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Electric orbit transfer vehicle cryogenic propellant system
Electric orbit transfer vehicle cryogenic propellant system* J.R. Schuster, C.T. Huynh and G.E. Williams General Dynamics Space Systems Division, San Diego, CA 92123, USA
An electric orbit transfer vehicle (EOTV) is intended to transfer payloads from low Earth orbit (LEO) to higher orbits using low-thrust solar-electric propulsion and hydrogen propellant. Because of its high specific impulse and synergistic sharing of power supply, attitude control and communication systems with the payload, the highly efficient EOTV transfer stage permits use of a smaller, less costly launch vehicle than if orbit transfer were accomplished using chemical propulsion. Study of the propellant storage and supply system for an EOTV intended to fly a 168 day spiral trajectory from LEO to geosynchronous orbit (GEO) reveals that the low propellant flow rate needed by the thrusters can be supplied by the boil-off from the storage tank, eliminating the need for any overboard venting. The tank can be fabricated under the same pressure-stabilized, thin, stainless steel monocoque construction as the current Centaur upper stage, and insulated with Centaur fixed foam and MLI. The tank contains a thermodynamic vent system (TVS) for control of tank pressure in zero and low gravity and for supply of propellant to the thrusters. An external compressor, accumulator and regulator condition the hydrogen boil-off provided by the TVS and provide for start-up and shut-down transients. The resulting system is simple, has a very low structural mass fraction and builds on the Centaur cryogenic upper stage technology, which has been operational for over 25 years.
Keywords: space cryogenics; cryogenic propellant systems; electric orbit transfer vehicles
We are in an era of expanding space activity, as the commercial, scientific, tactical and strategic benefits of Earth-orbiting systems are being realized, and the data returns from deep space missions are rapidly increasing our scientific knowledge of the solar system, our galaxy and the universe. The launch and positioning of Earthorbiting satellites has become routine and there are a substantial number of boosters, upper stages, transfer stages, motors and thrusters used worldwide to provide the needed propulsion. For many years orbit transfer vehicles (OTVs) have been studied as an effective means for moving payloads from one orbit to another. The OTV is different from an upper stage in that it is not involved in placing the payload in its initial orbit and therefore the OTV design can be optimized for operation in space. This includes low-thrust propulsion, low structural mass fraction and insulation systems that take advantage of the space vacuum, Most OTV concepts that have been proposed are based on high-energy chemical propulsion, such as used by the hydrogen-fuelled Centaur upper stage. The cur* Paper presented at the 1992 Space Cryogenics Workshop, 1 5 - 16 June 1992, M0nchen, Germany
rent practical upper performance limit for chemical propulsion is specific impulse in the range 4 6 0 - 4 8 0 s. In contrast, a solar-powered electric orbit transfer vehicle (EOTV) using hydrogen propellant and low-thrust arcjet propulsion could provide a specific impulse in the range 1 3 0 0 - 1 4 0 0 s. Because of its high specific impulse and synergistic use of the payload power supply, attitude control and communication systems, the highly efficient EOTV enables a smaller, less costly hmnch vehicle to be used for a given payload. Alternatively, a given launch vehicle could accommodate heavier payloads. This paper addresses the hydrogen propellant storage and supply system for an EOTV intended to fly a six month spiral trajectory from low Earth orbit (LEO) to geosynchronous Earth orbit (GEO). The system concept is simple, has specific features to enable such a long mission with alternating low gravity and zero gravity, and has modest development requirements because it utilizes technology that is currently available.
Mission description The mission addressed involves the use of an EOTV to transport a satellite from LEO to GEO. It is assumed that the EOTV and satellite are launched into LEO with a
0011 2275/93/040423-06 (c) 1993 Butterworth - Heinemann Ltd
Cryogenics 1993 Vol 33, No 4
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EOTV cryogenic propellant system: J.R. Schuster et al. conventional multistage launch vehicle, such as one from the Atlas family of launchers that combine an Atlas booster with a Centaur cryogenic upper stage. Figures 1 and 2 illustrate the EOTV mission and its phases, respectively. After being placed in LEO by the launch vehicle, the EOTV, along with its satellite payload, spirals outward around the Earth until the final orbit is reached. Three options could be considered for the EOTV: 1, the EOTV simply provides transportation and is expended after satellite deployment in GEO; 2, upon reaching GEO, the EOTV 'donates' some of its subsystems to the satellite, such as power, control, etc., prior to satellite deployment; or 3, the EOTV is fully integrated with the satellite, achieving maximum synergism.
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An Atlas launch vehicle is assumed, for which Figure 3 presents the payload capability for a direct ascent (single Centaur burn) to circular orbit. A payload of 7200 kg can be placed in a 500 km orbit. The Atlas IIAS can also perform a parking orbit ascent (two Centaur burns), and with this approach can place 7200 kg in a 1000 km orbit or 7800 kg in a 500 km orbit.
