CALCULATION OF ROCKET NOSE FAIRING SHELLS AERODYNAMIC CHARACTERISTICS

In order to study the influence of elastic forward-swept wing (FSW) with single control surface, the computational fluid dynamics/computational structural dynamics (CFD/CSD) loose coupling static aero elastic numerical calculation method was adopted for numerical simulation. The effects of the elastic FSW with leading- or trailing-edge control surface on aero elastic characteristics were calculated and analysed under the condition of high subsonic speed. The result shows that, the deflection of every single control surface could change the aero elastic characteristics of elastic FSW greatly. Compared with the baseline model, when leading-edge control surface deflected up, under the condition of small angles of attack, the aerodynamic characteristics was poor, but the bending and torsional deformation decreased. Under the condition of moderate angles of attack, the aerodynamic characteristics was improved, but bending and torsional deformation increased; When leading-edge control surface deflected down, the aerodynamic characteristics was improved, the bending and torsional deformation decreased/increased under the condition of small/moderate angles of attack. Compared with the baseline model, when trailing-edge control surface deflected down, the aerodynamic characteristics was improved. The bending and torsional deformation increased under the condition of small angles of attack. The bending deformation increased under the condition of small angles of attack, but torsional deformation decreases under the condition of moderate angles of attack. So, for the elastic FSW, the deflection of trailing-edge control surface play a more important role on the improvement of aerodynamic and elastic deformation characteristics.

References 1 Moore, F. G., Hymer, T. C, and Mclnville, R. M., Improved Aeroprediction Code: Part I—Summary of New Methods and Comparison with Experiment, Naval Surface Warfare Center Dahlgren Division, NSWCDD/TR93/91, Dahlgren, VA, May 1993. Moore, F. G., Mclnville, R. M., and Hymer, T. C., Improved Aeroprediction Code: Part II—Computer Program User's Guide and Listing, Naval Surface Warfare Center Dahlgren Division, NSWCDD/TR-93/241, Dahlgren, VA, Aug. 1993. Moore, F. G., Hymer, T. C., and Mclnville, R. M., A Planar Nonlinear Missile Aeroprediction Code for all Mach Numbers, AIAA Paper 94-0026, Jan. 1994. Fidler, J. E., and Bateman, M. C., Methods for High Incidence Missile Design, Journal of Spacecraft and Rockets, Vol. 12, No. 3, 1975, pp. 162-168. Moore, F. G., and Mclnville, R. M., A New Method for Calculating WingAlone Aerodynamics to Angle of Attack 180 deg, Naval Surface Warfare Center Dahlgren Division, NSWCDD/TR-94/3, Dahlgren, VA, March 1994. Stallings, R. L., Jr., and Lamb, M., Wing-Alone Aerodynamic Characteristics for High Angles of Attack at Supersonic Speeds, NASA Technical Paper 1889, July 1981. Baker, W. B., Jr., Static Aerodynamic Characteristics of a Series of Generalized Slender Bodies with and without Fins at Mach Numbers from 0.6 to 3.0 and Angles of Attack from 0 to 180 deg, Arnold Engineering Development Center, TR-75-124, Vols. I and II, Tullahoma, TN, May 1976. Nielsen, J. N., Hemsch, M. J., and Smith, C. A., A Preliminary Method for Calculating the Aerodynamic Characteristics of Cruciform Missiles to High Angles of Attack Including Effects of Roll Angle and Control Deflections, Office of Naval Research, Report ONR-CR215-216,4F, Arlington, VA, Nov. 1977.

