Robert H. Frisbee

SP-100 nuclear electric propulsion for Mars cargo missions

This paper summarizes an evaluation of mission performance (in terms of vehicle mass and trip time) of the use of the near-term SP-100 reactor technology for nuclear electric propulsion for Mars cargo missions, and of the technology requirements for the propulsion and dynamic power conversion systems of the vehicle. The reactor power system uses dynamic power conversion (Rankine), and the propulsion system uses lithium-propellant magnetoplasmadynamic (MPD) thrusters. Three reactor power modules are used to give a total bus power of 1.7 MWe. The total power, power conditioning, and propulsion systems specific mass is 24.8 kg/kWe; the propellant tankage factor is 2.8 percent. The power conditioning system has an efficiency of 90.2 percent and the MPD thrusters an efficiency (electric-to-jet) of 60 percent at a nominal specific impulse of 5000 lb(f)-s/lb(m). Rankine, Brayton, and Stirling dynamic power conversion systems were compared, and the Rankine was found to give the best performance in terms of smallest specific mass and volume; however, it has the longest development time requirement. 6 refs.

The VHITAL Program to Demonstrate the Performance and Lifetime of a Bismuth-Fueled Very High Isp Hall Thruster

In the Very High Isp Thruster with Anode Layer (VHITAL) Program the performance, plume and lifetime capability of the radiatively-cooled two stage, bismuth-fueled VHITAL-160 will be characterized in the US and Russia.

Oxygen-propellant plasma thrusters for cis-lunar electric propulsion missions

This paper presents an analysis of the mission benefits and technology requirements of electric propulsion thrusters designed to use oxygen (O2) as propellant, and an overview of the status of current research in this area. Such engine technology, when used in solar or nuclear electric propulsion transfer vehicles, could enable the use of lunar-produced oxygen propellant as a means of dramatically reducing the mass of vehicles and propellants required to support the build-up of a lunar base. It is found that there are mass benefits, but these benefits are countered somewhat by long trip times due to the low electric-to-jet power efficiency of O2-propellant electric thrusters.

Technology requirements for high-power Lithium Lorentz Force accelerators

Lithium Lorentz Force Accelerators (LFA’s) are capable of processing very high power levels and are therefore applicable to a wide range of challenging missions. An analysis of a reusable orbit transfer vehicle with a solar or nuclear electric power source was performed to assess the applicability of high-power LFA’s to this mission and to define engine performance and lifetime goals to help guide the technology development program. For this class of missions, the emphasis must be on achieving high efficiency at an Isp of 4000–5000 s at power levels of 200–250 kWe. The engines must demonstrate very reliable operation for a service life of about 3000 hours. These goals appear to be achievable with engine technologies currently under development.

Evaluation of cryogenic power conditioning subsystems for electric propulsion spacecraft

The power requirement of vehicles designed to transport cargo supporting a piloted expedition to Mars is in the range of megawatts. Therefore, it is imperative that the megawatt class power processing unit designed for high-power nuclear electric propulsion vehicles using turboalternators and advanced magnetoplasmadynamic (MPD) thrusters should be such that the overall system efficiency is as high as possible with minimum system specific mass. This paper examines the use of cryogenic power conditioning subsystems to achieve that goal since they have very high efficiency. In the past, cryogenic power conditioning have shown complexity of design and implementation and were costly and somewhat uncertain. With recent advances in materials, devices used in power conversion and cooling methods, further improvements in efficiency and specific mass are realizable. Cryogenically cooled MOSFETs and MCTs are considered for power conversion and two configurations have been examined. It is found that a system efficiency of 92.67% and specific mass of 9.99 kg/kW/sub e/ can be realized using MOSFET-based cryogenic power conditioning systems for electric propulsion spacecraft using MPD thrusters. With cryogenically cooled MCTs, the specific mass decreases to 9.77 kg/kW/sub e/, but the efficiency also decreases to 90.94%.

