Subject: Space-tech Digest #131 Contents: P2-E Booster (2 msgs) ------------------------------------------------------------ Date: Mon, 9 Nov 92 22:22 PST To: space-tech@cs.cmu.edu Subject: P2-E Booster (1/2) From: Bruce_Dunn@mindlink.bc.ca (Bruce Dunn) A NOTE ON THIS POSTING: This posting covers a proposed design for a family of large, pressure fed, liquid fueled boosters which in conjunction with pump-fed upper stages can be used to create a variety of medium and heavy lift launchers. The design is a descendant of the "P2" proposal that I circulated several months ago, and on which I got extensive feedback (thank you everyone!). The major aim of the redesign is to lower development costs and to make a minimum of extrapolations from the current "state of the art" in booster construction. The following changes have been made from the original proposal: 1) The basic vehicle is now totally expendable, although it retains the potential for future recoverability. 2) A family of vehicles with different payload sizes is now proposed, with the minimum vehicle sized to compete with existing commercial launchers. 3) The pressurization system has been changed from hydrogen to helium (I would rather switch than fight). 4) The steel used for the tanks has been changed from 18Ni250 maraging steel to Ladish D6aC steel, as used in the shuttle SRBs (heavier, but cheaper and proven technology) 5) The upper stage has been changed to LOX/kerosene in order to eliminate the toxic hydrazine and nitrogen tetroxide propellants used in the original design. Similarly the hydrazine used to power thrust vector control APUs has been eliminated in favor of hydrogen peroxide, and the use of a hydrazine starting slug for engine ignition has been eliminated. Although still capable of producing a fireball, the vehicle is no longer capable of producing a Bhopal-like toxic cloud in the event of a tank rupture or valve leak. 6) The engine combustion chamber is ablatively cooled, rather than regeneratively cooled. 7) The nozzle of the engine features a larger expansion ratio (12:1) for sea level operation. In the interests of lower development costs, the same nozzle is used for engines igniting at altitude. GENERAL DESIGN OF P2-E BOOSTERS: The P2-E family consistes of large, inexpensive, expendable pressure-fed, liquid-fueled boosters, which can be used as first and second stages of launch systems. The boosters burn propane and hydrogen peroxide (hence the P2 part of the name) and are expendable (hence the E part of the name). An individual pressure fed stage can be produced in sizes from approximately 500 to 1500 or more metric tons mass. All sizes use a single common engine and are the same diameter - the only difference is tank length. Propellants are contained in high strength steel tanks, and are pressurized by room temperature helium gas. Propellants are routed through hydraulically actuated throttling valves, are burnt in an ablatively cooled combustion chamber, and are exhausted through a steerable ablatively cooled nozzle derived from the Shuttle solid rocket booster (SRB). Chamber pressure at full thrust is 6 MPa (approximately 900 psi, the same as the SRB), supplied by propellants pressurized to 7.5 MPa. This gives a sea level thrust of 13,500 kN per booster and a vacuum thrust of 15,500 kN (roughly 3 million pounds force). TANK DESIGN: Separate fuel and oxidizer tanks with no common bulkhead are used. Tanks are welded ultra-high-strength steel, with a diameter of 6 meters. Depending on the stage size, fuel tanks range from approximately 6 to 15 meters length, while oxidizer tanks range from 12 to 35 meters in length. The propane tank is on top of the peroxide tank, and an internal fuel line descends through the oxidizer tank from the propane tank. Tank mass is dominated by the requirements of pressurization. Tank masses were calculated by standard pressure vessel formulas for cylindrical tanks and spherical end caps, with an allowance of an extra 5% mass for local strengthening in critical areas and for attachments etc. Tanks are designed with 5% ullage, and a safety factor of 1.2 (calculated yield stress divided by normal operating stress). Before each use, tanks are tested hydrostatically to 10% over their maximum operating pressure. The P2-E design assumes the use of Ladish D6aC steel, heat treated to a minimum yield strength of 1340 MPa. This gives tank walls ranging from 10 to 20 mm in thickness. This steel has been widely used at this yield strength for the construction of solid rocket motor casings such as those for the Titan and Shuttle. TANK PRESSURIZATION SYSTEM: Propellant tanks are pressurized with room temperature helium gas, generated