Tang Guojian

Nonlinear hierarchy-structured predictive control design for a generic hypersonic vehicle

A hierarchy-structured predictive controller is designed and analyzed for rotation motion dynamics of a generic hypersonic vehicle (GHV). This vehicle model has fast variability, is highly nonlinear, and includes uncertain parameters. The controller contains two subsystems, the inner-fast-loop nonlinear generable predictive controller (NGPC) and the outer-slow-loop NGPC, both of which are designed by the closed-form optimal generable predictive control method. Thus, the heavy on-line computational burden in the classical predictive control method is avoided. The hierarchy structure of the control system decreases the relative degree of each subsystem and helps increase the dynamic response speed of the attitude controller. In order to improve the robustness of the control system, a feedback correction algorithm is proposed that corrects the calculation error between the predictive model and the real dynamic model. Simulation studies are conducted for the trimmed cruise conditions of an altitude of 33.5 km and Mach 15 to investigate the responses of the vehicle to the step commands of angle of attack, sideslip angle, and bank angle. The simulation studies demonstrate that the proposed controller is robust with respect to the parametric uncertainties and atmospheric disturbance, and meets the performance requirements of GHV with acceptable control inputs.

A multi-impulse extended method for low-thrust trajectory optimization

The low-thrust trajectory optimization is replaced by a multi-impulse one. After the multi-impulse trajectory was obtained by solving a nonlinear programming problem, the multi-impulse is extended directly to the low-thrust by Q-guidance. Starting with the first segment of the multi-impulse trajectory, the thrust direction is set to consist with the impulse direction. The thrust magnitude is at its maximum. By propagating the initial state with the thrust for 1.0 day, the state of the spacecraft will update to a new one. A new impulse is obtained by solving the Lamberts problem between the new state and the final state. The process is then repeated until the impulse reduces to zero. Finally, the multi-impulse trajectory is converted to the low-thrust one by applying this process to all the other segments. Numerical results for two rendezvous missions fully validate the availability and efficiency of the Multi-Impulse Extended method.

A novel low-thrust trajectory optimization approach based on virtual gravitational body

A novel approach based on virtual gravitational body (VGB) is proposed for low-thrust trajectory optimization in this paper. To represent the low-thrust trajectory in an analytical form, a VGB, whose gravity on the spacecraft is equivalent to the resultant of the Sun gravity and the low-thrust, is introduced. Firstly, it is assumed that the VGB stays stationary in each trajectory segment to recover the low-thrust. Secondly, to satisfy the low-thrust magnitude constraint, the VGB is further assumed to follow a uniform rectilinear motion in each segment. It can be proved that, in both cases, the low-thrust trajectory is still a conic arc when expressed in a translational reference frame centered at the VGB. Finally, the original optimal control problem is converted into a nonlinear programming problem, which could be solved very fast owing to the analytical nature of conic arcs. Compared with a spherical shaping approach, numerical results of the Earth to asteroid 1989ML rendezvous mission demonstrate the availability and high efficiency of this VGB based approach.

Analysis of the autopilot design

The attitude autopilot and the velocity orientation autopilot of the flying vehicle swerve phase are designed and compared in this paper. The analysis result shows that the response speed of the velocity orientation autopilot is far faster than the attitude autopilot by reasonably designing, and the stability of the velocity orientation autopilot is better than the attitude autopilot.

Design and realization of terminal guidance law for passive homing interceptors

It is of the essence for passive homing interceptors to acquire information required to implement guidance laws. In practical applications, the accuracy of obtained information directly exerts an influence on terminal guidance performance. In order to achieve hit-to-kill, a modified three-dimensional proportional guidance law designed to enhance observability of passive guidance of interceptors is presented in this paper. The dynamic model of three-dimensional interception based on Singer model and observation equation of infrared seeker are established. A real-time two-step estimator is adopted to estimate the states used to realize the modified guidance law, solving the problem of practical application of guidance law. Simulation results indicate that the modified guidance law in cascade with the real-time two-step estimator produces small miss distance, satisfying the requirement of terminal guidance accuracy of interception.