During spacecraft reentry, the aerodynamic characteristics of the reentry capsule with a sizeable obtuse head inverted cone (OBHIC) in a hypersonic state are the key factors in the design of controlled reentry trajectory, the design of trim angle of attack (AoA), and the accuracy of fall point of capsules. This paper establishes a numerical simulation algorithm of a hypersonic chemical nonequilibrium flow field by considering high-temperature real gas effects. The reentry capsule’s aerodynamic and flow field characteristics are predicted and analyzed. The component transport equations for chemical reactions are added to the Navier-Stokes equations, enabling precise modeling of gas molecule vibrational excitation and ionization processes. A half-model multi-block structured mesh is employed to construct the aerodynamic profile of the reentry capsule with OBHIC. The aerodynamic characteristics and pressure distribution of the reentry capsule were calculated and analyzed under the typical conditions of flight altitude of 40~80 km, flight velocity of 6~27 Ma, and AoA of 15°~24°. The results of the aerodynamic characteristics prediction show that the aerodynamic characteristics of the reentry capsule are little affected by the incoming flow velocity in a hypersonic state but significantly affected by the AoA. As the AoA increases, the lift coefficient CL increases linearly, the drag coefficient CD decreases linearly, the lift-to-drag ratio K rises linearly, and K increases from 0.19 to 0.29. In its hypersonic state, the reentry capsule exhibits a low lift-to-drag ratio, characterized by high drag and low lift coefficients. The results of this paper play an essential role in the reentry flight dynamics and reentry trajectory design of the reentry capsule. Moreover, it can further ensure the safety and reliability of the reentry flight of spacecraft.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

To determine the flight of a ski jumper it is essential to know what aerodynamic forces are acting on the ski jumper. However, few data on this are available, especially for a V-style ski jumping flight. We have measured the aerodynamic forces during the free flight phase for a V-style, as well as a parallel-style, ski jump by employing a full-size model in a wind tunnel. The aerodynamic force data, (drag, lift and pitching moment) were obtained to create an aerodynamic database. These forces are given in polynomial form as functions of the angle of attack, the body-ski (forward leaning) angle and the ski-opening (V-style) angle. Using the polynomial form database is a convenient way of obtaining the aerodynamic forces. Moreover, the wind tunnel was equipped with a ground effect plate to measure the aerodynamic forces during the landing phase. It was found that the difference between the lift with and without the ground effect plate increases with the ski-opening angle. The longitudinal stability in the pitching motion of a body-ski combination is also discussed on the basis of the pitching moment data. This indicates that a stable pitching oscillation of the body-ski combination may arise around an equilibrium point in the angle of attack, the trim angle of attack, during flight.

An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

Circulation control has been applied to airfoils since the late 1960's, and has been proven to change the aerodynamic performance by altering the interaction of the streamlines without changing the physical characteristics of the airfoil itself. This has many applications in fluid dynamics; the focus of this application is for the replacement of the conventional helicopter rotor blade system with a fly-by-wire, active circulation controlled system. Conventional helicopters use a swashplate and a series of mechanical linkages, bearings, and dampers to create a fully articulated rotor hub system. This system is required to achieve the blade characteristics required for stable flight. The need for such a system stems from the asymmetric lift developed in maneuvering flight conditions, which requires the angle of attack of the blades to be changed based on the rotational position within the rotor plane, also known as the azimuth angle. By alternatively activating blowing slots along the leading and trailing edges of the airfoil, the aerodynamic parameters (i.e. lift and drag) can be changed, effectively changing the angle of attack through streamline alteration thus eliminating the need for physical blade pitch changes.;Mathematical models/codes are used to model and simulate the complex blade dynamics of a full-scale rotorcraft. Many of these codes use a blade element method that separates the rotor into small segments and evaluates the aerodynamic characteristics of these elements as two-dimensional airfoils at different local conditions. These local conditions include, but are not limited to, extreme angles of attack, reverse flow, compressibility effects, dynamic effects, and other aerodynamic phenomena.;This research investigated the reverse flow aerodynamic characteristics of a 10:1 elliptical airfoil with various leading and trailing edge blowing pressures. The testing conditions for the aerodynamic investigation were: angle of attack [154 to 196 degrees]; blowing coefficient [0 to 0.009] and [0 to 0.014] for the leading and trailing edge blowing slots, respectively; and Mach number [0.073 and 0.109].;It is concluded that the potential exists for increasing forward flight speeds for helicopters using circulation control in the reverse flow regions of the helicopter rotor environment. Specifically, it is concluded that positively altering the aerodynamic characteristics, primarily the lift coefficient, in reverse flow, is possible through circulation control. Through this investigation, the general trends were found for the aerodynamic characteristics of a 10:1 circulation controlled elliptical airfoil in reverse flow. These trends led to the selection of blowing configurations to decrease the asymmetric loading condition based

