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APPENDIX D Possible Impacts of Effluents from SDI Systems SPACE SHUTTLE EXPERIENCE RELEVANT TO POSSIBLE IMPACTS OF EFFLUENTS PROJECTED FOR SDI SYSTEMS Space Shuttle experience is relevant to projecting likely effects of effluent dump rates for Strategic Defense Initiative (SDI) systems. A good source of data on effluent rates for the Space Shuttle was provided by Pickett et al. (19853. Sources of effluent on the Shuttle include outgasing (estimated at about 3 x 10-4 kg/s at mid-mission), Flash Evaporator System (FES) operations, water dumps, and the primary and vernier Reaction Control System (RCS) engines. Data found for the STS-3 mission are presumed here to be typical of those for other Space Shuttle missions. The FES system dumps water vapor in pulses of duration 200 ~ 30 ms at a maximum pulse rate of 4 Hz, yielding a release rate of 22.7 kg/in (= 6.3 x 10-3 kg/s). The average FES release rate is 2.3 kg/in (= 6.3 x 10-4 kg/s) at 0.4 Hz. There were 20 FES releases during the STS-3 mission, with durations ranging from 1 min to about 2.5 h. The Space Shuttle also dumps liquid water at an average rate of 64 kg/in (= 1.8 x 10-2 kg/s). During the STS-3 mission, there were nine water dumps, each of duration 45-60 min. which released 121
122 APPENDIX D a total of 41-93 kg of water. If the 93 kg is assumed to have been dumped in 60 min. this gives a maximum rate of 2.6 x 10-2 kg/s. The RCS consists of 38 primary thrusters (395 kg of thrust each) and 6 vernier thrusters (11 kg of thrust each). These thrusters use monomethyl hydrazine (MMH) fuel with an NO3 oxidizer, and have (calculated) effluent as shown in Table D-1 (Pickett et al., 1985, Table 2~. MMH is N2H3CH3. Minimum pulse duration is 80 ms. Longest pulse on STS-3 was about 30 s. Mass efflux rate for the vernier thrusters is 4 x 10-2 kg/s per engine, and 1.4 kg/s per engine for the primary thrusters. The velocity of the released gases is estimated as 3.5 km/s. Finding estimates for SDI system effluents proved somewhat more difficult. Review of the Space Power Architecture System (SPAS) summary reports (1988) yielded a quotation from TRW, which performed one of the three SPAS studies, of 7.5 kg/s of H2 for cooling a 180-MW free-electron laser (FEL), and 13.5 kg/s of H2 for cooling a 400-MW neutral-particle beam (NPB). Martin-Marietta gives estimates in the 4~100 kg/s range for various NPB systems (Tables D-2 and D-3), and General Electric (GE) quotes efflux rates in the 20 to 40 kg/s range, though these are associated with power systems, so this may amount to comparing apples with oranges (Ta- ble D-43. The wide Variation in these numbers is due to different assumptions about power levels and system designs among the con- tractors. It appears that, with the exception of the primary RCS thrusters, none of the efflux rates associated with the Shuttle approach those estimated for SDI systems. Neither is there much H2 efflux from the Shuttle. The L2U burn (described in the following section), during which the three PROS engines were fired over a period of 1.5 min. was the longest PROS operation identified by the committee. Data from this burn have been used in the following section to estimate the usefulness of the spherical assumption employed for efflux expansion. ESTIMATION O1? THE IMPACT OF EFFLUENT ON PROPAGATION OF A NEUTRAI~PARTIC[E BEAM Following is a calculationusing a spherical approximation to es- timate the impact of effluent on propagation of a neutral-particle beam. Assume that the effluent originates at the origin and expands radially with velocity ur. The resulting mass density p at distance R · . IS given By
