Wing Tip Missiles Research Articles

Underwing Missile Aerodynamic Effects on Flight-Measured Limit-Cycle Oscillations

F-16 testing revealed differences in limit-cycle oscillation response characteristics associated with subtle aerodynamic variations of the underwing missiles. Physical length differences between long and short missiles led to the accidental discovery of the limit-cycle oscillation sensitivity. In addition to the length difference, the long missiles were found to have builtup collar sections, to which the fins and canards were attached to the missile body. Two cases are presented where both long and short missiles were carried underwing. For the case with no wingtip missiles present, carriage of the long missiles showed significant limit-cycle oscillation response, while the short missiles showed very little response. For the case with wingtip missiles present, the limit-cycle oscillation response levels were slightly higher for carriage of the short missiles but were otherwise quite similar to the long-missile responses. Computational fluid dynamics on a rigid wing suggest that the differences between the long and short missiles alter the flowfield by changing the impinging shock strength and location on the lower wing surface outboard of the underwing missile. For the empty wingtip launcher case, the location of the impinging shock on the lower surface influences the strength of the shock on the upper surface. This influence on the upper shock strength could be a key contributor to the limit-cycle oscillation mechanism when wingtip aerodynamics do not inhibit the flowfield.

Aeroelastic simulations using gridless boundary condition and small perturbation techniques

This paper examines the use of stationary Cartesian mesh for non-linear flutter computations involving complex geometries. The surface boundary conditions are implemented using reflected points which are determined via a gridless approach. The method uses a cloud of nodes in the vicinity of the surface to get a weighted-average of the flow properties using radial basis functions. To ensure computational efficiency and for local grid refinements, multigrid computations within an embedded grids framework are used. As the displacements of moving surfaces from their original position are typically small for flutter problems, a small perturbation boundary condition method is used to account for the moving surfaces. The method therefore does not require repeated grid re-generation for the deforming surfaces. The overall method is both accurate and robust. Computations of the well-known Onera M6 wing, RAE wing-body configuration, the AGARD 445.6 wing flutter test case show good accuracy and efficiencies. Simulations of the aeroelastic behavior of a complete fighter-type aircraft with wing tip missiles at high transonic speeds further demonstrate the practical usefulness of the present boundary conditions technique.

Assessment of Uncertain External Store Aerodynamics Using mu-p Flutter Analysis

F LUTTER testing is rarely performed on full-scale aircraft, due to the high risk for structural damage and failure. Instead, flutter boundaries are computed based on a numerical model of the aircraft, and flight testing under less critical conditions can be performed to collect data for validation of the numerical model. Clearly, problems arise when the flight-test data show deviation from the numerically predicted data. Deviations are likely to occur for any aircraft configuration, because model imperfections and simplifications always lead to some uncertainty in the numerical model, and the question arises as to the impact that these uncertainties have on the predicted flutter boundaries. Inmost cases, some parts of the numerical model, such as themass properties of different components, are well known, whereas other elements, such as the aerodynamic loads in some region of a wing with complex geometry, are known to be subject to uncertainty. Earlier studies [1–3] have shown how aerodynamic uncertainties can be introduced in the numerical model based on physical reasoning and known modeling difficulties. The same technique for modeling of aerodynamic uncertainty will be applied in this study, but it will also be shown how data points collected at subcritical conditions can be used to establish an uncertaintymodel that is capable of producing reliable flutter boundaries. In the present study, a wind-tunnel model was specifically designed to demonstrate and evaluate uncertainty modeling approaches. A fairly simple wing geometry and structure is chosen to minimize the errors introduced by modeling simplifications. An external store in the form of a wing-tip missile is used to increase the model complexity gradually, and the impact on the flutter behavior is investigated both numerically and experimentally. In cases in which the numerical predictions deviate from the experimental results, an uncertainty description is developed and -p flutter analysis and model validation [4] are applied to compute bounds on the flutter speed.

