Midcourse Maneuver Research Articles

Research on Divert and Attitude Control System Technology of Ballistic Missile Midcourse Maneuver Penetration Warhead

With the continuous development of American midcourse interception technology, it is particularly important to promote the midcourse penetration capability construction of ballistic missiles. The use of maneuver penetration technology in the midcourse of ballistic missiles can effectively deal with the midcourse interception of foreign missiles. In order to promote the development of midcourse maneuver penetration technology, this paper takes the Divert and Attitude Control System (DACS) of maneuver penetration warhead as the starting point, refers to the development of DACS of Kinetic Kill Vehicle in the United States, and summarizes the research status of gas valve, the key structure of DACS. On this basis, combined with the structural characteristics of maneuver penetration warhead, the key technologies related to the construction of DACS of penetration warhead are summarized, which lays a foundation for the development of DACS of penetration warhead in the future.

Realizing Midcourse Penetration With Deep Reinforcement Learning

A midcourse maneuver controller is obtained using deep reinforcement learning to maintain the survivability of a ballistic missile. First, the midcourse is abstracted as a Markov decision process (MDP) with an unknown system state equation. Then, a controller formed by the Dueling Double Deep Q (D3Q) neural network is used to approximate the state-action value function of the MDP. In order to make the controller’s intelligence improved by deep reinforcement learning, the state space, action space, and instant reward function of the MDP are customized. The controller uses a real-time situation as input and outputs the ignition states of pulse motors. Offline training shows that deep reinforcement learning can achieve the optimal strategy’s convergence after approximately 65 hours. Online tests demonstrate the controller’s ability to avoid an interceptor intelligently and to account for an entry error. In scenarios with multiple random factors, the controller achieved a penetration probability of 100% and a mean re-entry error of less than 5000 m.

Response to 2019 Barringer Medal award

Response to 2019 Barringer Medal award

Real Time Mid-course Maneuver and Guidance of a Generic Reentry Vehicle

The aim of any mission is to accomplish the final objective with desired accuracy and the same is valid for a generic launch vehicle. In many missions it is necessary to execute mid-course maneuvers with an intentional diversion trajectory to create a counter measure or to avoid certain specific known geographical locations. The current work elaborates a novel and practically implementable mid-course maneuver and an ascent phase guidance of a reentry vehicle executing an in-flight determined mid-course maneuver (trajectory reshaping) without compromising the accuracy of the final achieved target position. The robustness of the algorithm is validated with 6DoF simulation results by considering the dispersion of the burnout state vector conditions which arises due to variations in thrust profile, aerodynamics characteristics of the vehicle, atmosphere, etc. Defence Science Journal, 2013, 63(4), pp.346 -354 , DOI:http://dx.doi.org/10.14429/dsj.63.4207

Targeting low-energy transfers to low lunar orbit

A targeting scheme is presented to build 21-day launch periods for low-energy trajectories from any specified Earth parking orbit to any specified low lunar orbit, using up to two mid-course maneuvers. A total of 288 launch periods are constructed for transfers to a variety of different targeted low lunar orbits, arriving at different times during the month. Each orbital insertion occurs at an altitude of 100km above the lunar surface and each orbit sampled in this paper is polar; the longitude of ascending node and the argument of periapse vary for each sampled orbit. An analysis is presented to characterize the total transfer ΔV required to build a launch period, as well as the total transfer ΔV required to transfer from any given Earth parking orbit inclination to each sampled lunar orbit.

Multiobjective genetic optimization of Earth–Moon trajectories in the restricted four-body problem

In this paper, optimal trajectories of a spacecraft traveling from Earth to Moon using impulsive maneuvers (ΔV maneuvers) are investigated. The total flight time and the summation of impulsive maneuvers ΔV are the objective functions to be minimized. The main celestial bodies influencing the motion of the spacecraft in this journey are Sun, Earth and Moon. Therefore, a three-dimensional restricted four-body problem (R4BP) model is utilized to represent the motion of the spacecraft in the gravitational field of these celestial bodies. The total ΔV of the maneuvers is minimized by eliminating the ΔV required for capturing the spacecraft by Moon. In this regard, only a mid-course impulsive maneuver is utilized for Moon ballistic capture. To achieve such trajectories, the optimization problem is parameterized with respect to the orbital elements of the ballistic capture orbits around Moon, the arrival date and a mid-course maneuver time. The equations of motion are solved backward in time with three impulsive maneuvers up to a specified low Earth parking orbit. The results show high potential and capability of this type of parameterization in finding several Pareto-optimal trajectories. Using the non-dominated sorting genetic algorithm with crowding distance sorting (NSGA-II) for the resulting multiobjective optimization problem, several trajectories are discovered. The resulting trajectories of the presented scheme permit alternative trade-off studies by designers incorporating higher level information and mission priorities.