APPENDIX D TABLE D-1 Thruster Plume Characteristics for Primary (PROS) and Vernier (VRCS) Thrusters Effluent Species Molecular Weight Mole Fraction Composition, Neutrals H 2 O 18 0.328 N 2 28 0.306 CO2 44 0.036 O 2 32 0.0004 CO 28 0.134 H2 2 0.17 H 1 0.015 MMH-NO3 108 0.002 Total 0.9914 Composition, Dominant Ions NO 30 1.7 x 10 8 CO2 44 2.7x 10 10 OH 17 4.3 x 10 10 Electrons -- 24 x 10 9 Number of Number of Ions Thruster Neutrals (electrons) Firing Ejected Ejected VRCS Typical b Longest PROS Typical d Longest 1.3 x 1025 1.7 x 1026 9.2 x 1024 5.5 x 1025 3.1 x 1018 3.8 x 10 2.1 x 10187 1.2 x 10 NOTE: For the primary thruster, m = 1~419.8 g/s/engine, where m is tile mass efflux rate; for the vernier thruster, m = 40.8 g/s/engine. PRCS = primary reaction control system; VRCS = vernier reaction control system. MMH = monomethylhydrazine. b Based on 2 firings ejecting 163 g over 2 s. c Based on 14 firings ejecting 2,100 g over 30 a. d Based on 1 firing ejecting 114 g over 80 me. Based on 5 firings ejecting 682 g over 720 me. SOURCE: Pickett et al. (1985~: Table 2. 123
124 APPENDIX D TABLE D-2 Assumed Effluent Compositions Power System Effluent Compositions in Weight Percent Nuclear Brayton turboalternator Nuclear Rankine turboalternator Liquid metal reactor Combustion Brayton turboalternator Combustion Brayton turboalternator (water collected) Combustion-dri~ren MHD Fuel cell Fuel cell, radiator cooled Fuel cell, water collected 100% H No effluent No effluent FEL: 52.8% H2, 47.2% H2O EML/NPB: 58.6% H2, 41.458 H2O 100% H2 FEL: 20.2% H2, 66.6% H2O, 4.2% CsOHa EML/NPB: 36.156 H2, 60.1% H2O, 3.8% CsOHa FEL: 59.1% H2, 40.9% H2O EML/NPB: 67.6% H2, 32.4% H2O No effluent 10056 H2 NOTE: EML = electromagnetic launcher; FEL = free-electron laser; MHD = magnetohydrodynamics; NPB = neutral-particle beam. aActual cesium-containing species may differ from this. SOURCE: Martin-Marietta Space Power Architecture System (1988) report, Task 3: Table III-2.1. p(R)= 47rR2~t, (1) where m is the mass flow rate. In propagating from Ro to a target at Rf, the neutral-particIe beam traverses a mass per square centimeter given by < pR f pdR= m ~ ~ ~ ), which, for t ~ Ro/u and Rf ~ Ro, becomes m < pR >= (2) For hydrogen exhaust at 1000°K, u # 4.5 x103 m/s. The cross section for 100 MeV Ho on Ho is 1.23 x 10-~9 cm2 in most recent work (Johnstone, 1988~. The corresponding integrated mass density of effluent required to ionize 50 percent of the neutral-particle beam is S.9 x 10-6 g/cm2. Approx~nately 7 percent of the beam is stripped
APPENDIX D TABLE D-3 Assumed Effluent Initial Temperatures (OK), Initial Pressures (atm), and Flow Rates (kg/s) 125 FEL NPB Temp. Flow Pressure Temp. Flow EML Temp. Flow Nuclear Brayton turboalternator 1.0 586 52 416 44 855 44 Nuclear Rankine turboalternator -- -- -- -- -- - - - - Liquid metal reactor -- -- -- -- -- -- -- Combustion Brayton 1.0 570 89 407 70 818 70 Combustion Brayton, water collected 1.0 639 48 447 41 996 41 Combustion-driven MHD 1.5 3000 178 3000 122 3000 122 Fuel cell 1.0 450 81.7 450 61.8 450 61 8 Fuel cell, radiator cooled -- -- -- -_ _- __ __ Fuel cell, water collected 1.0 450 48.3 450 41.8 450 41.8 NOTE: EML = electromagnetic launcher; FEL = free-electron laser; MHD = magnetohydrodynamics; NPB = neutral-particle beam. SOURCE: Martin-Marietta Space Power Architecture System (1988) report, Task 3: Table III-2.2. at 1 x 10-6 g/cm2. For purposes of the present calculation, the integrated mass density for which beam stripping becomes a concern is taken to be < pR >= 10-6g/cm2 = 10~5kg/m2. Ensuring undisturbed beam propagation requires m << 4?ruRO < pR >= 0.56Rokg/s, or, for Ro = 10 m, m << 5.6 kg/s. As noted above, estimates of coolant mass efflux rates for neutral-particle beams used in the Space Power Architecture System (SPAS) studies were in the range of about 10 to 100 kg/s, somewhat greater than the 5.6 kg/s estimated above. Thus, if the spherical approximation is at all reasonable, neutral- particle beam stripping should be a serious concern. A zero-order check on the spherical approximation can be made based on Space Shuttle data from the STS-3 flight. During this flight, an engine test of the ShuttIe's Primary Reaction Control System (PRCS) was conducted. Observations of pressure in the bay