Aerodynamic characteristics of a guided rocket projectile are numerically simulated. Simulations are performed in a range of flight conditions using a commercial CFD software FLUENT, and variations of lift coefficients and drag coefficients with Mach numbers and angles of attack were demonstrated. The numerical results also indicate that the angle of canard deflection has significant influence on the flow field near tailfins.

Analytical prediction of trajectories of ice pieces after release in an airstream

PurposeThe purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing.Design/methodology/approachSimulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference.FindingsIt was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs.Practical implicationsFor accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation.Originality/valueThe presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.

Long-range rotating wrap-around-fin rockets may exhibit non-convergent conical motion at high Mach numbers, causing increased drag, reduced range, and potential flight instability. This study employs the implicit dual time-stepping method to solve the unsteady Reynolds-averaged Navier–Stokes (URANS) equations for simulating the flow field around a high aspect ratio wrap-around-fin rotating rocket at supersonic speeds. Validation of the numerical method in predicting aerodynamic characteristics at small angles of attack is achieved by comparing numerically obtained side force and yawing moment coefficients with experimental data. Analyzing the rocket’s angular motion process, along with angular motion equations, reveals the necessary conditions for the yawing moment to ensure stability during angular motion. Shape optimization is performed based on aerodynamic coefficient features and flow field structures at various angles of attack and Mach numbers, using the yawing moment stability condition as a guideline. Adjustments to parameters such as tail fin curvature radius, tail fin aspect ratio, and body aspect ratio diminish the impact of asymmetric flow induced by the wrap-around fin on the lateral moment, effectively resolving issues associated with near misses and off-target impacts resulting from dynamic instability at high Mach numbers.

Aerodynamic characteristics of an Ejection Seat System at different angles of attack are studied by the numerical method and the flow mechanisms for such flows are carefully analyzed. The governing equations are Reynolds-averaged Navier-Stokes equations which are solved by the unstructured finite volume method. Upwind Osher scheme is used for spatial discretization and five-stage Runge-Kutta scheme is applied for temporal discretization. The DES model based on S-A one equation turbulence model is adopted. Parallel computation is based on the domain decomposition method and multi-block is achieved by using METIS system. The experimental data is used to validate this method. This research is helpful to understand the aerodynamic characteristics and flow mechanisms of Ejection Seat System at different angles of attack and Mach numbers, and can provide reasonable reference for Ejection Seat System design.