Optimal power-limited rendezvous for linearized equations of motion

The form of solution of the optimal power-limited rendezvous problem for linear equations of motion is made to conform to previous developments by the author. The solution is then applied to the problem of rendezvous of a spacecraft with an object in the vicinity of a nominal Keplerian orbit. For the case in which the nominal orbit is circular, the controllability matrix can be inverted symbolically, resulting in a closed-form solution for the thrusting function. This result can be expressed as a closed-loop feedback law. For noncircular nominal orbits, a feedback law can be approximated by repeated numerical matrix inversions and reinitialization of the problem. PTIMAL power-limited terminal rendezvous of a spacecraft based on linearized equations of motion offers the mathematical advantage of closed-form or nearly closed-form solutions of the linearized trajectory optimization problem. In many cases this leads to thrust control in the form of optimal feedback laws, or at least, control functions that can be easily reinitialized on a regular basis. This gain in adaptivity may, to some extent, compensate for the loss in accuracy due to the linearization. There are many potential applications of a linearized analysis of optimal power-limited rendezvous. These typically involve the maneuvering of a low-thrust spacecraft near a real or fictitious object in a nominal orbit. Examples include maneuvers in the vicinity of artificial satellites, space stations, small comets, or asteroids. Examples where the object is fictitious include low-thrust satellite station keeping or interplanetar y missions in which the usual high-thrust midcourse maneuver is replaced by a continuous long-term lowthrust correction toward the nominal orbital position and velocity. Early work on the computation of power-limited , variable exhaust-velocity spacecraft trajectories was performed by Irving,1 Ross and Leitman,2 Edelbaum,3 Saltzer and Fetheroff,4 Melbourne,5 and Melbourne and Sauer.6'7 Immediately after these initial studies, investigators began to look at various linear models.812 Billik8 and Gobetz9 used linearization about nominal circular orbits, Edelbaum10 used linearization with respect to the orbital parameters of an ellipse, Gobetz11 linearized about an ellipse of low eccentricity, and Euler12 used the Tschauner-Hempel13'14 linearization for general elliptical orbits. More general discussions that include linear and nonlinear models for both power-limited and constant-exhaust-velocity problems can be found in the works of Edelbaum15 and Marec.16 Recently, optimal power-limited low-thrust trajectories to return to an original station in orbit after an impulsive maneuver away from this orbit were examined by Lembeck and Trussing17 using equations linearized about a circular orbit. The same linearized equations were used by Coverstone-Carroll and Trussing18 to investigate the cooperative power-limited rendezvous of two spacecraft near a nominal circular orbit. Power-limited trajectories in an inverse-square gravity field have also been considered by Prussing.19 In the present paper, we consider the fixed-time linear, powerlimited rendezvous problem. We utilize a somewhat more general cost function than usual, an integral of the product of a weighting

Navigation and Guidance of Japanese Deepspace Probes Encountering Halley's Comet

The first Japanese deepspace probes called Sakigake and Suisei were launched from Uchinoura Launching Site, located at 1,500 km south-west of Tokyo, respectively on January 8 and August 19, 1985. The Institute of Space and Astronautical Science has developed the new four-stage solid propellant boosters called M3S-II for this purpose and constructed a gigantic antenna - 64 m in diameter - in Usuda, located at 150 km north-west of Tokyo. These probes successfully encounted Halley’s comet in March, 1986. In this paper, the guidance and navigation problems, namely the guidance of the last stage rocket in the direct ascent scheme, the midcourse maneuvers and the tracking and navigation of the probes throughout the mission, will be discussed.