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APPENDIX D 127 were made by the neutral pressure gauge on the Plasma Diagnostics Package (PDP) during this engine test, known as the L2U burn. This burn lasted for about 1.5 min. and involved pulsed firing of the L2U and R1U thrusters and continuous firing of the F2U thruster. Pressure at the PDP's location was about 3 x 10-4 tort (com- pared to a background pressure of 10-7 torr) during the L2U burn (Murphy et al., 1983; Pickett et al., 1985; Shawhan and Murphy, 1983~. The mass flow rate of each of the PROS engines is m = 1.42 kg/s (Murphy et al., 1983, Pickett et al., 1985; Table D-1), which is high enough that the exhaust gas is collisional when it exits, as ex- pected for coolant efflux associated with SD! systems. The L2U and R2U thrusters are located to the port and starboard, respectively, of the ShuttIe's vertical stabilizer on the orbital maneuvering system (OMS) pods and fire upward, while the F2U engine IS on the upper surface of the ShuttIe's nose and also fires upward (Murphy et al., 1983~. During the L2U burn, the engines were fired to obtain net zero torque on the Shuttle, with R1U c~celling the rod induced by L2U and F2U cancelling pitch from both rear thrusters. The PDP is located on-axis, somewhat aft of the center of the payload bay, at about the level of the longerons. Estimates of the distances from the PDP to the three engines are as follows: F2U: 17 m forward, on-axis, 1.5 m down; R1U: 12 m aft, 2.7 m to starboard, 1.5 m up; and L2U: 12 m aft, 2.7 m to port, 1.5 m up. These estimates yield distances of RF ~ 17 m to F2U and RR ~ 13 m to R1U and L2U. The mass density at Ro, the location of the pressure gauge, is computed using equation (1) as the sum of contributions from the three engines. The result for mass density is: p(Ro) = 4 (R2 ~ n~ J To relate this result to pressure, recall that p3pv, (3) (4) where v = >/~ and T is the temperature of the gas at the point in question. An upper limit on the pressure can be obtained by assuming that v = u; that is, there is no cooling of the effluent. Malting this assumption, and using u _ 3.5 km/s (Pickett et al., 1985) yields, for the Shuttle case above,
128 APPENDIX D pma,:(Ro) # 2.1 nt/m2 # 1.6 x 10-2toIT. A reasonable lower bound on the expected pressure may be obtained by assuming that the effluent thermalizes; that is, T # 200°K-300°K at the pressure gauge. The velocity u = 3.5 km/s corresponds to an average kinetic temperature of 22000°K for the PROS exhaust (Murphy et al., 1983; Pickett et al., 19853. Using this correspondence, and taking T = 250°K for the thermalized exhaust, yields Ptherma~ # 1.8 X 10 torn The measured pressure for the L2U burn was 3 x 10-4 torr, which is very close to the pressure computed above for the thermaTized case. This suggests that the spherical approximation is reasonable, which in turn suggests that neutral particle beam stripping by effluent is an issue that must be addressed. The committee considers that a flight experiment to characterize the behavior of the H2 effluent would be appropriate. Altitude is an important consideration here. At Shuttle altitudes, orbital vehicles move at about 7.5 km/s, and the residual atmosphere is dense enough to cause rapid dissipation of effluent clouds. At higher altitudes, the residual atmosphere becomes very tenuous, so the effluent will remain near the vehicle longer. To the extent that the orbits of interest are in the 1,000+ km range, useful early data might be obtained from a sounding rocket experiment, because sounding rocket velocities relative to the background residual atmosphere are on the same order as the random thermal velocity of the background atoms (about 1 km/s). Even though the residual atmospheric densities are high at typical sounding rocket altitudes of a few hundred kilometers, the low relative velocity of the rocket should make the dissipation of rocket- generated effluent comparable to that expected of vehicles orbiting at higher altitude (greater than 1,000 km).