SP‐100 Dynamic Power and Lithium‐Propellant MPD Nuclear Electric Propulsion Technology Requirements

The objective of this study was to evaluate the requirements (including system integration, design, test requirements, and schedule) for the propulsion and power conversion systems of a nuclear electric propulsion (NEP) vehicle using an SP‐100 reactor with a dynamic power conversion system, Li‐propellant magnetoplasmadynamic (MPD) thrusters, Li‐propellant storage and feed systems, and the power conditioning electronics required to convert the power output from the power system to the form (voltage, current) needed by the thrusters. Potassium‐Rankine power conversion systems have the potential for the greatest mission benefit in terms of minimum mass and volume (as compared to Brayton or Stirling power conversion systems), but they require the most development. High‐current, low‐voltage turboalternators are needed for the MPD thruster system envisioned here, although one alternative would be to use more near‐term high‐voltage alternators at the potential cost of higher rectifier losses or added transformer...

Design of a Retro Rocket Earth Landing System for the Orion Spacecraft

The Orion spacecraft is tasked with taking humans to space and returning them safely to Earth at a site in the continental United States. To ensure that the crew returns safely to Earth, a landing system must be designed to reduce the terminal velocity so that, upon landing, the impact loads are minimal and the capsule does not roll over. The retro rocket landing system presented here has been currently baselined because of its performance capabilities over other landing attenuation options such as airbags and landing struts. Preliminary fault analysis and mass and volume sizing led to a down-select of various rocket configurations to base mounted vertical rockets that would be exposed when the heatshield separated. Additionally, the design contains horizontal rockets through the backshell that would be exposed by blowout ports. The trades performed for the retro rocket system led to a design consisting of four 7000 lbf vertical rockets sized to reduce the vertical velocity from 25 ft/s to 5 ft/s. Additionally, four 9400-lbf horizontal rockets were sized to reduce horizontal velocities ranging from 0 to 58 ft/s to less than 20 ft/s. This paper summarizes the analysis that led to the selection of this concept as a preliminary design. This design will change as the Orion spacecraft design and landing system requirements are modified.

Solar Sails for Mars Cargo Missions

This paper presents an analysis of Solar Sails for the Mars Cargo Mission. The figures-of-merit used are the total system Initial Mass in Low Earth Orbit (IMLEO) and Trip Time. The total IMLEO includes the payload, solar sail, and any orbit transfer vehicle (OTV) required to move the sail and its payload to the operational altitude of the sail (e.g., 2.000 km minimum altitude Earth orbit for the solar sail due to air drag). Once the sail and its payload are transported to the sail’s minimum operational orbit by the OTV, the sail begins its Earth-escape spiral and heliocentric transfer to the orbit of Mars. In order to minimize the payload’s Earth-to-Mars trip time, the sail does not perform a Mars orbit insertion capture spiral but rather deploys its payload during Mars flyby. The payload then aerobrakes into Mars orbit or to the surface to await arrival of the crewed portion of the mission. The sail loiters in heliocentric space until it is time to return to Earth. Note that one important constraint on t...

Summary of the NASA/JPL workshop on advanced quantum/relativity theory propulsion

NASA and the Jet Propulsion Laboratory (JPL) sponsored a workshop on advanced quantum/relativity theory propulsion in May 1994 to consider the possibilities of faster-than-light (FTL) travel and/or communication. The workshop specifically focused on three “scientific windows” that might permit FTL travel: (1) tunnels through spacetime; (2) a hypothetical physics where the speed of light is a lower bound; and (3) the physics of additional space dimensions. A number of open issues in physics were noted that may leave open the possibility of FTL travel or communication although no obvious method to achieve such travel or communication was found. Several experiments were identified that would help clarify the possible existence of FTL phenomena.

Orion spacecraft nominal and contingency earth landing retro rocket system options

Retro rocket Earth landing system architectures for the Orion crew module were developed for the Constellation Program during the Landing System Advanced Development Project design trades. The architectures include both variable and fixed impulse base-mounted solid-propellant retro rocket motors for vertical and horizontal landing velocity control. Monte Carlo performance analyses were conducted, detailed motors were designed and support system configurations were developed for mass and performance characterization. Preliminary penetration schemes for blow-out ports through the TPS were designed for the motor exhaust plumes. The requirements, configurations, elements, mass and performance of the retro rocket landing system architectures for multiple landing scenarios are discussed.