Euler's equations, the shock wave relations, and the surface boundary condition are used to derive a hypersonic slender body theory applicable to angles of attack from zero to 90 degrees. The theory is not based on small disturbances, but it gives Van Dyke's theory for small angles of attack, Sychev's similarity for moderate angles of attack, and strip theory for large angles of attack. The new similarity parameters are Meo~, 0 cos a/~, and y (where ~=o cos a + sin a). They are shown to be a combination of Sychev's parameters and can be used to modify equations for pressure and force coefficients at small angles of attack for moderate and large angle of attack applications. B(X,y,Z) CA CA CD CD CN CN Cp Cp d D Db A n p,p S(x,y,Z) T Nomenclature scaled function for body geometry axial force coefficient, based on ref. area ::: 2Ld scaled axial force coefficient given by eq. (50) drag coefficient, based on ref. area ::: 2Ld scaled drag coefficient given by eq. (51) normal force coefficient based on ref. area ::: 2Ld scaled normal force coefficient given by eq. (19) pressure coefficient, eq. (47) ... scaled pressure coefficient given by eq. (47) base height of lower surface, see Fig. 1 parameter defined by eq. (2) parameter defmed by eq. (44) unit vectors in x, y, z directions, respectively similarity parameter, 0 cos a~ similarity parameter, Moo~ body length freestream Mach number unit vector normal (inner) to surface static pressure and scaled static pressure, eq. (25) scaled function for shock surface temperature * Professor of Mechanical and Aerospace Engineering, Associate Fellow AlAA. Copyright© 1992 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. 542 u,v,w u,v,w ~ Veo,VOQ x,y,z x,y,z a ~ y o ~ Cartesian velocity components in x,y,z directions, respectively scaled (nondimensional) velocity components freestream velocity vector and magnitude Cartesian coordinates, see Fig. 1 x/L, y/d, z/d, respectively angle of attack shockwave angle ratio of specific heats, 1.4 for air lower surface thickness ratio, d/L, see Fig. 1 --'> lower surface thickness ratio relative to V 00' see Fig. 1 and eq. (15) p,p density and scaled density, eq. (22) circumferential angle, =0 in windward plane SUbscripts c Cone b body s shock wave 00 freestream value Introduction Analysis of inviscid, hypersonic flow past slender bodies has been developed by Van Dyke l for small angles of attack and by Sychev2 for large angles of attack. The small angle of attack analysis is based on small disturbance theory and the similarity parameters are Moo 0, a/o, and y. At large angles of attack the Sychev similarity parameters are Moo sin a, 0 cot a, and y. Hemsch3 showed that the large angle of attack similarity parameters can be combined to yield the small angle of attack parameters when a is small, i.e. (MOQ sin a) (0 cot a) --'> Moo 0 and 0 cot a --'> o/a. Cox and Crabtree4 and Hayes and Probstein5 discuss hypersonic small disturbance theory and Sychev similarity. BarnweU6 showed that Sychev's analysis is applicable to all slender-body flows with crossflow Mach numbers greater than sonic and hence is not restricted to hypersonic values of the cross flow Mach number. In addition, he showed that Sychev similarity applied to a number of slender-body flows with subsonic crossflow Mach numbers, including incompressible flow. Hemsch3 demonstrated empirically that Sychev similarity holds for any value of the cross-flow Mach number if the axial flow component is 4

This paper describes research undertaken to validate specific computational tools that predict the aerodynamic and, to a lesser extent, aerothermodynamic characteristics of blunt, capsule-type vehicles during atmospheric re-entry. To this end, several candidate capsules were considered, with varying side wall angles and symmetrical and asymmetrical heat shields. The results illustrate the correlations obtained from analytical predictions against actual test data collected through two separate wind tunnel tests that span the range of Mach numbers of interest. The research also assessed the effects of critical Outer Mould Line parameters on the aerodynamic performance of the vehicle, including its stability and control characteristics, and compared the results against the historical Apollo configuration. The candidate vehicles’ center of gravity envelopes for single trim attitudes during re-entry were also determined to evaluate their monopole stability tendencies. The results show that, for the configurations with symmetrical heat shields, an Apollo concept offered the best liftto-drag ratio and the highest trim angle of attack for a given center of gravity position. The capsule configuration with the asymmetrical heat shield, while offering the best overall liftto-drag ratio, exhibited a relatively large trim angle of attack changes as a function of roll attitude. The center of gravity longitudinal displacement required to achieve monopole stability of a single angle of attack trim point, increased with decreasing side wall angle.

An experimental investigation was made to determine the variation of shock shape and extent of subsonic region behind the shock wave with angle of attack for a 70° cone at various Mach numbers. The main interest was centered on those Mach numbers which produced detached shock waves or for which the possibility of detachment at angle of attack existed. The tests were conducted in the GALCIT 2.5 Supersonic Wind Tunnel at angles of attack of 0°, 3°, 6°, 9°, 12°, and 15° and Mach numbers of 1.438, 1.544, 1.584, 1.857, 1.986, and 3.01. It was found that, with increasing angle of attack and constant Mach number, the subsonic region behind the shock wave increased in the lower portion and decreased in the upper portion. With increasing Mach number the subsonic region decreased for all angles of attack. Interaction between the upper and lower portions of the shock wave affects the subsonic region behind the wave, suppressing its appearance on the upper surface, retarding its disappearance from the lower surface.