as needed from liquid helium. Liquid helium is stored in an insulated light-weight non-pressurized tank. The liquid helium is pumped to high pressure using a turbopump powered by the decomposition of hydrogen peroxide. The stream of liquid helium is vaporized and warmed to room temperature in flow-through gas generator which burns pressurized propane and peroxide in a stoichiometric flame, then quenches the flame with liquid helium. Peroxide and propane for the pump and gas generator system are stored in small secondary propellant tanks, and are pressurized by gaseous helium from a high pressure tank. The tank pressurization system and its propellant and helium supply tanks are located in the space between the hemispherical top of the oxidizer tank and the hemispherical bottom of the propane tank. This space is surrounded by a cylindrical fairing which has the same diameter as the fuel and oxidizer tanks. This fairing has a structural and aerodynamic function in joining the fuel and oxidizer tanks, but also acts to form a bay for the tank pressurization equipment, avionics etc. For the smallest booster sizes, the helium volume is low enough that the propane and peroxide tanks can be butted up against one another. For larger sizes of booster, in order to accommodate the volume of liquid helium needed, the inter-tank space is enlarged by spacing the main propellant tanks apart slightly. ENGINE AND NOZZLE: The P2-E boosters use a new pressure fed engine which utilizes the steerable, ablatively cooled nozzle from the Shuttle solid rocket booster (SRB). The nozzle is mated to a newly designed ablatively cooled combustion chamber. The resulting engine should have low development costs as it has no gas generator system, no pumps, no cooling passages (other than in the injector area), burns dense, storable, and relatively non-toxic propellants, and uses an existing nozzle and thrust vector control system. The major component needing development is an injector and combustion chamber to burn hydrogen peroxide and propane. On an SRB, the nozzle is bolted to the bottom of the steel motor casing and has an ablatively cooled throat which protrudes into the central hole in the propellant grain. The casing to nozzle joint is protected by insulation. The proposed P2-E combustion chamber has a length of approximately 2 meters and a diameter of 3 meters, and is made of steel with an interior lining of ablative material similar to that used on the nozzle. The inner surface of the ablative lining is given the same shape and interior dimensions as an unburned propellant grain on a SRB, and the SRB nozzle is bolted to the combustion chamber in the same way that it is bolted to the bottom segment of the SRB. The top of the chamber contains an injector system which is fed propellants via hydraulically actuated throttling valves. The top of the combustion chamber is rigidly bolted to the bottom of the peroxide tank. In estimating booster mass, this combustion chamber has been assigned a mass of 5 metric tons. Together with a nozzle assembly with a mass of 10 tons, the engine masses 15 tons in total. The P2E ablatively cooled nozzle is directly derived from the standard SRB nozzle by increasing the expansion ratio from 7.38 to 12 by extending the nozzle exit cone. The P2-E nozzle has a throat diameter of 1.38 m and an exit cone diameter of 4.78 m. The engine is estimated to have a specific impulse of 250 at sea level and 290 in a vacuum. Isp values are extrapolations based on the actual performance of the Rocketdyne RS-27A and F-1 kerosene/LOX engines. The extrapolation procedure assumes a factor of 0.926 to account for the use of peroxide rather than LOX as an oxidizer, a factor of 1.02 to account for the use of propane rather than RP-1 as a fuel, and a factor of 1.02 to account for the fact that unlike the Rocketdyne engines, the P2-E engine does not divert main propellants to power a turbopump. Thrust vector control is achieved by two hydraulic actuators powered by redundant auxiliary power units (APUs) consuming pressurized hydrogen peroxide from the main propellant supply - this is the same system employed on the SRBs except that the SRB APUs utilize pressurized hydrazine from a separate tank system. Propellant throttling valves take their power from the thrust vector control hydraulic system. Propellant flow and mixture ratio is adjusted on a feedback basis. If necessary, the engines can be throttled partway through the flight to lower aerodynamic forces or to limit final accelerations. For flight regimes where only a single P2-E booster engine is ignited, roll control is provided by mono-propellant thrusters utilizing hydrogen peroxide