Search for the Viking 2 Landing Site

The search for the landing site of Viking 2 was more extensive than the search for the Viking 1 site. Seven times as much area (4.5 million square kilometers) was examined as for Viking 1. Cydonia (B1) and Capri (C1) sites were examined with the Viking 1 orbiter. The B latitude band (40 degrees to 50 degrees N) was selected before the final midcourse maneuver of Viking 2 because of its high scientific interest (that is, high atmospheric water content, surface temperature, possible near-surface permafrost, and a different geological domain). The Viking 1 orbiter continued photographing the Cydonia (B1) site to search for an area large and smooth enough on which to land (three-sigma ellipse; 100 by 260 kilometers); such an area was not found. The second spacecraft photographed and made infrared measurements in large areas in Arcadia (B2) and Utopia Planitia (B3). Both areas are highly textured, mottled cratered plains with abundant impact craters like Cydonia (B1), but smaller sectors in each area are partially mantled by wind-formed deposits. The thermal inertia, from which the grain size of surface material can be computed, and atmospheric water content were determined from the infrared observations. A region in Utopia Planitia, west of the crater Mie, was selected: the landing took place successfully on 3 September 1976 at 3:58:20 p.m. Pacific Daylight Time, earth received time.

Off-optimum timing of interplanetary midcourse guidance maneuvers

The nature of the optimum solution with respect to the timing of interplanetary midcourse guidance maneuvers is investigated. The analysis provides a realistic cost function, defined as the sum of the root-mean-square mid-course maneuvers, for both the fixed-time-of-arrival and the variable-time-of-arrival guidance logics and uses the time of the last maneuver as a functional constraint. The results of the analysis indicate that only the timing of the first and of the last maneuvers is sensitive to perturbation about the optimal values and that a variation in the timing of the intermediate maneuvers produces small changes in the total midcourse velocity requirement.

Mariner mars 1969 navigation, guidance and control

Design, mechanization, and flight test results of the Mariner Mars 1969 navigation, guidance and control systems, are discussed. A trajectory design section describes near-earth launch trajectory, planetary targeting and constraints, as well as the tradeoffs made on the trajectory selection. Among the factors considered are reliability, direct ascent vs. parking orbit and higher spacecraft weight vs. extended launch window. Guidance and orbit determination factors are discussed, including earth-based radio and spacecraft optical navigation and the accuracy of orbit determination and maneuver execution. The control systems section identifies differences in the attitude control, midcourse maneuver, and science instrument scan pointing systems from those used in previous Mariner spacecraft. A description is given of the new Central Computer and Sequencer (CC&S) system, used for the first time on this mission, which allows extremely flexible spacecraft operation using in-flight reprogramming of the computer memory by radio command. Finally, reliability summaries and performance evaluations permit conclusions as to the effectiveness of the Mariner Mars 1969 navigation, guidance, and control systems to conduct near-term and more advanced planetary missions.

Mariner Mars 1969 Navigation, Guidance, and Control

Design, mechanization, and flight test results of the Mariner Mars 1969 navigation, guidance, and control systems are summarized. A trajectory design section describes near-earth launch trajectory, planetary targeting, and constraints, as well as the tradeoffs made on the trajectory selection. Among the factors considered are reliability, direct ascent vs parking orbit, and higher spacecraft weight vs extended launch window. Guidance and orbit determination factors are discussed, including earth-based radio and spacecraft optical navigation and the accuracy of orbit determination and maneuver execution.The control systems section identifies differences in the attitude control, midcourse maneuver, and science instrument scan pointing systems from those used in previous Mariner spacecraft. A description is given of the new Central Computer and Sequencer (CC&S) system, used for the first time on this mission, which allows extremely flexible spacecraft operation using in-flight reprogramming of the computer memory by radio command. Finally, reliability summaries and performance evaluations permit conclusions as to the effectiveness of the Mariner Mars 1969 navigation, guidance, and control systems to conduct near-term and more advanced planetary missions.

The Voyage of Mariner II

A detailed account is presented of the events leading to the successful rendezvous of Mariner II with Venus on Dec. 14, 1962. The development of the spacecraft is described. The launching and some of the adventures in space, such as the mid-course maneuver, are described. Finally, the rendezvous and some of the measurements are discussed. (D.L.C.)