EOTV and transfer trajectory parameters A number of analytical tools are being used throughout the aerospace community to estimate the trajectories and performance of low-thrust space transfer vehicles. One
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Figure 3
At/a_s IIAS direct ascent to circular orbit
tool particularly suited for parametric orbit transfer vehicles studies is the electric vehicle analyser (EVA) code ~. This code was used to estimate EOTV design parameters and determine a trajetory to be used as a basis for studying the EOTV liquid hydrogen (LH2) propellant storage and management features. Table 1 summarizes the mission trajectory and vehicle parameters from the EVA analysis. A total mass (EOTV plus satellite) of 7200 kg and a 500 km initial circular orbit were assumed. For a total thrust of 2.79 N, a specific impulse of 1250 s and a total trip time of 168 days, a satellite weighing 2410 kg can be transferred to GEO using an EOTV having a dry mass of 2010 kg and 2780 kg of LH2 propellant. Figure 4 presents the history of the transfer trajectory parameters. Table 1
EVA analysis results
Orbit parameters
Figure 1
EOTV mission profile
® EOTI//MTEU.ITEIF.GIN OBIIIT RNISING 0 500 KM, 28.5 DEG
® ~
® EOTV/SATELUTE SPIRALTO FINAL ORBIT
All.AS IIAS DELIVIDRS EoI~r~ATLqi ITE TO 800 KM, 29.5 DEG
® - PLAIIEClMNGE BElaJlS@ 000 IOA,29.5 9EG
mP4~ATWJ~ IUIN~ @ 3S775 I(M, 0 mEG
Figure
424
2
Major phases of an EOTV mission to GEO
Cryogenics 1993 Vol 33, No 4
LEO orbit fraction in s h a d o w /3 = 0.0 ° /3 = 52 ° GEO orbit fraction in s h a d o w /3 = 0.0 ° /3 = 2 3 . 5 ° Total number of orbits Total trip time Total time in sun Total time in s h a d o w Initial orbit Final orbit Required A V
3 8 % (35.7 min) 3 0 % (28 min) 5% (71.6 min) 0 . 0 % (0.0 rain) 1200 168 days 142 days 26 days 500 k m / 2 8 . 5 ° inclination 35 8 0 0 k m / 0 . 0 ° inclination 5990 m s 1
Spacecraft parameters Specific impulse Total thrust Total p o w e r Excess p o w e r Total initial mass Propellant mass EOTV mass (dry) Net payload
1250 S 2.79 N 4 2 . 0 kWe 2.0 kWe 7 2 0 0 kg 2 7 8 0 kg 2 0 1 0 kg 2 4 1 0 kg
EOTV cryogenic propellant system: J.R. Schuster e t
al.
Space heating The EOTV hydrogen propellant tank is subject to environmental heating while in space, due to solar radiation, reflected solar radiation off the Earth and thermal radiation emitted by the Earth. The outer surface of the tank system will also emit thermal radiation, and the surface temperature will vary with location around the tank and with time, due to the angular relationships with the Earth and Sun and the distance from the Earth. The tank will be insulated so that boil-off rate is limited and the surface temperature of the insulation will depend on the ratio of surface solar absorptivity to thermal emissivity
300 280 260240 220w =c .~200-
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System
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The EVA code analysis assumes that the propulsion system is used only when the EOTV solar panels are illuminated by sunlight. Therefore, the propulsion system cycles off and on as the EOTV and its payload pass in and out of the Earth's shadow as it spirals from LEO to GEO. The amount of time per orbit that the spacecraft spends in the Earth's shadow depends on the altitude and the beta angle (the angle the solar vector makes with the orbital plane). As indicated in Table 1, for a launch latitude of 28.5 °, depending on time of year, the time in the shade can vary from 30% (28 min) to 38% (35.7 rain) in LEO and between 0% and 5% (71.6 rain) at GEO. The propellant storage and management system contains the liquid hydrogen propellant required for the mission. It has the features needed to store the cryogenic
ALTITUDE HISTORY 40
INCLINATION HISTORY
~ 20
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Figure 5
Tank environmental sink temperature
propellant and supply it to the propulsion system in gaseous form at the appropriate temperature, pressure and flow rate. Due to the nature of the spacecraft trajectory, the propellant management and storage system must be capable of periodic start-up and shut-down, and operation in a fluctuating and gradually changing thermal environment. Its functional requirements include thermal control, pressure control, slosh control, propellant supply and propellant quantity and flow rate gauging. Most system functions can be verified through ground testing; however, both pressure control and slosh control are sensitive to gravity and may require flight testing for final verification. Figure 6 illustrates the system. The tank is about the same size as the hydrogen tank for the Centaur system now in use and builds on the cryogenic upper stage technology that has been operational for over 25 years. Built-in control features provide hydrogen boil-off at the rate used by the propulsion system: no propellant is wasted via venting to space. The tank is insulated to limit boil-off rate in both the ground and space environment. It also contains a thermodynamic vent system (TVS) and baffles to control tank pressure and liquid sloshing and supply propellant for the propulsion system. An external compressor, accumulator and regulator are required to condition the hydrogen gas vented from the tank and provide for startup and shut-down transients. As altitude increases, radiant heating from the Earth decreases and a tank heater is used to maintain the boil-off rate at required levels. Table 2 provides an estimated mass summary of the propellant storage and management system. An inert mass fraction of 15% should be easily achievable.
o
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Figure 4
/
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(c~/e). For conceptual estimates it has been found that for a well insulated tank the equivalent sink temperature of the outer surface can be used to analyse thermal management features and boil-off. The General Dynamics Vector Sweep computer code 2 was used to estimate environmental heating both in LEO and GEO, and the various orbital-averaged heat fluxes were combined and converted into orbital-averaged equivalent sink temperature, which was also averaged over the tank surface. Figure 5 presents tank insulation system effective surface temperature as a function of c~/e for both the initial and final orbits. As the EOTV spirals outward from the Earth, the tank experiences increasing direct solar heat load because of decreasing shadowing by the Earth; however, the albedo and Earth thermal heat loads simultaneously decrease, resulting in a net decrease in effective surface temperature.