from the main pressurized peroxide supply. Because the nozzle exit cone diameter (4.78 meters) is substantially less than the booster diameter (6 meters) the vehicle can be directly supported on the pad, with the nozzles protruding through holes in a platform. There is no need for the "skirt" assembly of the SRBs, which protrudes in order to provide pad support points for the booster, which is narrower than the SRB nozzle. In the launch sequence, the helium pump and gas generator system is activated to pressurize the tanks. As the propellants reach full pressure, the engine throttling valves are opened to allow propellant into the combustion chamber, and a pyrotechnic igniter starts the engine. Once full thrust is reached, the vehicle is launched. The launch sequence may be aborted by closing the throttling valves while simultaneously shutting down the helium pressurization system. MASS BUDGET (METRIC TONS): The following dry masses are independent of the size of a P2E stage - all masses are in metric tons. Masses are estimates, subject to revision upon execution of more detailed designs. Engine 15 Pressurization equipment, reactants and plumbing 5 Inter-tank fairing 3 Thrust vector control pumps, actuators 1 Avionics, range safety 1 Separation motors and stage attachments (booster) or upper stage adapter (core) 2 Miscellaneous Structure and Equipment 3 Growth Margin at 10% 3 Fixed mass total 33 Actual stage mass includes the 33 tons of fixed mass, plus propellants, plus tank mass, plus helium mass. The masses involved are listed below for a 600 metric ton and a 1600 metric ton stage. Total mass Propellant Tank Helium Fixed Mass Fraction 600 517.9 44.4 5.7 33 .861 1600 1418.7 132.2 15.5 33 .887 UPPER STAGES: For baseline performance calculations for launchers using P2-E boosters, a hypothetical kerosene/LOX upper stage of either 100 or 200 metric tons mass has been assumed. A mass fraction of 0.925 has been assumed, which is the approximate mass fraction of the core stage of the ex-Soviet Sputnik/Vostok/Soyuz launcher. This conservative design is approximately 30 years old and is man-rated. The engine for hypothetical upper stages would be an ex-Soviet RD-108, with a thrust of approximately 941 kN and a vacuum specific impulse of 315. This engine has been in mass production for over 3 decades and is the sustainer engine of the core of the Soyuz launcher. Its vacuum performance is modest as it was designed to operate at sea level. It does however have a thrust in the right range, is reliable, and if it could be obtained would likely be inexpensive. A single engine would be used for the 100 ton stage, and two engines for the 200 ton stage to maintain the thrust to mass ratio. An alternate to the RD-108 engine would be a Rocketdyne RS-27A which has approximately the same thrust. This engine is also designed for sea level operation, and its 12:1 nozzle generates a vacuum specific impulse of only 301. Since however it has only a single nozzle (unlike the 4 nozzle cluster of the RD-108) it could potentially be easily redesigned for a higher expansion ratio and specific impulse. For all launchers, payload is strongly dependent on the nature of the upper stage. To indicate the possible range of payloads which could be lifted using higher performance upper stages, calculations have also been done with a hypothetical 100 or 200 ton LOX/kerosene upper stage with a mass fraction of 0.95 and a specific impulse of 340. Finally to indicate the maximum potential payload, calculations have been done assuming a 200 ton hydrogen/oxygen upper stage with a mass fraction of 0.90 and a specific impulse of 450. The engine for such an stage would be a SSME (space shuttle main engine), or a STME (space transportation main engine). PERFORMANCE: Performance calculations were made using a spread-sheet which models the acceleration of a payload to a velocity of 9300 m/sec (low earth orbit velocity, taking into account losses during ascent and assuming a launch at 28.5 degrees latitude). For all configurations, total liftoff mass is governed by the sea level thrust of the P2-E engines in operation at takeoff. An initial acceleration of 1.35 g was required of the vehicle. For a given upper stage (100 or 200 metric tons depending on configuration), the spreadsheet "solver" and "goal seek" functions were used to vary the sizes of the pressure fed P2-E booster elements in order to maximize the payload. Payloads given are gross, without any allowance for shrouds - addition of allowances for shrouds carried partway to orbit might lower payloads by something like 10%. xxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxx Configuration 1: Minimum P2-E