Effect of atmosphere parameter oscillation at high altitude in the northern hemisphere for near space hypersonic flight aerothermodynamic prediction

For real atmosphere parameter fluctuates from the standard atmosphere model according the time and geography with stochastic atmospheric wave, it is important to predict the deviation effect on the aerothermodynamics environment for gliding hypersonic vehicle. However, according to the experts at NASA Johnson Space Center, at present neither the standard atmosphere model or the reference atmospheric model is able to describe the complex real atmosphere accurately. Especially, above 70 km altitude, the atmosphere parameter varies complicatedly, and deviates from standard atmosphere model intensely. The measurement results of atmosphere parameter from SABER (Sounding of the Atmosphere using Broadband Emission Radiometry) on TIMED satellite, launched at 2001, are testified widely with data from other sources. In this paper, based on measurement results between 2002 and 2010 from SABER on TIMED satellite, the atmosphere parameter statistics characters for 85 km altitude in typical months (January, April, July and October) and typical northern latitudes (0°, 20°, 40°, 60°, 80°) are obtained. In every research case, over 5000 stochastic samples are produced, which are satisfied with Max-Min distributing or normal distributing for the atmosphere model. The heating transfer on stagnation for stochastic sample in the on-the-spot survey parameter has been also studied. Then the statistics characteristics of the heating transfer on stagnation can be obtained and compared with the results based on U.S.1976 standard atmosphere, which shows the atmosphere oscillation statistics effect. The effect for latitudes and seasons on aerothermodynamics prediction has been investigated. It is shown that fluctuation for the atmosphere parameter plays an important role on the aerothermodynamics environment in near space flight. In the high altitude zone, season factor has a notable effect on the heating distribution. However, close to the equator, season factor has a little effect. At the same reason, heating transfer in higher altitude zone is usually greater than that from the standard model; on the contrary, heating transfer in lower altitude zone is general lower than that from the standard model. In summer (July), the higher latitude zone locates, the higher heating distribution it gets. The maximum heating transfer appears in July on 80° latitude. In extreme case, the maximum stagnation heating exceeds the result from the standard atmosphere model about 40%. At the same altitude, the maximum heating transfer appears in April or July. In the future work, the affection about the difference between Northern and Southern Hemisphere, between ocean and land and atmosphere parameter relativity will be researched.

Research on strategy to decrease wind load of solid rocket with multiple-stage boosters

For solid rocket Hying in the endoatmosphere, the disturbance of wind field movement on ballistic trajectory cannot be neglected. This paper presents the dynamical model of solid rocket, selects a typical wind field model and analyzes the effects on ballistic characteristics in boost phase caused by wind field disturbance. And then a coefficient is introduced to reflect the degree of wind field disturbance and used as a basic reference to research the transmission mechanism of the effects on angle of attack and aerodynamic loads of the wind field disturbance. At last, this paper proposes a strategy to decrease wind load and the simulations show the efficiency of the proposed strategy.

Libration point orbit rendezvous in real earth-moon system using terminal sliding mode control

This paper studies the problem of libration point orbit rendezvous using terminal sliding mode control. The underlying dynamics are the full nonlinear equations of motion in a complete Solar System model, taking into account all the gravitational forces from the Sun, the Moon, and the planets, whose positions are computed using the JPL DE405 ephemeris. The prominent advantage of terminal sliding mode control, owing to the specific design of the sliding surface, is that the time of convergence can be prescribed a priori, which means the rendezvous time is fully under control. Besides the sliding surface design, a controller suited to the dynamical model is also derived to insure the global stabilization of the approach. Preliminary results validate the availability of the approach, and comparisons are made with the results obtained using glideslope guidance in previous work.

Accuracy Analysis of Constellation Orientation Parameters Determination via X-Ray Pulsar Measurement

Pulsars are rapidly rotating, highly magnetized neutron stars that produce stable and predictable signatures. It can provide an external spatial reference to determine constellation orientation parameters. Based on the pulsar relative observation model and inter-satellite range model, Posterior Cramer-Rao Lower Bound (PCRLB) method is utilized to acquire orientation parameters accuracy in theory. Three factors, which are baseline, delta time of arrival (DTOA) estimation accuracy and directional angle, affecting the directional angle accuracy are analyzed in theory and simulation.

Improved direct-shooting method for low-thrust trajectories optimization

For low-thrust trajectories optimization, an improved direct-shooting method is proposed to overcome the two major drawbacks of the original one: the low-thrust trajectories are computed inefficiently using explicit numerical integration; and the control steering profile is more likely to converge at local optimum due to its local linear interpolation characteristics. The measures taken in this work include deducing an analytical solution of the low-thrust trajectory and representing the low-thrust control as a global fourth-order polynomial. The proposed method has been tested on two rendezvous missions: one from the Earth to Mars, the other to near-Earth asteroid 2001GP2.