/
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Transfer trajectory parameters
40
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The 42.8 m 3 tank has a diameter of 3.05 m and an overall length of 6.51 m with 1.38 ellipsoidal domes. It can be derived from the Centaur tankage illustrated in Figure 7, which uses a pressure-stabilized, thin stainless steel, monocoque construction 3. It is conceivable that composites might be lighter, but a composite tank would require considerable development. Centaur tankage is the lightest weight upper stage tank structure yet flown 4 and using the same diameter and tank domes as Centaur would enable use of common tooling and manufacturing
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EOTV cryogenic propellant system: J.R. Schuster et al. i • _ GROUNO • VENT MASSFLOW CONTROLLERS
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Figure 6 System schematic
methods, thus lessening considerably the required development for the EOTV tank.
Tank thermal control The insulation technology applied to Centaur has improved over the years to include foam insulation to limit boil-off on the launch pad and high performance multilayer insulation (MLI) of the radiation shield type tq provide low boil-off in space. An operational feature now used for Atlas/Centaur is closed-cell foam insulation bonded directly to the tank outer surface. This has permitted the elimination of the launch pad helium purge. The recommended insulation system for the EOTV hydrogen tank is therefore MLI over a layer of the bonded-on foam currently used on Centaur. It is assumed that the insulation system is protected from ascent heating and aerodynamic forces by an aerodynamic fairing that surrounds the EOTV and its satellite payload. The insulation system design must provide the capability to control boil-off rate over the entire mission such that propellant feed requirements of the propulsion system can be met without having to vent boil-off to space to achieve tank pressure control• This means that the heating environment in LEO must not produce an average boil-off rate greater than that needed for propulsion. Since the thrusting time per orbit increases with beta angle, and the tank heating also increases, the required amount of insulation is relatively insensitive to beta angle• Referring to Figure 5, for a beta angle of 0.0, a conservative value of 0.3 for c¢/e results in a value of 221 K for insulation system surface temperature. This
then becomes the basis, along with the orbital average propellant feed rate in LEO, for selecting the insulation. Figure 8 illustrates the ideal heat leak through high density MLI as a function of surface temperature and thickness. It accounts for edge seams, and conduction through spacers and mounting pins, but does not account for degradation of performance due to handling and installation. Allowing for half the tank heat leak to occur through tank structural and fluid penetrations, and doubling the ideal MLI heat leak to account for handling and installation, = 25 mm of MLI will produce the right boil-off rate in LEO. A tank electrical heater with a capacity of 100 W is used to adjust the boil-off rate to the desired value. At GEO the nominal dissipation of the heater is estimated to be 18 W. The Centaur bonded-on foam has a nominal thickness of 15 ram. For MLI/foam systems the foam thickness must be great enough to prevent ice formation at the MLI/foam interface while on the launch pad. Making the conservative assumption that gaseous conduction dominates on the launch pad, the foam should be at least half as thick as the MLI. The tank is vulnerable to damage by micrometeoroids and debris, and analyses may show that the protection offered by the insulation must be augmented by additional shielding. This can be achieved by incorporating a thin aluminium bumper or extra-heavy scrim rein-
Table 2 Propellant storage and management system mass summary 0.75,
Item
Mass (kg)
Hydrogen tank Insulation system Micrometeroroid/debris bumper Tank supports Thermodynamic vent system Compressor Accumulator Valves and plumbing
263 151 27 26 7 1 4 23
Total
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502 Figure 8
426
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Cryogenics 1993 Vol 33, No 4
13MMMLf
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EOTV cryogenic propellant system: J.R. Schuster et al. forcement in the outermost layer of the insulation system.
Tank pressure control Tank pressure control for the EOTV is achieved through mixing the tank contents to counter thermal stratification, and venting tank vapour to provide tank cooling. This vented vapour is the 'boil-off' commonly referred to. The Centaur uses its reaction control system (RCS) thrusters to provide --~ 10-3g to settle the liquid to enable venting. The EOTV will have much lower acceleration ( < 1 0 - 4 g ) and periods of zero-g coast while passing through the Earth's shadow, making it uncertain where the vapour is located in the tank. Additionally, there may be times during the EOTV mission when it will be necessary to coast for much longer periods. Thus, the EOTV should utilize a thermodynamic vent system (TVS). A passive TVS consists of a J o u l e - T h o m s o n ( J - T ) valve and a heat exchanger; an active TVS also includes a tank mixer. The heat exchanger of the passive TVS is vent tubing attached to the tank wall. An active TVS can use either wallmounted tubing or a compact heat exchanger. Figure 9 shows an integrated active TVS developed for the Shuttle/Centaur that was mounted inside the tank and incorporated a compact heat exchanger and mixer. This TVS incorporated a regulator that allowed it to accept liquid, vapour or two-phase fluid, while ensuring that the vent flt/id exiting the TVS was pure vapour. It weighed 17 kg, and was capable of venting over 3 g s ~. The = 0 . 2 g s ] system required for the EOTV could be much smaller. The Shuttle/Centaur TVS was extensively ground tested but never flown.