Core stage: P2-E, 907 metric tons Upper stage: LOX/kerosene, 100 metric tons, Isp=315, MF=.925 Payload to LEO: 12.7 metric tons Core stage acceleration: 1.35 g to 7.0 g, staging at 4100 m/sec Upper stage acceleration: 0.9 g to 4.8 g Payload using alternate upper stages (lower stage re-optimized): 18.6 metric tons when using 100 ton, LOX/kerosene, Isp=340, MF=.95 24.2 metric tons when using 200 ton, LOX/kerosene, Isp=340, MF=.95 36.0 metric tons when using 200 ton, LOX/H2, Isp=450, MF=.90 The core stage and upper stage are stacked vertically, with the payload on top of the upper stage. With a conservative upper stage, this vehicle has roughly the payload capability of an Ariane 44L, and more than the capability of an Atlas or Delta. xxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxx Configuration 2: Boosted P2-E, series burn Booster stages: 2 x P2-E, 640 metric tons each Core stage: P2-E, 525 metric tons, igniting after booster burnout Upper stage: LOX/Kerosene, 200 metric tons, Isp=315, MF=.925 Payload to LEO: 36.8 metric tons Booster stage acceleration: 1.35 g to 3.4 g, staging at 2100 m/sec Core stage acceleration: 2.0 g to 4.9 g, staging at 4700 m/sec Upper stage acceleration: 0.8 g to 3.7 g Payload using alternate upper stages (lower stages re-optimized): 49.2 metric tons when using 200 ton, LOX/kerosene, Isp=340, MF=.95 67.7 metric tons when using 200 ton, LOX/H2, Isp=450, MF=.90 The core stage, the upper stage, and the payload are stacked vertically, as for configuration 1. The core stage is flanked by two boosters, which ignite at takeoff and accelerate to booster staging. Booster separation is performed by solid motors, using motors from the Shuttle SRB, which has approximately the same dry mass as the P2-E boosters. The core stage then ignites. Using a conservative upper stage, this vehicle has about 75% more capacity than the Titan IV or Shuttle. xxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxx Configuration 3: Boosted P2-E, parallel burn Booster stages: 2x P2-E, 596 metric tons each Core stage: P2-E, 1619 metric tons, igniting at takeoff Upper stage: LOX/Kerosene, 200 metric tons, Isp=315, MF=.925 Payload to LEO: 51.7 metric tons Booster stage acceleration: 1.35 g to 3.13 g, staging at 2100 m/sec Core stage acceleration: 1.2 g to 3.6 g, staging at 5200 m/sec Upper stage acceleration: 0.8 g to 3.0 g final Payload using alternate upper stages (lower stages re-optimized): 62.6 metric tons when using 200 ton, LOX/kerosene, Isp=340, MF=.95 84.9 metric tons when using 200 ton, LOX/H2, Isp=450, MF=.90 The core stage, the upper stage, and the payload are stacked vertically, as for configuration 1. The core stage is flanked by two boosters. At takeoff, both of the boosters and the core stage ignite. The use of all three engines at launch allows a substantial increase in the total vehicle liftoff mass. No cross-feeding of propellants is used during the parallel burn. This vehicle makes maximum use of the P2-E stages and a 200 metric ton kerosene/LOX upper stage to launch a very heavy payload. xxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxxx These performance estimates indicate that heavy payloads can be orbited with an extremely unsophisticated, albeit very large vehicle. The performance of the P2-E vehicle is not very sensitive to the assumptions made. Dry mass and specific impulse are more critical for the P2-E minimum vehicle, which has only two stages, than they are for the three stage boosted configurations. The sensitivity of the 2 stage vehicle is as follows: Change from Baseline Design Percent of Original Payload 10% increase in core burnout mass 93% 10% increase in upper stage burnout mass 94% decrease of 10 in core specific impulse 92% decrease of 10 in upper stage specific impulse 91% all of above simultaneously 70% These figures may be contrasted with those of SSTO vehicles, where in many cases a small increase in dry mass or a loss of specific impulse can completely eliminate the payload. COMPARISON WITH EXISTING LAUNCHERS: The P2-E minimum launcher may be compared with the following competitors: Vehicle Cost, $M LEO payload Vehicle Mass metric tons metric tons P2-E with 10% shroud allowance ??? 