Electric propulsion for cargo missions

ASEN 5053 Final Project Amal Chandran

May 5, 2007

1 Contents

1. Introduction……………………………………..3 2. Mission Scenario………………………………..4 3. cargo mission mass requirements…………….....5 4. Cargo mission Launcher………………………...6 5. Electric propulsion……………………………....7 6. Solar Electric propulsion………………………..8 7. Nuclear electric propulsion…………………….10 8. Emerging Technologies……………………..…13 9. Conclusion……………………………………..14 10. References……………………………………..15

List of Figures

Fig 1: Mars Cargo Mission [1]………………………………………………….…4 Fig 2: Mars mission stages 3, 4, and 5 [1] ……………………………….…….….5 Fig 3: ARES V launch image Courtesy NASA……………………………..….….6 Fig 4: Electric propulsion systems[2] ……………………………………….….…7 Fig 5:Variation in initial mass in LEO with trip time for different Isp[3] ….….….8 Fig 6: Schematic of SEP cargo vehicle[3] ………………………………………...9 Fig 7: Mass break down for SEP Mars cargo vehicle[3] ……………………….…9 Fig 8: NEXIS 57 cm diameter ion thruster[5] …………………………………....10 Fig 9: Schematic of a NEP cargo vehicle[6] ……………………………………..12 Fig 10. Conceptual design of ALFA2 Thruster[5] ………………………………..13 Fig 11: A Mars Exploration Camp courtesy www.marsproject.com…………..14

List of Tables

Table 1: Mass breakdown for Chemical Rocket………………………………….5 Table 2: Mass breakdown for Nuclear thermal Propulsion……………………….6 Table 3: Mass breakdown for conceptual NEP vehicle………………………….12 Table 4 : CaLiPPSo and ALFA2 comparison [5] ………………………………...13

2 1. Introduction

The human is felt by many to be long overdue. When the first humans walked on the surface of the moon, the exploration of Mars was expected to follow inevitably, but this did not happen. Nevertheless, there remains a strong feeling that a human expedition to Mars will happen someday. One of the major obstacle involved in a human exploration mission to Mars is the huge payloads that have to be transported from Earth and inserted into a Martian orbit. This will require a huge amount of propellant and drive up the cost. A solution to this would be to ferry the propellant for the return journey as well as Martian exploration vehicles in advance on a unmanned cargo space craft into Martian orbit and then later let the piloted crew vehicle rendezvous with it in Martian orbit. Electric propulsion offers a cost effective and promising alternative to conventional propulsion because of the impulses it generates.

A major drawback with Electric propulsion though is the low thrust it generates. In electric propulsion, electric energy from solar cells or nuclear-electric reactor is used to energize the propellant working fluid to yield specific impulses much higher than those available from chemical reactions.[2] This reduces the propellant requirement for a given spacecraft ∆V. Electric propulsion devise are “power limited” since the rate at which energy from the external source is supplied to the propellant is proportional to the mass of the power system. This results in very low thrust levels for a given vehicle mass. Hence electric propulsion vehicles are low acceleration vehicles with low thrust to weight (T/W) ratio. However Electric propulsion systems are ideal for non time critical interplanetary trajectories like cargo transfer missions.

This Research paper analyses the different Electric propulsion techniques and takes an in depth look at a few Mars cargo ferries using electric propulsion that have been proposed in the past.

3 2. Mission Scenario First stage – A cargo transport carrying the landing vehicle and the earth return propellant, which has been assembled in LEO (500 km for this project) will be launched via an expendable vehicle on a minimum energy trajectory.

Second stage- Upon arrival at Mars this vehicle will be placed in Low Mars Orbit (6000 km for this project) to await piloted human crew vehicle. An orbit of 6000 km is the same as that of Mar’s inner moon Phobos. The cargo vehicle could be landed on Phobos and power from the Electric propulsion devices could be used to extract resources such as water from Phobos for production of propellant or other useful materials.

Third stage – Approximately 15 months after the first launch a vehicle carrying the crew members will be launched on a high energy low duration trajectory called a ‘sprint’ trajectory

Fourth stage – Upon arrival at Mars the piloted vehicle will rendezvous with the cargo vehicle and depart for exploration of surface. The crew members remaining in the crew vehicle will transfer the propellant and other cargo from the cargo vehicle to piloted vehicle to prepare for the return journey to Earth.

Fifth Stage – The surface crew will rendezvous with the orbiting piloted vehicle to depart for Earth, arriving about 5 months later.

Figure 1 illustrates the mission scenario.

Fig 1: Mars Cargo Mission [1]

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Figure 2 illustrates Mars Orbit Capture, piloted crew vehicle docking with the cargo, going to Martian surface. Ascent vehicle returning after exploration and finally the crew vehicle returning to Earth.