Liquid dynamics and slosh control The EOTV is expected to roll periodically to realign the solar panels, and this is expected to cause fluid motion and excite propellant slosh modes when coupled to vehicle acceleration. In addition, liquid tends to cover the walls of the tank in zero-g, and will reorient and slosh when the thrust is initiated upon exit from the Earth's shadow. This propellant motion can lead to shifts in vehicle centre-of-mass, affecting attitude and pointing. Tank ring baffles will probably be needed to limit propellant motion.
Propellant supply The tank thermal control and pressure control systems will be designed to achieve the proper boil-off characteristics to meet the overall propellant needs of the propulsion system. However, in order to supply propellant on demand at the temperature, pressure and flow rate needed, three additional elements are required: a compressor, an accumulator and a flow regulator. The compressor raises the pressure of the hydrogen vapour from the TVS to a pressure level adequate to provide for stable, efficient thruster operation. It enables a lightweight, low-pressure hydrogen tank to be used. The pressure ratio for the compressor would be in the range 2 - 4 and would require power of the order of 100 W. This piece of equipment should be quite small and light, and could possibly be a derivative of hardware used for cryogenic sensor cooling.
Figure 9
Shuttle/Centaur thermodynamic vent system
The accumulator is a pressure vessel used to store compressed, gaseous hydrogen boil-off, damp out any supply fluctuations and provide for smooth start-up and shut-down of the propulsion system. Its size will depend on the amount of storage desired, but 0.25 m 3 should be sufficient to store enough hydrogen to operate the thrusters for an hour. This vessel is probably an off-theshelf, flight-qualified bottle. The flow regulator actually regulates pressure and thus regulates flow rate according to the pressure drop characteristic of the propulsion system. The desire to maintain a fixed flow rate could be met with either a mechanical or electronic pressure regulator. A mechanical pressure regulator is used on Centaur to control tank venting. Electronic pressure regulators have been developed for low-temperature vapour service, including hydrogen. Some development might be needed, however, to achieve control for the relatively low flow rates needed for this application, and the regulator would need to be flight-qualified. An alternative approach could be to use a flow network of parallel fixed orifices, with each branch controlled by a shut-off valve.
Quantity and flow gauging Gauging of the quantity of hydrogen in the tank and the flow rate to the propulsion system will be needed on the EOTV to enable spacecraft control and support mission operations. Off-the-shelf gauges of several types are available for accurate measurement of hydrogen vapour flow rate in the range needed. If there is not an appropriate flight-qualified gauge, qualification should be straightforward. Mass gauging the hydrogen in the storage tank might be done with the hydrogen liquid settled while the thruster is operating, using the same capacitance method employed by the Centaur, if the capacitance probe will drain under the imposed acceleration. Alternative laser and microwave sensors have been demonstrated in the laboratory and might be flight-qualified for future application to an EOTV. Some development work has been done in industry on mass gauging techniques for zero-gravity conditions. At this time, however, no zerogravity technique is sufficiently developed to be assured of its success.
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EOTV cryogenic propellant system: J.R. Schuster et al.
Conclusions A low-thrust, solar-electric orbital transfer vehicle is a concept for highly efficient transfer of payloads from LEO to GEO. A high specific impulse and synergistic sharing of subsystems between the E O T V and its payload allows a smaller launch vehicle to be used and/or a heavier payload to be launched than for current LEO-to-GEO propulsion stages. The hydrogen propellant storage and supply system for the E O T V is simple and has modest development requirements because it can use currently available technology such as from Centaur tankage, insulation and pressure control systems. Use of this technology could result in an E O T V
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with a small availability.
inert
mass
fraction
and
short-term
References 1 Dickey, M.R., Klucz, R.S., Ennix, K.A. and Matuszak, L.M.
Developmentof the electric vehicle analyzer, Air Force Astronautics Laboratory Report AL-TR-90-006, Edwards Air Force Base, California, USA (June 1990) 2 O'Neill, R.F. and Zich, J.L. Vector sweep user's guide. General DynamicsReportGDSS-SP-85-011,GeneralDynamicsSpaceSystems Division, San Diego, California, USA (December 15 1985) 3 Dunbar, D.R. Mission planner's guide for the Atlas launch vehicle family, revision 2, General Dynamics Commercial Launch Services, Inc., San Diego, California, USA (July 1990).
An electric orbit transfer vehicle (EOTV) is intended to transfer payloads from low Earth orbit (LEO) to higher orbits using low-thrust solar-electric propulsion and hydrogen propellant. Because of its high specific impulse and synergistic sharing of power supply, attitude control and communication systems with the payload, the highly efficient EOTV transfer stage permits use of a smaller, less costly launch vehicle than if orbit transfer were accomplished using chemical propulsion. Study of the propellant storage and supply system for an EOTV intended to fly a 168 day spiral trajectory from LEO to geosynchronous orbit (GEO) reveals that the low propellant flow rate needed by the thrusters can be supplied by the boil-off from the storage tank, eliminating the need for any overboard venting. The tank can be fabricated under the same pressure-stabilized, thin, stainless steel monocoque construction as the current Centaur upper stage, and insulated with Centaur fixed foam and MLI. The tank contains a thermodynamic vent system (TVS) for control of tank pressure in zero and low gravity and for supply of propellant to the thrusters. An external compressor, accumulator and regulator condition the hydrogen boil-off provided by the TVS and provide for start-up and shut-down transients. The resulting system is simple, has a very low structural mass fraction and builds on the Centaur cryogenic upper stage technology, which has been operational for over 25 years.