11.4 1000 Ariane 44L/LH2-LO2 110 9.6 469 Atlas2/Centaur 80 7.1 163 Delta2/PAM-D 52 5.0 204 It is premature to estimates costs for the P2-E vehicle. However, as a very rough rule, costs are more associated with complexity and the number of parts in a vehicle, rather than sheer size. Complexity drives up inspection and quality control costs and gets in the way of reliability - this is one of the basic problems of the Shuttle. On the basis of complexity, the P2-E compares very well with the commercially successful Ariane 44L vehicle, which has slightly less payload capability to low earth orbit. Vehicle Tank sets Separation Planes Engines Ariane 44L 7 6 10 P2-E 2 1 2 The P2-E vehicle avoids the use of solid rocket boosters, eliminating the pollution problems associated with their exhaust (Titan, Delta, Shuttle, some Ariane configurations). It avoids the use of toxic propellants such as nitrogen tetroxide and hydrazine derivatives (Titan, Ariane), and avoids the use of a hydrogen/oxygen stage (Titan IV, Atlas, Ariane, Shuttle). -- Bruce Dunn Vancouver, Canada Bruce_Dunn@mindlink.bc.ca ------------------------------ Date: Mon, 9 Nov 92 22:30 PST To: space-tech@cs.cmu.edu Subject: P2-E Booster (2/2) From: Bruce_Dunn@mindlink.bc.ca (Bruce Dunn) CONTRUCTION AND OPERATIONAL ASPECTS: The P2-E boosters are very simple. The tanks are heavy walled and welded, and need neither the aircraft-style light weight construction of conventional boosters, or the carefully shaped, elaborately jointed heavy segments needed to construct a large solid rocket. It would probably be easiest to weld together the tank sets near the launch pad, using shaped steel segments shipped from a primary supplier. The different P2-E boosters vary only in the length of the barrel shaped section of the propane and peroxide tanks, therefore a stage of any desired size can be assembled on demand. A lower limit of 525 metric tons stage mass is set by the fact that at this size and a diameter of 6 meters, the propane tank becomes a sphere. Tanks are hydro tested before any flight to ensure their integrity. The development of the P2-E family would start by building the minimum 2 stage launcher, which even conservative calculations suggest has the capability to lift most current payloads with considerable spare capacity. The main items requiring development are the combustion chamber and the helium gas generator system. Engine testing should be easy, as no turbo-pumps are involved. The gas generator system would be fully testable without actual engine operation. These two key elements, combined with high strength steel tanks, would be the basis of the pressure fed P2-E stages. At the same time, an upper stage would be designed built using an existing engine. Once the basic P2-E vehicle was operational, it would go through an initial flight test program designed to demonstrate its performance, and establish a preliminary reliability record. Once several successful flights have been achieved, it would be able to compete with existing launchers. Until a more extensive reliability record has been built up, the launcher would probably have to compete on the basis of superior price or payload. Because of its inherent simplicity however, it is likely that in the long term the launcher will be very reliable, and become attractive on that basis as well. The minimum P2-E would compete most directly with the Ariane IV family of launchers, although if costs were low enough it could also compete with the Atlas II and Delta II. With the establishment of ongoing production of P2-E pressure-fed booster and pump-fed upper stages for the minimum launcher, then the boosted P2-E variants would become feasible with little development cost. This would provide a clear path to the development of heavy launchers using components from a medium launcher. Such boosted vehicles would compete with the projected National Launch System, or with commercial ventures such as heavy lift Delta. ENHANCEMENTS: The calculations used here assume either a 100 or 200 ton upper stage without optimization of the exact stage size. In the real world, ascent trajectory modeling would be used to also optimize the upper stage mass. The optimum upper stage mass will be determined by a trade off between the benefits of increasing the size of the high performance upper stage relative to that of the lower stages, and the penalties of a decreasing thrust to mass ratio when using larger tanks with the same engine. Optimization of upper stage mass would increase performance from that calculated here for arbitrary upper stages. The calculations assume the use of room temperature helium for pressurization. Slightly higher performance and considerably lower helium costs might be achieved by using hot helium. The limitation on helium temperature will probably involve the avoidance of temperature-induced decomposition of the hydrogen peroxide. The booster stages of the series and parallel burn versions of the boosted P2-E have roughly the same dry mass and terminal velocity as the burned-out Shuttle solid rocket boosters, and thus could be recovered by SRB-derived parachutes and reused in order to lower launch