Evolutionary use of nuclear electric propulsion

Hack, Kurt J.; George, Jeffrey A.; Riehl, John P.; Gilland, James H.

doi:10.2514/6.1990-3821

Cited by 19 publications

(7 citation statements)

References 9 publications

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“…For the more realistic NEP spacecraft (case 2), the optimal power range is from 3600 to 980 kW, which is consistent with projected NEP power levels. 5 SEP spacecraft exhibits a power range from approximately 72 to 9 k\V, which correspond to the shortest and longest trip times, respectively. Figure 3 shows how the optimal 7 sp increases at a nearly linear rate with increasing trip time.…”

Section: Resultsmentioning

confidence: 99%

Electric-propulsion spacecraft optimization for lunar missions

Kluever

Chang

1996

Journal of Spacecraft and Rockets

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A quick and efficient method for computing the optimal vehicle-trajectory combination for lunar missions using electric propulsion is developed. The problem involves computing the optimal spacecraft sizing parameters and trajectory shaping parameters that maximize the payload for a one-way, fixed-trip-time, planar transfer from circular Earth orbit to circular low lunar orbit. The computational load of the problem is reduced by utilizing universal low-thrust trajectory solutions to approximate the Earth-escape and moon-capture spiral trajectories. Several maximum-payload solutions are obtained for both nuclear electric-propulsion and solar electric-propulsion spacecraft. The nuclear results show a very good match with published exact trajectories. A sensitivity analysis of the assumed electric-propulsion technology level is also performed. Nomenclature a T -thrust acceleration, m/s 2 b = propellant-dependent coefficient c = engine exhaust velocity, km/s D = constant Earth-moon separation distance, km d = propellant-dependent coefficient, km/s g = sea-level gravitational acceleration, m/s 2 7 sp = specific impulse, s Kt -tankage fraction m net = net mass of spacecraft, kg Wp p = mass of power and propulsion system, kĝ prop . = propellant mass, kĝ tank = mass of tank and propellant feed system, kg m 0 = initial mass of spacecraft in Earth orbit, kg P = input power, kW r = radial position, km T/W = thrust-to-weight ratio kapt = powered moon-capture spiral time, days koast = translunar coast time, days £ esc = powered Earth-escape spiral time, days tf = total trip time, days v r = radial velocity, km/s VQ = circumferential velocity, km/s a = power and propulsion system specific mass, kg/kW Y] = thruster efficiency 9 = polar angle, deg IJL = gravitational parameter net = net mass fraction CD = constant Earth-moon system angular rate, rad/s Subscripts e = Earth m = moon LEO = low Earth orbit LLO = low lunar orbit Superscript * = nominal or optimal value

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Section: Resultsmentioning

confidence: 99%

Electric-propulsion spacecraft optimization for lunar missions

Kluever

Chang

1996

Journal of Spacecraft and Rockets

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“…«, = \/e 0 kT e /n e e 2 (4) Ignoring the weak dependence of /;, A on n e , Eq. (2) shows that the electron Hall parameter is proportional to electron temperature raised to the 3/2th power and the magnetic field strength, and is inversely proportional to electron number density.…”

Section: Introductionmentioning

confidence: 98%

Use of permanent magnets to reduce anode losses in MPD thrusters

Gallimore

Kelly

Jahn

1994

Journal of Propulsion and Power

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Results from previous studies indicate that the anode fall, the principal source of anode heating in MPD thrusters, increases monotonically with the electron Hall parameter calculated from electron temperature, number density, and magnetic field data obtained near the anode. In an attempt to reduce the anode fall by decreasing the local electron Hall parameter, a proof-of-concept test was performed in which an array of 36 permanent magnets were embedded within the anode of a high-power quasisteady MPD thruster to decrease the local azimuthal component of the induced magnetic field. The modified thruster was operated at power levels between 150 kW and 4 MW with argon and helium propellants. Terminal voltage, triple probe, floating probe, and magnetic probe measurements were made to characterize the performance of the thruster with the new anode. Incorporation of the modified anode resulted in a reduction of the anode fall by up to 15 V with argon and 20 V with helium, which corresponded to decreased anode power fractions of 40 and 45% with argon and helium, respectively. Nomenclature B = magnetic field strength, T e = elementary charge, 1.6 x 10~1 9 C j a = anode current density, A/cm 2 k = Boltzmann's constant, 1.38 x 1Q-23 J/K m e = electron mass, 9.11 x 10~3 1 kg n e = electron number density, m~3 q a -anode heat flux, W/cm 2 q c = anode heat flux from convection, W/cm 2 q r = anode heat flux from radiation, W/cm 2 T e = electron temperature, K V a = anode fall, V Z = axial direction £ () = permittivity of free space, 8.85 x 10~1 2 F/m 0 = azimuthal angle A = plasma parameter \ e = electron Debye length, m (| > = anode material work function, 4.62 V £l e = electron Hall parameter Introduction I T has long been established that a propulsion system of high exhaust velocity is desirable for interplanetary space missions. A review of the rocket equation shows that in order for a propulsive system not to require an inordinate amount of propellant for a given space mission, its exhaust velocity should be at least of the same order as the characteristic velocity increment required for the mission. For most interplanetary missions, a characteristic velocity increment of at least 10 km/s is necessary. 1 " 4 Because of its potential for achieving extremely large exhaust velocities (>40 km/s), the Presented as Paper 92-3461 at the AIAA/SAE/ASME/ASEE 28th

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“…1 ' 2 Recent low-thrust research has involved a range of piloted and cargo lunar and Mars missions. 3 - 4 Aston 5 has advocated the use of low-thrust propulsion for establishing a permanent lunar colony.…”

Section: Introductionmentioning

confidence: 99%

Optimal low-thrust three-dimensional Earth-moon trajectories

Kluever

Pierson

1995

Journal of Guidance, Control, and Dynamics

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Minimum-fuel, three-dimensional trajectories from a circular low Earth parking orbit to an inclined circular low lunar parking orbit with a fixed thrust-coast-thrust engine sequence are computed for a low-thrust spacecraft. The problem is studied in the context of the classical restricted three-body problem. The minimum-fuel transfer to a polar lunar orbit is obtained by successively solving a sequence of fixed lunar inclination problems. Minimumfuel three-dimensional transfers to a polar lunar orbit with meridian plane constraints are also computed. The minimum-fuel trajectory problems are solved using a "hybrid" direct/indirect method. The hybrid method utilizes the costate time histories to parameterize the three-dimensional thrust steering angle time histories. Numerical results are presented for the optimal three-dimensional Earth-moon trajectories.

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Evolutionary use of nuclear electric propulsion

Cited by 19 publications

References 9 publications

Electric-propulsion spacecraft optimization for lunar missions

Electric-propulsion spacecraft optimization for lunar missions

Use of permanent magnets to reduce anode losses in MPD thrusters

Optimal low-thrust three-dimensional Earth-moon trajectories

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