Keywords: space cryogenics; cryogenic propellant systems; electric orbit transfer vehicles
We are in an era of expanding space activity, as the commercial, scientific, tactical and strategic benefits of Earth-orbiting systems are being realized, and the data returns from deep space missions are rapidly increasing our scientific knowledge of the solar system, our galaxy and the universe. The launch and positioning of Earthorbiting satellites has become routine and there are a substantial number of boosters, upper stages, transfer stages, motors and thrusters used worldwide to provide the needed propulsion. For many years orbit transfer vehicles (OTVs) have been studied as an effective means for moving payloads from one orbit to another. The OTV is different from an upper stage in that it is not involved in placing the payload in its initial orbit and therefore the OTV design can be optimized for operation in space. This includes low-thrust propulsion, low structural mass fraction and insulation systems that take advantage of the space vacuum, Most OTV concepts that have been proposed are based on high-energy chemical propulsion, such as used by the hydrogen-fuelled Centaur upper stage. The cur* Paper presented at the 1992 Space Cryogenics Workshop, 1 5 - 16 June 1992, M0nchen, Germany
rent practical upper performance limit for chemical propulsion is specific impulse in the range 4 6 0 - 4 8 0 s. In contrast, a solar-powered electric orbit transfer vehicle (EOTV) using hydrogen propellant and low-thrust arcjet propulsion could provide a specific impulse in the range 1 3 0 0 - 1 4 0 0 s. Because of its high specific impulse and synergistic use of the payload power supply, attitude control and communication systems, the highly efficient EOTV enables a smaller, less costly hmnch vehicle to be used for a given payload. Alternatively, a given launch vehicle could accommodate heavier payloads. This paper addresses the hydrogen propellant storage and supply system for an EOTV intended to fly a six month spiral trajectory from low Earth orbit (LEO) to geosynchronous Earth orbit (GEO). The system concept is simple, has specific features to enable such a long mission with alternating low gravity and zero gravity, and has modest development requirements because it utilizes technology that is currently available.
Mission description The mission addressed involves the use of an EOTV to transport a satellite from LEO to GEO. It is assumed that the EOTV and satellite are launched into LEO with a
0011 2275/93/040423-06 (c) 1993 Butterworth - Heinemann Ltd
Cryogenics 1993 Vol 33, No 4
423
EOTV cryogenic propellant system: J.R. Schuster et al. conventional multistage launch vehicle, such as one from the Atlas family of launchers that combine an Atlas booster with a Centaur cryogenic upper stage. Figures 1 and 2 illustrate the EOTV mission and its phases, respectively. After being placed in LEO by the launch vehicle, the EOTV, along with its satellite payload, spirals outward around the Earth until the final orbit is reached. Three options could be considered for the EOTV: 1, the EOTV simply provides transportation and is expended after satellite deployment in GEO; 2, upon reaching GEO, the EOTV 'donates' some of its subsystems to the satellite, such as power, control, etc., prior to satellite deployment; or 3, the EOTV is fully integrated with the satellite, achieving maximum synergism.
Launch
0
100
L~I
I
200 I
NAUTICALMILES 300 400 I I I L
10000-
I
000 I
I
I
600 I
I 20
0 000-
\ •
8 000-
~
\
-18
\
-
-16 i r~ []
-14 ~
5000-
An Atlas launch vehicle is assumed, for which Figure 3 presents the payload capability for a direct ascent (single Centaur burn) to circular orbit. A payload of 7200 kg can be placed in a 500 km orbit. The Atlas IIAS can also perform a parking orbit ascent (two Centaur burns), and with this approach can place 7200 kg in a 1000 km orbit or 7800 kg in a 500 km orbit.
EOTV and transfer trajectory parameters A number of analytical tools are being used throughout the aerospace community to estimate the trajectories and performance of low-thrust space transfer vehicles. One
EOIV DS KXMIIOED
RnMim,I P ~ i
mum1 .....
1.1~41W 1~ m I ~
CHECKOUT EOIVAIIO SATB.UTE ~
.................
