic line

Within the framework of the ideal, i.e., inviscid and non-heat conducting, gas model we consider the problem of designing the supersonic section of a two-dimensional or axisymmetric nozzle realizing a uniform supersonic flow limitingly similar with a sonic flow when the choked flow involves a curvilinear sonic line. Emphasis is placed on nozzles with abruptly or steeply converging subsonic sections and a strongly curved sonic line formed by the C ⁻-characteristics of the expansion fan with the focus at the lower bend point of the vertical section of the subsonic contour. In the two-dimensional case, the least possible greater-than-unity Mach number M _em at the nozzle exit corresponds to the flow in which the first intersection of the C ⁺-characteristics originated at the closing C ⁻-characteristic of the expansion fan falls on the unknown contour of its supersonic part. For a uniform flow with M _e < M _em the intersection of C ⁺-characteristics beneath the unknown contour make impossible its construction. A part of the contour realizing a uniform flow with M _em > 1 ensures a limitingly rapid flow acceleration and forms the initial region of the supersonic generator of a maximum-thrust nozzle. For this reason, in the case of a curvilinear sonic line the supersonic generators of these nozzles have two, rather than one, bends, which, however, is interesting only for the theory. At least, in the calculated examples the thrusts of the nozzles with one and two bends differ only by a hundredth or even thousandth fractions of per cent.     

Numerical simulation of gap effect in supersonic flows

The gap effect is a key factor in the design of the heat sealing in supersonic vehicles subjected to an aerodynamic heat load. Built on S-A turbulence model and Roe discrete format, the aerodynamic environment around a gap on the surface of a supersonic aircraft was simulated by the finite volume method. As the presented results indicate, the gap effect depends not only on the attack angle, but also on the Mach number.