Multidisciplinary Design Optimization of a Built-Up Wing Structure with Tip Missile

The influence of a wing tip missile on the design optimization of a wing structure is studied. Finite element models of a realistic built-up wing structure are used to represent stiffness and mass properties. The store location and the effect of the store aerodynamics and mass are the variables included. A multidisciplinary optimization technique is used to compensate/restore the lost aeroelastic performance due to the presence of the store. Missile locations are the only configuration variables addressed besides the structural variables. The built-up wing box structure is optimized with constraints on the static strength and flutter speed. The thickness and the cross-sectional areas of the structural elements are the primary variables in the optimization. The aerodynamics of the tip missile has a significant effect on the flutter characteristics. In addition, the flutter behavior of the optimized structure is very sensitive to the tip missile movement along the tip chord. The results indicate that the effect of the tip missile aft movement must be examined in conjunction with the store aerodynamics. ASTROS is the primary tool used.

Robust Aeroservoelastic Design with Structural Variations and Modeling Uncertainties

An aeroservoelastic analysis procedure has been integrated with structural optimization and robust-control design tools. The multidisciplinary analysis and design process can address structural constraints, performance, and robust aeroservoelastic stability requirements simultaneously. The design process is expanded to account for structural variations and modeling uncertainties utilizing a reduced-order state-space model of the aeroelastic plant that is advantageous for efficient low-order controller design. Reduced-order modeling introduces unstructured modeling errors. Structural mass variations, such as those associated with a wide range of changeable external stores of a fighter aircraft, are also interpreted as model uncertainties. The models are expressed in a form compatible with standard robust control design procedures such as the μ synthesis and analysis technique incorporated in this study. The robust control system can then be included in an optimization process where more elaborate aeroelastic models are used for tuning selected structural variables and control gains for minimizing specific design objectives under stability, performance, and structural integrity requirements. To provide enhanced physical insight, the μ analysis procedure is supplemented by a technique for computing stability boundaries in multidimensional parametric uncertainty spaces that is based on the classical single-input/single-output gain-margin approach. The integrated design procedure is demonstrated with a realistic model of a fighter aircraft with four wing control surfaces and a wing-tip missile with variable inertial properties.

Aeroservoelastic Structural and Control Optimization Using Robust Design Schemes

The procedure for structural and robust control design of multi-input/multi-output aeroservoelastic systems presented contains modeling of uncertainties, synthesis of a robust controller, and a unified structural and control optimization process under stress, flutter, and control performance constraints. Robust control design techniques can enhance the preliminary structural design for cases where certain system parameters are not known exactly or are uncertain. An aeroservoelastic interaction module is used for mediation between the structural optimization code and the control synthesis tools in an iterative coupled process. It provides the reduced-size state-space aeroelastic models needed for control synthesis, including model sensitivities to structural uncertainties, and integrates the resulting control model in structural optimization that includes robust-control considerations in terms of singular value constraints. The efficient uncertainty modeling developed is based on the readily available structural sensitivity data and the linear fractional transformation tools. Modeling of inertial uncertainties of the controlled structure was performed to account for relatively large inertial deviations of a fighter aircraft wing-tip missile. The numerical example clearly demonstrates the effectiveness of the proposed design scheme and its usefulness at preliminary design stages of aircraft with multiple external store loads.

Viscous simulation of F-16 like configuration at high AoA

A viscous simulation of the flow over an F-16A model with inlet and the ventral fins, which were normally not considered in previous studies, is described. The method uses a patch grid multi-block, multi-grid Navier-Stokes solver. A novel method to grid the intake and diverter is used. The arrangement of the gridding system adopted minimises the number of grids needed particularly for viscous computations. A modified form of the Baldwin-Lomax algebraic turbulence model was used for the high AoA studies. To simplify the incorporation of turbulence model, the blocking arrangement is such that it allows a thin-layer Navier-Stokes equation formulation in the calculation of the eddy viscosity. A series of four flow conditions were studied; M = 0.9, AoA = 4°, 6.0°; and M = 1.2 at AoA = 6.0°; and M = 1.85 at AoA = 16°. Extensive comparisons with experimental data of the pressure distribution on the wing and on different streamwise stations show the modelling to be reasonably accurate. For the results presented here, the wing tip missile is ignored.