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-10 pLANAR~SCENT
I
4000200
400
000
800
1000
1200
CIRCULARORBITALTITUDE(KM)
Figure 3
At/a_s IIAS direct ascent to circular orbit
tool particularly suited for parametric orbit transfer vehicles studies is the electric vehicle analyser (EVA) code ~. This code was used to estimate EOTV design parameters and determine a trajetory to be used as a basis for studying the EOTV liquid hydrogen (LH2) propellant storage and management features. Table 1 summarizes the mission trajectory and vehicle parameters from the EVA analysis. A total mass (EOTV plus satellite) of 7200 kg and a 500 km initial circular orbit were assumed. For a total thrust of 2.79 N, a specific impulse of 1250 s and a total trip time of 168 days, a satellite weighing 2410 kg can be transferred to GEO using an EOTV having a dry mass of 2010 kg and 2780 kg of LH2 propellant. Figure 4 presents the history of the transfer trajectory parameters. Table 1
EVA analysis results
Orbit parameters
Figure 1
EOTV mission profile
® EOTI//MTEU.ITEIF.GIN OBIIIT RNISING 0 500 KM, 28.5 DEG
® ~
® EOTV/SATELUTE SPIRALTO FINAL ORBIT
All.AS IIAS DELIVIDRS EoI~r~ATLqi ITE TO 800 KM, 29.5 DEG
® - PLAIIEClMNGE BElaJlS@ 000 IOA,29.5 9EG
mP4~ATWJ~ IUIN~ @ 3S775 I(M, 0 mEG
Figure
424
2
Major phases of an EOTV mission to GEO
Cryogenics 1993 Vol 33, No 4
LEO orbit fraction in s h a d o w /3 = 0.0 ° /3 = 52 ° GEO orbit fraction in s h a d o w /3 = 0.0 ° /3 = 2 3 . 5 ° Total number of orbits Total trip time Total time in sun Total time in s h a d o w Initial orbit Final orbit Required A V
3 8 % (35.7 min) 3 0 % (28 min) 5% (71.6 min) 0 . 0 % (0.0 rain) 1200 168 days 142 days 26 days 500 k m / 2 8 . 5 ° inclination 35 8 0 0 k m / 0 . 0 ° inclination 5990 m s 1
Spacecraft parameters Specific impulse Total thrust Total p o w e r Excess p o w e r Total initial mass Propellant mass EOTV mass (dry) Net payload
1250 S 2.79 N 4 2 . 0 kWe 2.0 kWe 7 2 0 0 kg 2 7 8 0 kg 2 0 1 0 kg 2 4 1 0 kg
EOTV cryogenic propellant system: J.R. Schuster e t
al.
Space heating The EOTV hydrogen propellant tank is subject to environmental heating while in space, due to solar radiation, reflected solar radiation off the Earth and thermal radiation emitted by the Earth. The outer surface of the tank system will also emit thermal radiation, and the surface temperature will vary with location around the tank and with time, due to the angular relationships with the Earth and Sun and the distance from the Earth. The tank will be insulated so that boil-off rate is limited and the surface temperature of the insulation will depend on the ratio of surface solar absorptivity to thermal emissivity
300 280 260240 220w =c .~200-
~
System
description
The EVA code analysis assumes that the propulsion system is used only when the EOTV solar panels are illuminated by sunlight. Therefore, the propulsion system cycles off and on as the EOTV and its payload pass in and out of the Earth's shadow as it spirals from LEO to GEO. The amount of time per orbit that the spacecraft spends in the Earth's shadow depends on the altitude and the beta angle (the angle the solar vector makes with the orbital plane). As indicated in Table 1, for a launch latitude of 28.5 °, depending on time of year, the time in the shade can vary from 30% (28 min) to 38% (35.7 rain) in LEO and between 0% and 5% (71.6 rain) at GEO. The propellant storage and management system contains the liquid hydrogen propellant required for the mission. It has the features needed to store the cryogenic
ALTITUDE HISTORY 40
INCLINATION HISTORY
~ 20
10
~s
NOTE: TANK AXIS ALIGNED WITH VELOCITY VECTOR - LEO, 13 = 52 DEG - - - - - - LEO, ~ = 0.0 DEG ----GEO
120100-
o'.,
0'.0
;.,
1'.o
ode
Figure 5
Tank environmental sink temperature
propellant and supply it to the propulsion system in gaseous form at the appropriate temperature, pressure and flow rate. Due to the nature of the spacecraft trajectory, the propellant management and storage system must be capable of periodic start-up and shut-down, and operation in a fluctuating and gradually changing thermal environment. Its functional requirements include thermal control, pressure control, slosh control, propellant supply and propellant quantity and flow rate gauging. Most system functions can be verified through ground testing; however, both pressure control and slosh control are sensitive to gravity and may require flight testing for final verification. Figure 6 illustrates the system. The tank is about the same size as the hydrogen tank for the Centaur system now in use and builds on the cryogenic upper stage technology that has been operational for over 25 years. Built-in control features provide hydrogen boil-off at the rate used by the propulsion system: no propellant is wasted via venting to space. The tank is insulated to limit boil-off rate in both the ground and space environment. It also contains a thermodynamic vent system (TVS) and baffles to control tank pressure and liquid sloshing and supply propellant for the propulsion system. An external compressor, accumulator and regulator are required to condition the hydrogen gas vented from the tank and provide for startup and shut-down transients. As altitude increases, radiant heating from the Earth decreases and a tank heater is used to maintain the boil-off rate at required levels. Table 2 provides an estimated mass summary of the propellant storage and management system. An inert mass fraction of 15% should be easily achievable.
o
4o
80
12o 1;o
Propellant tank o
40
TRIP TIME (DAYS)
so
12o 1~o
TRIP TIME (DAYS)
/"V HISTORY 8
CUMULATIVE ORBITS 1500
-°L,o, o
°°l,OooOo 40
-co l~iO 16o
o
TRIP TIME (DAYS)
Figure 4
/
140-
30
~, 20 0
180160-
(c~/e). For conceptual estimates it has been found that for a well insulated tank the equivalent sink temperature of the outer surface can be used to analyse thermal management features and boil-off. The General Dynamics Vector Sweep computer code 2 was used to estimate environmental heating both in LEO and GEO, and the various orbital-averaged heat fluxes were combined and converted into orbital-averaged equivalent sink temperature, which was also averaged over the tank surface. Figure 5 presents tank insulation system effective surface temperature as a function of c~/e for both the initial and final orbits. As the EOTV spirals outward from the Earth, the tank experiences increasing direct solar heat load because of decreasing shadowing by the Earth; however, the albedo and Earth thermal heat loads simultaneously decrease, resulting in a net decrease in effective surface temperature.
