Pitch Damping Derivatives Research Articles

Analysis of uncertainty factors in the prediction of missile pitch damping derivatives by unstructured CFD solver

Based on the transient planar pitching method, the dynamic derivative calculation function is implemented on an unstructured computational fluid (CFD) software NNW-FlowStar. Then the pitch-damping dynamic derivatives of the standard model of the ANF missile at zero angles of attack and different Mach numbers (0.5∼4.5) are calculated. The influence of various numerical and physical parameters on the calculation results of dynamic derivatives is studied in detail. It is found that the dynamic derivatives of pitch damping are independent of the values of amplitude (0.125° ≤ A ≤ 1°) and reduction frequency (0.025 ≤ k ≤ 0.2). When the number of calculation steps of the sine oscillation period is 200 and the number of inner iterations is 40, the prediction accuracy and computing efficiency can meet the engineering requirements simultaneously in this case. The influence of different grid resolutions on the dynamic derivative calculation results is studied. Results show that the grid requirements of dynamic derivative calculation are consistent with those of steady-state calculations. Finally, the calculation results are compared with the free flight test results, which verifies the reliability of the methods and parameters suggested in this paper.

Analysis of Damping Derivatives for Delta Wings in Hypersonic Flow for Curved Leading Edges with Full Sine Wave

In this study, an attempt is made to evaluate the effect of first arched ends on the damping derived due to the pitch rate aimed at the variable sine wave bounty, flow deflection angle δ, pivot position, and the Mach numbers. Results show that with the escalation in the bounty of the complete sine wave (i.e., positive amplitude) there is an enlightened escalation in the pitch damping derivatives from h = 0, later in the downstream in the route of the sprawling verge it decreases till the location of the center of pressure and vice versa. At the location where the reasonable force acts, when we consider the stability derivatives in damping for the rate of pitch q, there is a rise in the numerical tenets of the spinoffs. This increase is non-linear in nature and not like for position near the leading edges. The level of the stifling derivatives owing to variations in Mach numbers, flow bend approach δ, and generosity of the sine wave remained in the same range.

Research on Damping Derivatives for Delta Wings in Hypersonic Flow for Curved Leading Edges with Full Sine Wave

In this study, an attempt is made to evaluate the effect of first arched ends on the damping derived due to the pitch rate aimed at the variable sine wave bounty, flow deflection angle δ, pivot position, and the Mach numbers. Results show that with the escalation in the bounty of the complete sine wave (i.e., positive amplitude) there is an enlightened escalation in the pitch damping derivatives from h = 0, later in the downstream in the route of the sprawling verge it decreases till the location of the center of pressure and vice versa. At the location where the reasonable force acts, when we consider the stability derivatives in damping for the rate of pitch q, there is a rise in the numerical tenets of the spinoffs. This increase is non-linear in nature and not like for position near the leading edges. The level of the stifling derivatives owing to variations in Mach numbers, flow bend approach δ, and generosity of the sine wave remained in the same range.

Investigation of pitch damping derivatives for the Standard Dynamic Model at high angles of attack using neural network

Numerical investigations of flow field around the Standard Dynamic Model (SDM) are performed using computational fluid dynamics approach. Initially, the static SDM is studied at various angles of attack up to 70∘ and an agreement between normal force and pitching moment predictions with experimental data is ensured, thanks to the polyhedral grids. Subsequently, the response of the SDM under single frequency sinusoidal pitching motions is computed and the associated pitching moment coefficient damping is obtained using two methods of classical Fourier coefficients and multilayer perceptron (MLP) artificial neural network. The results are compared to published experimental values. In the final stage, frequency sweep sinusoidal excitation in pitch axis is conducted with 30∘ amplitude in transonic flow and the MLP is exploited to calculate variable stability derivatives. It is observed that damping derivatives are highly dependent on both amplitude and frequency of oscillation. Also, an increase in the frequency of motion lowers the pitching moment damping. As the motion frequency rises, the pitching moment amplitude increment is seen to be greater than that of normal force. Polyhedral mesh as well as overset grid technique are adopted in flow field computations, leading to high fidelity of numerical simulations.

Research on Numerical Simulation of Support Interference for Dynamic Derivatives at Hypersonic Flight Speeds

The research on dynamic derivative under the dynamic unsteady condition is one of the most difficult aspects of the aircraft development process. For the special experiment of dynamic derivatives, no numerical simulation of support interference has yet been systematically studied. The numerical simulation of support interference for identification of dynamic derivatives with 7o blunted cone under the hypersonic condition was done in this paper. The 2-order Roe scheme and the dual-time-step method based on LU-SGS were respectively applied to discrete of the spatial and time derivative of the unsteady flow. A least square method based on the Etkin unsteady aerodynamic model was used for identification of dynamic derivatives. Hypersonic missile module was adopted as a verification example, and the numerical calculation results are consistent with the experimental results. For two different sting support forms with 7o blunted cone, the impact of sting support interference on dynamic derivatives was studied. Results show the moment coefficients of two support forms under static conditions are essentially the same, while there is a big difference in the pitch damping derivatives under dynamic conditions. The support interference of nonlinear aerodynamic loads resulted from the shock wave induced separation and other unsteady flow structures under dynamic conditions is far more complicated than that under static conditions, and the correction law of support interference under static conditions cannot be applied directly to the unsteady dynamic situations.

