A new automated multi-level of fidelity Multi-Disciplinary Design Optimization (MDO) methodology has been developed at the MDO Laboratory of K.N. Toosi University of Technology. This paper explains a new design approach by formulation of developed disciplinary modules. A conceptual design for a small, solid-propellant launch vehicle was considered at two levels of fidelity structure. Low and medium level of fidelity disciplinary codes were developed and linked. Appropriate design and analysis codes were defined according to their effect on the conceptual design process. Simultaneous optimization of the launch vehicle was performed at the discipline level and system level. Propulsion, aerodynamics, structure and trajectory disciplinary codes were used. To reach the minimum launch weight, the Low LoF code first searches the whole design space to achieve the mission requirements. Then the medium LoF code receives the output of the low LoF and gives a value near the optimum launch weight with more details and higher fidelity.
To propel a spacecraft away from the Sun, a magneto plasma sail (MPS) spacecraft produces an artificial magnetic cavity to block the hypersonic solar wind. To make a large magnetic cavity sufficient to obtain significant thrust, the MPS spacecraft increases the magnetic cavity size using an onboard coil with assistance from a plasma jet. This process is called magnetic field inflation. In this study, we performed ideal and resistive magnetohydrodynamic (MHD) analyses to investigate the magnetic diffusion effect on the magnetic field inflation process. Our results indicate that a dipole-like magnetic field is drastically deformed by a plasma jet; when the magnetic Reynolds number Rm was 10 or more, the magnetic field lines were nearly identical to the streamlines of the plasma jet. Hence, no magnetic diffusion effect appeared for Rm>10. Meanwhile, when Rm is an order of unity, the magnetic diffusion effect was remarkable in the current sheet formed around equatorial region. For example, when the divergence angle of a plasma jet in the polar direction was 30°, the magnetic field strength at 40 m from the spacecraft (calculated by resistive MHD model) was 19% smaller than the ideal MHD model (Rm=∞).
This paper focuses on flow visualization, normal force and pitch moment testing of the NASA TP-1803 strake-wing model at high attack angles. The dynamic aerocharacteristics for pitching with various reduced frequencies and two sideslip angles β=0° and 10° in the water tunnel are compared with those for static case. For α=20°–50°, the strake/wing vortices breakdown positions occur later for β=10° than for β=0°. The value of the normal force coefficient under sideslip angle β=10° is greater than β=0° at high attack angles. In the pitch-down process, the aerodynamic center creates a nose-up pitching moment and the model becomes unstable compared to the static condition. As the pitch reduced frequency increases, the wing vortices sustain longer flow lines and provide more normal force during pitch-up motion. In addition, the hysteresis loop of the normal force curve is larger for higher reduced frequencies.
The Fuzzy logic controller (FLC) is well-known for robustness to parameter variations and ability to reject noise. However, its design requires definition of many parameters. This work proposes a systematic and simple procedure to develop an integrated fuzzy-based guidance law which consists of three FLC. Each is activated in a region of the interception. Another fuzzy-based switching system is introduced to allow smooth transition between these controllers. The parameters of all the fuzzy controllers, which include the distribution of the membership functions and the rules, are obtained simply by observing the function of each controller. Furthermore, these parameters are tuned by genetic algorithms by solving an optimization problem to minimize the interception time, missile acceleration commands, and miss distance. The simulation results show that the proposed procedure can generate a guidance law with satisfactory performance.
This paper presents a numerical method for deriving a symplectic state transition matrix for high-fidelity Earth orbits subject to non-dissipative perturbation forces. By taking advantage of properties of Hamiltonian systems, this method provides an exact solution space mapping of linearized orbital dynamics, preserving the symplectic structure that all Hamiltonian systems should possess by nature. This method can be applied to accurate, yet computationally efficient dynamic filters, long-term propagations of the motions of formation flying spacecraft and the eigenstructure analysis of N-body dynamics, etc., when the exact structure-preserving property is crucial. We show the derivation of the numerical method of symplectic state transition matrix, and apply it to Earth orbits with perturbation forces based on real ephemerides. These numerical examples reveal that this method shows improvements in preserving the structural properties of the state transition matrix, and in the computational efficiency compared to the conventional linear state transition matrix with Euler or Runge-Kutta integration.
This paper presents experimental results of unsteady aerodynamic interactions including Shock/Shock Interaction (SSI) and Shock/Boundary Layer Interaction (SBLI) between two bodies at hypersonic speed. These interactions can be seen in space vehicles consisting of multi-bodies, such as a TSTO, or at a scramjet engine inlet. The present study considers the effect of a flat plate below the SSI where a boundary-layer is developed on the plate surface. More specifically, the interacted flow for a combination of a flat plate (FP) and a hemi-circular cylinder (HCC) is examined at a hypersonic speed (M∞=8.1); the distributions of surface pressure and heat transfer rate are measured. To obtain various SSI patterns, the clearance between two bodies (FP and HCC) is changed. Results show that unsteadiness at the SSI point causes a feedback loop between the two bodies; a jet flow impinges on the FP, the effect of which propagates upstream where the jet impinges on the FP, and the aerodynamic and aerothermodynamic loads reach their maxima. Finally, we found that the feedback loop can be destroyed by installing a fence on the FP to reduce unsteadiness of flow field.
Three-dimensional flow simulations through a turbine cascade were carried out for an exit isentropic Mach number of 0.79 and a Reynolds number of 2.8×106. The main objective is to increase the base pressure. Calculations were carried out using an in-house numerical code, where a 2nd order Roe’s flux-difference splitting for inviscid numerical fluxes, a 2nd order implicit dual time method for time integration, and Delayed Detached Eddy Simulation for turbulence were employed. The present idea to increase the base pressure is to connect the pressure and suction sides of the blade using microtubes distributed uniformly along the span on the trailing edge. Simulated results show that this modification does not affect the blade load, increases the base pressure by 0.7%, and decreases the overall loss by 3%.
The paper presents a new approach to analysis of dynamic properties of the rotating airplane. The analysis is based on the rate of rotational energy and a full non-linear model of airplane dynamics. It can be applied to a whole range of attack angles. To provide stable rotation, it is sufficient to use algebra and know only the inertial properties of the airplane and the functional relations of either aerodynamic moments or angular accelerations. We prove that the results are fully congruent with those obtained by conventional methods within the scope of validity of the simplified model.
The paper describes how to simulate the flight of a flapping-wing micro-aerial vehicle (MAV). It uses an aerodynamic database generated using three-dimensional Navier-Stokes code. The database is composed of the time mean aerodynamic forces and moments generated at various flapping wing motions in various flight modes. Flight is simulated utilizing the database by interpolation. The procedure is applied to transition flight of a dragonfly-type MAV with two-pairs of resonance-type flapping wings. The present MAV attains the mission of hovering, transition and cruising flights successfully with stable attitude.
In this paper, variation in cushion pressure is investigated theoretically to achieve SES ride comfort. A simple and practical control method is proposed to control cushion pressure changes caused by pumping that appears in SES cruising over waves. Cushion pressure is controlled by varying the discharge height of the vertical nozzle. In this method, the nozzle height (hover height) is kept constant according to craft motion. As the result of investigations, cushion pressures are controlled successfully by the proposed method. In addition, we show that the pressure variation depends both on hover heights and on the rates of change in cushion volume. Therefore, it is necessary to consider phase differences between craft motion and vertical nozzle displacements.