/
o..
Transfer trajectory parameters
40
I ao
12o 16R
TRIP TIME (DAYS)
The 42.8 m 3 tank has a diameter of 3.05 m and an overall length of 6.51 m with 1.38 ellipsoidal domes. It can be derived from the Centaur tankage illustrated in Figure 7, which uses a pressure-stabilized, thin stainless steel, monocoque construction 3. It is conceivable that composites might be lighter, but a composite tank would require considerable development. Centaur tankage is the lightest weight upper stage tank structure yet flown 4 and using the same diameter and tank domes as Centaur would enable use of common tooling and manufacturing
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EOTV cryogenic propellant system: J.R. Schuster et al. i • _ GROUNO • VENT MASSFLOW CONTROLLERS
ACCUMULATOR COMPRESSOR~ Q I J A N T I T Y S A G E
FOAM,MU A WM/DSHIELD J-T VALVE ~
Ir HYDROGEN suscRmcALi f I I =
__
~
'
~l__~
-
~
~
ELECTRICTHRUSTERS
i-1
J.~-PROPELLANT STORACE TANK BAFFLES PACT THERMOOYNAMIC VENTSYSTEMA TANKMIXER FILLLINE
Figure 6 System schematic
methods, thus lessening considerably the required development for the EOTV tank.
Tank thermal control The insulation technology applied to Centaur has improved over the years to include foam insulation to limit boil-off on the launch pad and high performance multilayer insulation (MLI) of the radiation shield type tq provide low boil-off in space. An operational feature now used for Atlas/Centaur is closed-cell foam insulation bonded directly to the tank outer surface. This has permitted the elimination of the launch pad helium purge. The recommended insulation system for the EOTV hydrogen tank is therefore MLI over a layer of the bonded-on foam currently used on Centaur. It is assumed that the insulation system is protected from ascent heating and aerodynamic forces by an aerodynamic fairing that surrounds the EOTV and its satellite payload. The insulation system design must provide the capability to control boil-off rate over the entire mission such that propellant feed requirements of the propulsion system can be met without having to vent boil-off to space to achieve tank pressure control• This means that the heating environment in LEO must not produce an average boil-off rate greater than that needed for propulsion. Since the thrusting time per orbit increases with beta angle, and the tank heating also increases, the required amount of insulation is relatively insensitive to beta angle• Referring to Figure 5, for a beta angle of 0.0, a conservative value of 0.3 for c¢/e results in a value of 221 K for insulation system surface temperature. This
then becomes the basis, along with the orbital average propellant feed rate in LEO, for selecting the insulation. Figure 8 illustrates the ideal heat leak through high density MLI as a function of surface temperature and thickness. It accounts for edge seams, and conduction through spacers and mounting pins, but does not account for degradation of performance due to handling and installation. Allowing for half the tank heat leak to occur through tank structural and fluid penetrations, and doubling the ideal MLI heat leak to account for handling and installation, = 25 mm of MLI will produce the right boil-off rate in LEO. A tank electrical heater with a capacity of 100 W is used to adjust the boil-off rate to the desired value. At GEO the nominal dissipation of the heater is estimated to be 18 W. The Centaur bonded-on foam has a nominal thickness of 15 ram. For MLI/foam systems the foam thickness must be great enough to prevent ice formation at the MLI/foam interface while on the launch pad. Making the conservative assumption that gaseous conduction dominates on the launch pad, the foam should be at least half as thick as the MLI. The tank is vulnerable to damage by micrometeoroids and debris, and analyses may show that the protection offered by the insulation must be augmented by additional shielding. This can be achieved by incorporating a thin aluminium bumper or extra-heavy scrim rein-
Table 2 Propellant storage and management system mass summary 0.75,
Item
Mass (kg)
Hydrogen tank Insulation system Micrometeroroid/debris bumper Tank supports Thermodynamic vent system Compressor Accumulator Valves and plumbing
263 151 27 26 7 1 4 23
Total
E ~ O.SO. ~ 0.25O 150
1~5
2;0 2~5 2;0 I.SUL~TiORSORF,CETEMPE.~TORE(N)
502 Figure 8
426
~
Cryogenics 1993 Vol 33, No 4
13MMMLf
NOTE: • OOUGLEALUMINIZEO KAPTON • DACRONNETTINGSPACER • 2.6LAYERSPERMM • ~/e = 0,3
Heat flux through MLI
25MM 38MM 50MM
2%
sos
EOTV cryogenic propellant system: J.R. Schuster et al. forcement in the outermost layer of the insulation system.
