Nuclear Electric Propulsion Spacecraft Research Articles

Optimum design of nuclear electric propulsion spacecraft for deep space exploration

Future space exploration is now focused on new field-deep space planets. Deep space exploration and the development and utilization of resources are inestimable strategic significance for the country to seize the initiative and command heights of deep space exploration. Nuclear electric propulsion (NEP) systems convert heat from the fission reactor to electrical power and then use the electrical power to produce thrust. Compared with traditional propulsion technology, the NEP system with variable Isp can be used for several applications. The NEP spacecraft is more suitable for deep space exploration missions due to the advantages of high specific impulse, high power, and long life. This paper analyzes the relationship between the transfer travel time, specific mass, power, and payload ratio of the NEP spacecraft through simple performance models. Use the mass optimization and specific mass optimization models based on the NEP system composition and small thrust orbit theory to maximize the payload ratio for a given transfer time and technological characteristics. Finally, applied this NEP model to discuss the feasibility of Mars, Jupiter, and Saturn transfer missions and compared it with the Tianwen-1, Juno, and Cassini–Huygens spacecraft, respectively. Results show that when the NEP spacecraft specific mass reaches 4.77 kg/kW, the Earth–Mars transfer time can be changed to 331.31 days, the payload increases to 1650 kg, the transfer time for Jupiter and Saturn mission would be shorted to 661.31 days and 1131.31 days, the corresponding payload significantly increases which achieved to 1270 kg and 2981 kg. The nuclear electric propulsion spacecraft dramatically improves the detection capability of the spacecraft and provides a reference for the feasibility demonstration and subsequent design of the deep space and the extrasolar planet exploration.

Low-thrust trajectory design as a constrained global optimization problem

The design of a spacecraft trajectory can be formulated as a global optimization task. The complexity of the resulting problem depends greatly on the final target planet, the chosen spacecraft intermediate route, and the type of engine and power system available on-board. Few attempts have been made to directly use a global optimization framework to design trajectories that make use of low-thrust propulsion because of the large scale and extreme complexity of the resulting non-linear programming problem. The presence of non-convex constraints, in particular, requires the use of solvers able to deal with such an added complexity. Here, the Sims–Flanagan transcription method is proposed to model the low-thrust trajectory design as a constrained global optimization problem. Then, two different solvers are applied: basin hopping and simulated annealing with adaptive neighbourhood. Both algorithms are hybridized with a local search. Two different interplanetary trajectories are considered: an Earth–Earth–Jupiter transfer with a nuclear electric propulsion spacecraft inspired by the Jupiter Icy Moons Orbiter and a transfer to Mercury inspired by the BepiColombo mission. For both problems, the proposed approach proves to be able to explore automatically the vast solution space, producing a large number of trajectories in a large range of final mass and flight times, proving the possibility to apply global optimization techniques directly to the low-thrust problem.

Optimal Earth-Moon Trajectories Using Nuclear Electric Propulsion

Minimum-fuel, two-dimensional and three-dimensional, Earth ‐moon trajectories are obtained for a nuclear electric propulsion spacecraft. The initial state is a circular low Earth parking orbit and the terminal state is a circular low lunar parking orbit. For the three-dimensional problem, the lunar orbit is inclined 90 deg to the Earth‐moon plane. The trajectory has a e xed thrust-coast-thrust engine sequence and is governed by the classical restricted three-body problem dynamic model. An analytical expression for quasicircular low-thrust transfers is usedtoapproximateandreplacethehundredsofinitialande nalorbitsabouttheEarthandmoon.Acostate-control transformation is also utilized to enhance convergence to the optimal solution. Numerical results are presented for the optimal Earth ‐moon trajectories.

Arcjet propulsion system for an SP-100 flight experiment

The design and performance of a arcjet nuclear electric propulsion spacecraft, suitable for use in a space reactor power system (SRPS) flight experiment, are outlined. The vehicle design is based on a 92-kW ammonia arcjet system operating at a specific impulse of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed system are described. A 100-kW SRPS is assumed and the spacecraft mass is baselined at 5250 kg, excluding the propellant and propellant feed system. A radiation/arcjet efflux diagnostics package is included in the performance analysis. This spacecraft, assuming a Shuttle launch from Kennedy Space Center, can perform a 35-deg inclination change and reach a final orbit of 35,860 km with a 120-day trip time, thus providing a four-month active load for the SRPS. Alternatively, a Titan IV launch could provide a mass margin of 120 kg to a 1000km, 58-deg final orbit in 74 days.

Safety Considerations in a Nuclear Electric Propulsion Spacecraft

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.