Effect of Angle of Attack on Stability Derivatives of a Delta Wing with Straight Leading Edge in Supersonic Flow

In the Present paper effect of angle of incidence on pitching derivatives of a delta wing with Straight leading edges for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude to give a piston theory. From the results it is seen that with the increase in the Mach number, there is a decrement of stiffness as well as the damping derivatives in pitch for all the Mach number tested, however, the magnitude of decrement for different inertia level will differ. It is seen that with the increase in the angle of attack both stiffness and damping derivatives increase linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees. When these stability derivatives were considered at various pivot positions, namely h = 0.0, 0.2, 0.4, 0.6, 0.8, and 1.0. After scanning the results it was observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of both the stability derivatives were decreasing progressively with the pivot position. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, and secondary wave reflections are neglected.

Predictions of Dynamic Damping Coefficients of Basic Finner Based on CFD

A method to predict pitch and roll damping derivatives of aircraft geometries with fins using an unsteady RANS solver is presented. A three-dimensional structured RANS solver based on the arbitrary Lagrangian-Eulerian (ALE) formulation with a dynamically deforming mesh algorithm is used and validated with the wind tunnel and ballistic range data available in the literature. Roll and pitch damping derivatives are calculated from load history of the unsteady flow around the model. A standard research configuration, known as the Basic Finner, is studied under forced pitching and rolling conditions. Pitching and rolling motions with oscillation are analyzed at supersonic Mach numbers ranging from 1.5 to 2.5. Predicted results showed good agreement with the available wind tunnel data.

Two complementary approaches to estimate downwash lag effects from flight data

ALIDATION of the wind-tunnel and analytical estimates of the aerodynamic derivatives with estimates from flight test data is an important application of the system identification methodology. Reliable and accurate estimates of a large number of aerodynamic derivatives are obtained from flight data using the maximum likelihood method, although, routinely, explicit accounting for certain unsteady aerodynamic effects such as downwash and sidewash interferences pose difficulties. Recent advances in both parameter estimation methods and in flight test techniques have provided a new impetus for modeling and identification of such unsteady interference effects in aircraft dynamics.1'3 This Note addresses in some depth the problem of accounting for downwash lag effects in aircraft parameter estimation from flight data. One possibility of including such an effect is to proceed from the fundamental and to model the forces acting on the wing and tail plane separately. In such a case it becomes possible to estimate explicitly the downwash angle and also to include the lag effect. It leads, however, to a model that is nonlinear in parameters and necessarily requires an estimation program capable of handling nonlinear model postulates. The other approach is a simplification based on linearization. It leads to the first-order approximation of aerodynamic derivatives with respect to translational acceleration. The derivatives due to the body-fixed translational acceleration in the vertical direction are equivalent in the stability axis to those with respect to the rate of change of angle of attack. The second approach was first investigated in Ref. 1 to estimate from flight data the pitching moment derivatives with respect to pitch rate and rate of change of angle of attack of a highly maneuver able aircraft with large roll angle capabilities. A parameter estimation program capable of handling linear as well as nonlinear system models is used here to compare the two approaches of accounting for the downwash. Further, an attempt is made to investigate the possibility of using the linearized approach to estimate separately the two pitch damping derivatives from flight tests with a larger aircraft having limited roll angle capabilities.

Experimental Missile Pitch- and Moll-Damping Characteristics at Large

Pitch- and roll-damping test mechanisms were developed for obtaining stability characteristics of high fineness ratio missiles at angles of attack up to 90 deg. Wind-tunnel tests were conducted on two missile configurations at Mach numbers 0.22 through 2.5. Mechanism evaluation tests showed that strut interference was not appreciable. The pitch-damping derivatives were found to be highly nonlinear functions of angle of attack, and the roll-damping and Magnus characteristics were very nonlinear with spin rate'at high angles of attack and low Mach numbers. Large and erratic variations in side force and yawing moment with angle of attack and high autorotation characteristics also were found at the low Mach numbers.

Viking Pitch Damping Derivatives as Influenced by Support Interference and Test Techniques

Sting support interference effects on the pitch damping derivatives and a comparison of damping characteristics using forced and free oscillation test methods were investigated on 140° spherically blunted cone configurations at M = 1.76-3.0. Sting influences are minimal with sting/model diameter ratios =0.183 and effective sting lengths =3.55 dm. Increasing ds/dm to 0.53 provides a considerable influence to the damping parameter at a = 0°-3° at all Reynolds numbers but shows little or no effect at higher alphas. A good correlation of freeflight and small diameter sting captive model wake geometries (wake neck position and size) was obtained which adds confidence in obtaining damping derivatives through captive model testing. The damping characteristics obtained on 11.23% and 4.35% scale Viking configurations using forced and free oscillation test methods respectively in two different test facilities are compared. The agreement is quite good at almost all test conditions.