Tank pressure control Tank pressure control for the EOTV is achieved through mixing the tank contents to counter thermal stratification, and venting tank vapour to provide tank cooling. This vented vapour is the 'boil-off' commonly referred to. The Centaur uses its reaction control system (RCS) thrusters to provide --~ 10-3g to settle the liquid to enable venting. The EOTV will have much lower acceleration ( < 1 0 - 4 g ) and periods of zero-g coast while passing through the Earth's shadow, making it uncertain where the vapour is located in the tank. Additionally, there may be times during the EOTV mission when it will be necessary to coast for much longer periods. Thus, the EOTV should utilize a thermodynamic vent system (TVS). A passive TVS consists of a J o u l e - T h o m s o n ( J - T ) valve and a heat exchanger; an active TVS also includes a tank mixer. The heat exchanger of the passive TVS is vent tubing attached to the tank wall. An active TVS can use either wallmounted tubing or a compact heat exchanger. Figure 9 shows an integrated active TVS developed for the Shuttle/Centaur that was mounted inside the tank and incorporated a compact heat exchanger and mixer. This TVS incorporated a regulator that allowed it to accept liquid, vapour or two-phase fluid, while ensuring that the vent flt/id exiting the TVS was pure vapour. It weighed 17 kg, and was capable of venting over 3 g s ~. The = 0 . 2 g s ] system required for the EOTV could be much smaller. The Shuttle/Centaur TVS was extensively ground tested but never flown.
Liquid dynamics and slosh control The EOTV is expected to roll periodically to realign the solar panels, and this is expected to cause fluid motion and excite propellant slosh modes when coupled to vehicle acceleration. In addition, liquid tends to cover the walls of the tank in zero-g, and will reorient and slosh when the thrust is initiated upon exit from the Earth's shadow. This propellant motion can lead to shifts in vehicle centre-of-mass, affecting attitude and pointing. Tank ring baffles will probably be needed to limit propellant motion.
Propellant supply The tank thermal control and pressure control systems will be designed to achieve the proper boil-off characteristics to meet the overall propellant needs of the propulsion system. However, in order to supply propellant on demand at the temperature, pressure and flow rate needed, three additional elements are required: a compressor, an accumulator and a flow regulator. The compressor raises the pressure of the hydrogen vapour from the TVS to a pressure level adequate to provide for stable, efficient thruster operation. It enables a lightweight, low-pressure hydrogen tank to be used. The pressure ratio for the compressor would be in the range 2 - 4 and would require power of the order of 100 W. This piece of equipment should be quite small and light, and could possibly be a derivative of hardware used for cryogenic sensor cooling.
Figure 9
Shuttle/Centaur thermodynamic vent system
The accumulator is a pressure vessel used to store compressed, gaseous hydrogen boil-off, damp out any supply fluctuations and provide for smooth start-up and shut-down of the propulsion system. Its size will depend on the amount of storage desired, but 0.25 m 3 should be sufficient to store enough hydrogen to operate the thrusters for an hour. This vessel is probably an off-theshelf, flight-qualified bottle. The flow regulator actually regulates pressure and thus regulates flow rate according to the pressure drop characteristic of the propulsion system. The desire to maintain a fixed flow rate could be met with either a mechanical or electronic pressure regulator. A mechanical pressure regulator is used on Centaur to control tank venting. Electronic pressure regulators have been developed for low-temperature vapour service, including hydrogen. Some development might be needed, however, to achieve control for the relatively low flow rates needed for this application, and the regulator would need to be flight-qualified. An alternative approach could be to use a flow network of parallel fixed orifices, with each branch controlled by a shut-off valve.
Quantity and flow gauging Gauging of the quantity of hydrogen in the tank and the flow rate to the propulsion system will be needed on the EOTV to enable spacecraft control and support mission operations. Off-the-shelf gauges of several types are available for accurate measurement of hydrogen vapour flow rate in the range needed. If there is not an appropriate flight-qualified gauge, qualification should be straightforward. Mass gauging the hydrogen in the storage tank might be done with the hydrogen liquid settled while the thruster is operating, using the same capacitance method employed by the Centaur, if the capacitance probe will drain under the imposed acceleration. Alternative laser and microwave sensors have been demonstrated in the laboratory and might be flight-qualified for future application to an EOTV. Some development work has been done in industry on mass gauging techniques for zero-gravity conditions. At this time, however, no zerogravity technique is sufficiently developed to be assured of its success.
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EOTV cryogenic propellant system: J.R. Schuster et al.
Conclusions A low-thrust, solar-electric orbital transfer vehicle is a concept for highly efficient transfer of payloads from LEO to GEO. A high specific impulse and synergistic sharing of subsystems between the E O T V and its payload allows a smaller launch vehicle to be used and/or a heavier payload to be launched than for current LEO-to-GEO propulsion stages. The hydrogen propellant storage and supply system for the E O T V is simple and has modest development requirements because it can use currently available technology such as from Centaur tankage, insulation and pressure control systems. Use of this technology could result in an E O T V
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with a small availability.
inert
mass
fraction
and
short-term
References 1 Dickey, M.R., Klucz, R.S., Ennix, K.A. and Matuszak, L.M.
Developmentof the electric vehicle analyzer, Air Force Astronautics Laboratory Report AL-TR-90-006, Edwards Air Force Base, California, USA (June 1990) 2 O'Neill, R.F. and Zich, J.L. Vector sweep user's guide. General DynamicsReportGDSS-SP-85-011,GeneralDynamicsSpaceSystems Division, San Diego, California, USA (December 15 1985) 3 Dunbar, D.R. Mission planner's guide for the Atlas launch vehicle family, revision 2, General Dynamics Commercial Launch Services, Inc., San Diego, California, USA (July 1990).