Journal of Aerospace Science and Technology
Year:2005 | Volume:2 | Issue:4|Papers Count:6

A generic single coaxial jet model was designed, constructed, and tested in a water tunnel specially designed for STOVL (Short Take Off and Vertical Landing Aircraft) flow applications. Laser Doppler Anemometry (LDA) surveys and Laser Induced Florescence (LIF) image capture were used to visualize the global flow patterns and identify the mean velocity and turbulence structure of an inverted-profile coaxial impinging jet, with and without a crossflow. Effects of the outer/inner jet velocity ratio and vertical jet/ horizontal crossflow velocity ratio on the development of the coaxial jet flow field were studied. The results presented are suitable for CFD validation purposes.      

Details of pressure distributions on a two dimensional airfoil oscillating in pitch through the stall in a 0.8 x 0.8 m 2 low-speed wind tunnel are presented. Pitching occurred around the airfoils quarter-chord axis. Pitch rate, Reynolds number, and oscillation amplitudes were varied to determine the effects on pressure and lift distributions. It was found that the mean angle of attack and pitching amplitude had strong effects on the flow field hence pressure distribution in the immediate vicinity of the airfoil leading edge, x/c 0.4. For pressure ports located at x/c> 0.4, the aforementioned effects were not strong. It seems that during the oscillatory motions the flow was mostly separated. This investigation shows weak unsteady effects when the maximum dynamic angle of attack was below that of the static stall; i. e. α max.dynamic ≤. α max.static. For higher angles of attack, strong unsteady effects appear which depend on the mean angle of attack, frequency and amplitude of the oscillation. Dynamic stall and dynamic reattachment contribute to afavorable effect of unsteadiness on the surface pressure signature and hence the mean lift coefficient which increases as compared to the steady state lift and pressure data.      

In this paper a closed-loop time-optimal control strategy for the non-linear lunar lander mission is developed. Generally, determination of closed-loop feedback control law is not usually feasible for many non-linear dynamic systems. In addition, there exist certain difficulties associated with the numerical determination of open-loop optimal control solution for non-linear systems, such as slow convergence rate and high sensitivity to initial guesstimates. Besides, if one manages to overcome these inherent difficulties, the determined optimal control strategy will be in an open-loop form, and thus, fully dependent on the initial condition. Obviously, in this way perturbations and noise processes will make the optimal trajectory deviate from its ideal predicted values in any actual operating environment. Our study focuses on the planar trajectory and control optimization of a lunar lander spacecraft as a viable example of nonlinear dynamic system. A fuzzy algorithm is augmented to our variational formulation of the problem in an attempt to create a closed-loop fuzzy guidance logic. The training process of the fuzzy system is greatly reduced through the introduction of a set of states related non-dimensional variables. Simulation results indicate that the developed methodology can be successfully utilized in other flight scenarios with good robustness to the actuator and measurement system noise.      

 In this paper, slewing maneuver of a flexible spacecraft with a large angle of rotation is considered and, assuming structural frequency uncertainties, a robust minimum-time optimal control law is developed. Considering typical bang-bang control commands with multiple symmetrical switches, a parameter optimization procedure is introduced to find the control forces/torques. The constrained minimization problem is augmented with the robustness constraints, which in turn increases the number of switches in the bang-bang control input to match the total number of the constraint equations. The steps of the solution algorithm to obtain the time optimal control input are discussed next. The developed control law is applied to a given satellite during a slewing maneuver. The simulation results show that the robust control input with just few switching times can significantly lessen the vibrating motion of the flexible appendage in the presence of structural frequency uncertainties, which reveals the merits of the developed control law      

In this paper, two-dimensional and axisymmetric, time dependent transonic and supersonic flows over a projectile overtaking a moving shock wave are considered. The flow is simulated numerically by solving full time averaged Navier-Stokes equations. The equations are linearized by Newton approach. The Roe's flux splitting method, second order central difference scheme for the diffusion terms and the second order approximation for time derivatives are used. For the turbulence terms, the Baldwin-Lomax and mixing length turbulence models are used. The present algorithm captures complicated features of flow including moving shock waves, expansion waves, boundary layers and wakes and their interactions. The results show that as the projectile passes through the moving shock wave, it changes the flow field features and pressure distribution dramatically. The drag force decreases and even becomes negative while the projectile takes over the shock wave. The flow features and the aerodynamic forces in transonic flow changes much more than those in the supersonic flow as the projectile passes through the shock wave. The results show that when the shock wave passes though the projectile, the flow field structures and the aerodynamic forces change abruptly. The drag force reduces and the shock wave passes through the projectile. Such variations are more vociferous for transonic flows than the supersonic flows. For transonic flows, the drag force changes sign and accelerates the projectile. Such behavior is important when the stability and control of the projectile are studied. Also such changes in pressure around the projectile and in the wake region change the projectile projectory.      

Linear stability analysis of the three dimensional plane wake flow is performed using a mapped finite difference scheme in a domain which is doubly infinite in the cross-stream direction of wake flow. The physical domain in cross-stream direction is mapped to the computational domain using a cotangent mapping of the form y = - βcot (πς). The Squire transformation [2}, proposed by Squire, is also used to relate the three-dimensional disturbances to the equivalent two-dimensional disturbances. The compact finite difference scheme of Lele [3] and the chain rule of differentiation are used to solve the Orr Sommerfeld equation. The result of linear stability analysis indicates that streamwise and the spanwise component of velocity eigenmodes is antisymmetric and the cross stream velocity eigenmode is symmetric. This is consistent with the DNS requirement of plane wake flow pertaining to solvability conditions [5].