Figure 2: Mars mission stages 3, 4, and 5 [1]

3. Cargo Mission Mass requirements

Earlier designs proposed a chemical rocket as the piloted crew vehicle which requires huge amount of propellants of the order of ~ 400,000kg

Table 1: Mass breakdown for Chemical Rocket Payload Type Weights (kg) Earth return propellant 400,000 ERV 65,000 Mars excursion module (Lander) 55,000 Aeroshell 10,000 Propellant 50,000 Stage Dry mass 25,000 Total Mass 605,000

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Using a Nucler thermal Propulsion for piloted vehicle results in huge mass savings.For purposes of this project 65 MT payload derived from NASA Human Exploration of Mars Design Reference Mission (DRM) version 3.0 has been used[3].. This payload corresponds to delivery of an Earth Return Vehicle (ERV) into Mars orbit. In the nominal DRM 3.0 A Nuclear thermal propulsion stage is used for Earth escape and trans-Mars Injection (TMI). The NTP stage is then jettisoned and the payload is captured into orbit. Here we assume this function is performed by the cargo transport vehicle [3].

Table 2: Mass breakdown for Nuclear thermal Propulsion Payload Type Weight (kg) ERV 65,000 Mars excursion module (Lander) 55,000 Aeroshell 10,000 Propellant 50,000 Stage Dry mass 25,000 Total Mass 205,000

This payload can be brought in to Mars orbit by a single cargo vehicle or in two cargo ferries [3] and assembled in Mars orbit. One of which has the ERV as the main payload and the other has the Mars excursion module as the primary module an additional propellant mass of 45,000 and stage dry mass of 25,000 will be added in this case, with one vehicle weighing ~140,000 kg (ERV) and the other (ML) weighing ~ 135,000 kg 4. Cargo mission launcher for ETO The Ares V launch vehicle which has a payload capacity of 125,000 kg to LEO is chosen as the launch vehicle. Figure 3: illustrates an ARES V launch. For a Chemical rocket Earth Return Vehicle 6 launches of ARES V followed by in-space assembly is required for assembling the cargo ferry in LEO. For a Nuclear Thermal Propulsion ERV, Each Cargo ferry will have masses of the order of 130 tonnes to 140 tonnes. Each will require two launches each of the ARES V launch vehicle followed by assembly in LEO.

Figure 3: ARES V launch image Courtesy NASA

6 5. Electric Propulsion In electric propulsion, electric energy from solar cells or nuclear-electric reactor is used to energize the propellant working fluid to yield specific impulses much higher than those available from chemical reactions.[2] This reduces the propellant requirement for a given spacecraft ∆V. Electric propulsion devise are “power limited” since the rate at which energy from the external source is supplied to the propellant is proportional to the mass of the power system. This results in very low thrust levels for a given vehicle mass. Hence electric propulsion vehicles are low acceleration vehicles with low thrust to weight (T/W) ratio. Electric propulsion systems are therefore ideal for non time critical interplanetary trajectories like cargo transfer missions. Fig 4 illustrates an electric propulsion system. An Electric Propulsion system consists of a power (solar or nuclear) system power conditioning, thruster and propellant storage and feed subsystem.

Fig 4: Electric propulsion systems[2]

The Mars cargo mission is to be designed for transporting payload from LEO to a 6000 km altitude Low Mars Orbit (LMO). Typical ∆Vs for low T/W LEO to LMO transfers are of the order of 16 km/s[3]. As mentioned earlier, this LMO is at the same altitude as the inner moon, Phobos. An Earth to Mars trip time of 2.2 years to match the Earth-Mars synodic period is chosen. This makes it possible to launch the cargo vehicles during one Trans-Mars Injection (TMI) opportunity, travel to Mars, perform Mars orbit insertion and check out all the payload systems prior to launching the crew during the next mars TMI. The power requirement for these missions are of the order of 1000 – 2000 KW or 1- 2 MW as can be seen from figure 3. So, a one way expendable megawatt-class cargo vehicle is required. For this project an Ion (Herakles) thruster which was proposed for the Jupiter Icy Moons Orbiter (JIMO) mission and is considered a near term advanced thruster, is assumed to be used. Figure 5 plots the Initial mass in LEO against Earth–Mars trip time. It compares the Xe-Herakles thruster at different Isp’s of 6000, 7000 and 8000

7 seconds with the Bi-VHITAL (Very High Isp Thruster with Anode Layers) thruster at Isp’s of 5000, 6000 and 7000 seconds[3].

Fig 5:Variation in initial mass in LEO with trip time for different Isp[3]

Power systems for electric propulsion can be either solar or nuclear power based.

6. Solar Electric Propulsion (SEP)

In solar electric propulsion (SEP), solar photons are converted into electricity by solar cells. Figure 6 shows a conceptual SEP vehicle. It consists of a photo voltaic array that generates the necessary power, a main boom that is used both to support the power system’s radiators and to separate the spacecraft systems and payload from the reactor’s radiation, a power management and distribution (PMAD) system, and the spacecraft bus and payload. The spacecraft bus contains the reaction control system (RCS), various miscellaneous spacecraft systems like telecommunications, etc. and the electric propulsion system. Some electric propulsion options may require the use of a plume shield to protect sensitive spacecraft surfaces from the thrusters’ exhaust plume.

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Figure 6: Schematic of SEP cargo vehicle[3]

6.1 Solar array specific Mass – At an assumed efficiency of 25 percent for the power conditioning system, 3.6 m2 are required per kilowatt of electric power with 17.7 % distribution losses[4]. Weight of the cells and wiring would be about 0.45 kg/m2 and the addition of a supporting structure would raise this figure to 1.0 kg/m2. Thus for SEP’s specific mass of the solar array can be assumed at 3.6 kg/kW. Thus for a 1500 KW power requirement the solar array mass would be ~ 5500 kg. Mass breakdown for an SEP with XE-Herakles thruster at different Isp are given in figure 7.

Figure 7: Mass break down for SEP Mars cargo vehicle[3]

There would have to be at least 2 ARES V launches to launch the complete unit into LEO followed by assembly in Space.

9 6.2 Disadvantages of SEP :

1. Deployment of solar array in orbit poses challenges in structure, dynamics and control. Assuming 3.6 m2 of solar array is required per kilowatt of energy, for 1500 KW we need an area of 5400 m2 of solar cells to be deployed in a LEO. 2. There may be difficulties associated with packing a MW class solar array in a launch vehicle. 3. Power available at Mars will almost be half that at the Earth since solar power is inversely proportional to R2, the distance from the Sun. This might significantly affect extra mission capabilities like further Phobos exploration etc. 4. SEP vehicles have a longer trip time than NEP vehicles because of this reduction in power with distance. 5. The large are of the solar array may represent a significant debris impact concern in the debris rich LEO region. 6. The charged Particles in the Van Allen belt might destroy the Solar panel. The SEP vehicle does remain a viable contender for Mars cargo missions in spite of these disadvantages. 7. Nuclear electric propulsion (NEP)

High Power (> 500 KW) Nuclear Electric propulsion devices have never been built and exist only conceptually. The 57 cm diameter NEXIS (Nuclear Electric Xenon Ion System program) thruster has successfully demonstrated all the performance goals for NASA’ project Prometheus 1 (established in 2003 by NASA to develop nuclear-powered systems for long-duration space missions) Jupiter Icy Moon orbiter mission[5]. The NEXIS engine uses 8 thrusters operating simultaneously each generating 20.5 KW for a total of 164 KW of power. This falls significantly lower than the required power for a LEO to LMO transfer. Figure 8 shows a 57 cm diameter NEXIS engine after assembly in a clean room.

Fig 8: NEXIS 57 cm diameter ion thruster

10 7.1 A conceptual NEP Design for 1.5 MW.

The conceptual NEP vehicle configuration shown in fig 9 is based on the use of three SP- 100 Nuclear reactors (with Rankine dynamic power conversion) power modules[6] along with Li- MPD thrusters. In NEP vehicles the payload and the power processing module which contains the power processing unit as well as other spacecraft subsystems such as ACS navigation and propulsion systems, guidance etc are kept at a 24 m distance from the reactor and power conversion systems to minimize the radiation and thermal effects of the power system on the PPM and the payload. Similarly a 25 m distance is used between the PPM and the Li MPD thrusters in order to minimize contamination of the payload or the PPM radiator from the thrusters exhaust plumes.

Fig 9: Schematic of a NEP cargo vehicle[6]

The power system uses a dynamic cycle to convert thermal power from an SP-100 reactor into electricity for use by the MPD thrusters. The power modules are sized so as to provide a net power of at least 1.5 MW to the MPD thrusters after losses in the power system were accounted for. A maximum SP-100 reactor thermal power of 2.4 MW and outlet temperature of 1355 K and a minimum full power reactor operating life of 7 years are assumed.

7.2 Disadvantages of NEP

One disadvantage of a NEP vehicle is that the vehicle has to be based a Nuclear Safe Orbit (NSO) of typically 800 to 1100 km. This high altitude is required to ensure that, in the event of a catastrophic failure, there will be sufficient on-orbit stay time for radioactive components to decay to acceptable levels before reentering Earth’s atmosphere. Therefore a chemical propulsion system is required to put the vehicle in a 1000 km NSO.

11 The mass breakdown for the NEP vehicle is given in table 3.

Table 3: Mass breakdown for conceptual NEP vehicle. System Weight (tonnes) Power system 20 Reactor Booms 5 Power processing 5 Thruster booms 5 MPD thrusters 5 Li propellant tankage & feed 1 Li-MPD propellant 50 Chemical propulsion system 24 Payload 90 Total vehicle Mass 205

So as compared to the SEP system for the same Power this NEP design has a higher Mass mainly because of the extra chemical propulsion system that has to be carried to put it into a NSO. Generally Rankine dynamic power conversion system has a specific mass ranging from 3 kg/KW at 10 MW to 1.9 kg/KW at 200 MW. Therefore at high power requirements (when compared to 1-2 MW) there is the potential for this design to have significant mass and trip time savings over SEP.

12 8. Emerging technologies in High power thrusters exploration missions[5]

1. VHITAL – Very High Isp Thruster with Anode Layer is a 25-36 KW bismuth – fed Hall thruster.

2. ALFA2- Advanced Lithium-Fed, applied-field Lorentz Force Accelerator, 250 KW electromagnetic accelerator with an applied magnetic field being developed by JPL.This is a high power propulsion system that will enable many medium and far range missions and is highly suitable for Lunar and mars cargo missions as well as piloted mars missions and outer planet missions

Fig 10. Conceptual design of ALFA2 Thruster

3. CaLiPPSo - Cargo Vehicle Lithium Plasma Propulsion System, 500 KW electromagnetic thruster with self-magnetic fields. Its development is undertaken by Boeing in collaboration with JPL. Such high power, self field LFA thrusters is designed to provide the optimum Isp for Lunar and Mars cargo missions. The CaLiPPSo thruster will be all refractory metal construction that is radiation cooled. Unfortunately the CaLiPPSo program is currently on hold till interest in extending LFS technology to the 0.5 – 1 MW level remerges.

Table 4 compares CaLiPPSo and ALFA2

Table 4 : CaLiPPSo and ALFA2 comparison [5] ALFA2 CaLiPPSo Power/Thruster (KW) 250 500 Efficiency (%) 60-63 >60 Isp 6200 4500 Lifetime (years) >3 > 1

13 9. Conclusion

This research paper has tried to study electric propulsion systems available and those that have been proposed for a cargo ferry to Mars. If in the future there is a concerted effort for human exploration of Mars, a cargo ferry that will carry the Earth-return fuel for the crew as well as the Martian lander module in a non time critical trajectory to Mars is a logical design.

Electric propulsion is a viable concept for a cargo ferry design. Both SEP and NEP vehicles can function as cargo ferries. SEP designs with Ion thrusters and ALFA2 systems have comparable mission performances in terms of mass and trip time even without further advancements. SEP designs require robust solar array design and deployment in space. But with NEP designs there is no off the shelf technology available at present that can fulfill this requirement. Considerable development in a number of fields will have to be made for this. Development of high power thrusters such as CaLiPPSo will also have to be completed.

The availability of a heavy launcher such as the proposed ARES V that can put payloads of the order of 125 tonnes in LEO is a major requirement, but with the next Lunar program proceeding for a 2020 deadline there is every reason to hope that it will be used as a stepping stone for further solar system exploration.

Fig 11: A Mars Exploration Camp courtesy www.marsproject.com

14 10. References:

1. Darooka, D., Vincente, F., “Nuclear Electric Propulsion Some mission enabling concepts for Mars and Lunar Cargo Transport”, Presented at the AIAA/ASME/SAE/ASEE 26th Joint propulsion Conference, Orlando, FL, 16-18 July, 1990

2. Frisbee, R.H, Advanced Space Propulsion for the 21st century, Journal of propulsion and PowerVol.19, No. 6, Nov-Dec 2003

3. Frisbee, R.H., “Evaluation of High –Power Solar Electric Propulsion using Ion, Hall MPD and PIT thrusters for lunar and Mars cargo missions,” AIAA Paper AIAA-2006-4465, Presented at the AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Sacramento, California, 9-12 July 2006.

4. Frisbee, R.H., “Multi-megawatt Electric Propulsion for Mars missions,” AIAA Paper AIAA-91-3490, Presented at the AIAA/NASA/OAI Conference on Advanced SEI Technologies, Cleveland, Ohio, 4-6 September 1991

5. Goebel, D.M.,Katz, I., Ziemer, J., “Electric Propulsion Research and Development at JPL”, AIAA Paper AIAA-2005-3535, AIAA/ASME/SAE/ASEE Joint propulsion Conference and exhibit, Tucson, Arizona, July 2005

6. Frisbee, R.H., Hoffman, N.J., “Electric Propulsion options for Mars missions,” AIAA Paper AIAA-96-3173, Presented at the AIAA/ASME/SAE/ASEE Joint propulsion Conference and exhibit, Lake Buena Vista, FL, 4-6 September 1996

7. Martin J.L. Turner, “Expedition Mars”, Springer 2004

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