costs. If tanks valves are closed after propellants are used, helium used for pressurization could be recovered at the same time. Recovery of the core stage is also potentially possible, but would have to involve a maneuver in which the stage is rotated and reignited after upper stage separation in order to reduce velocity prior to parachute deployment. Total vehicle size is limited by the thrust of the first stage engines, which is in turn limited by chamber pressure. A logical extension of the P2 booster capability would be to employ higher chamber pressures and larger tanks. Alternately, if the full vehicle capability is not needed, propellant could be off-loaded from the vehicle and tank pressures reduced to give a greater margin of safety. The core stage of the series-burn boosted P2-E is ignited at high altitude after the boosters have burned out. A modest improvement in performance could be obtained by reworking the engine nozzle to give a higher expansion ratio for this stage. The P2-E design assumes the use of D6aC steel, chosen because it is already extensively used for solid rocket motor casings. A number of alternate steels are available which could also be used - some of these have higher strengths while retaining adequate toughness, and their use could provide thinner tank walls and a modest increases in performance. For extremely heavy payloads, four booster stages could be used, together with a core stage and an upper stage. A series burn vehicle using a conservative kerosene/LOX upper stage can boost 74 tons to low earth orbit, while a parallel burn vehicle can boost 93 tons. DESIGN ISSUES: The standard pump-fed rocket propellant combination of RP-1 and LOX, although having a higher Isp than propane/peroxide, is not easy to employ in pressure fed booster designs. The lower bulk density of the RP-1/LOX combination leads to larger tanks for a given mass of propellant. The LOX tank must be made of aluminum, as normal steels are too brittle at cryogenic temperatures and cryogenic compatible stainless steels do not have the tensile strengths required for lightweight tanks. The walls of an aluminum tank for the sizes of boosters used would be approximately 60 mm thick, leading to fabrication difficulties. The fuel and oxidizer tanks are at markedly different temperatures, leading to mechanical difficulties at attachment points due to differential expansion of materials, and requiring insulation where the fuel line passes through the LOX tank. Finally, cooling of pressurization gas by the cryogenic propellant increases the amount of helium needed for pressurization. For a given size of vehicle, whose mass is limited by the take-off thrust available, the performance of pressure-fed RP-1/LOX appears to be no better than that of propane/peroxide, and the design appears to be much harder to execute. The first and second stages burn propane/peroxide at a mixture ratio of 7.5 to 1. Propane is used rather than RP-1 as it is less expensive, gives about 2% higher Isp than RP-1, has a much lower viscosity (lowering pressure drops in fuel lines and injectors) and has an intrinsic vapor pressure which aids tank pressurization. Its only disadvantage is its lower density (0.5 vs. 0.8 for RP-1). In practice, the higher density of RP-1 is offset by its lower Isp and lower mixture ratio (7.0, vs. 7.5 for propane), and for a given vehicle mass propane/peroxide and RP-1/peroxide have virtually identical performance. Propane wins because of its other virtues. Hydrazine derivatives are rejected as fuels because of their toxicity and the fact that they are considerably more expensive than hydrocarbons. Hydrogen peroxide is used as an oxidizer as it gives good performance, is not cryogenic, is dense, and gives propellant combinations with a high oxidizer to fuel mixture ratio and thus high bulk densities (important for pressure fed rockets). N2O4 is rejected as an oxidizer because of its toxicity. Nitric acid is rejected as an oxidizer because of low performance, and because in its common form of IRFNA (Inhibited Red Fuming Nitric Acid) it contains a high proportion of releasable toxic NO2. Peroxide can be washed from tanks with water to allow internal inspections, and any accidental spills in the pad area can be flushed away with water (3% hydrogen peroxide in water is benign enough to be sold in drug stores for cleaning open wounds). Hydrogen peroxide is relatively non-toxic - spilled peroxide does not give off fumes, and skin contact results in a bleached patch of skin without permanent damage. 100% hydrogen peroxide is a relatively stable material, with decomposition rates of about 1% per year. Although hydrogen peroxide decomposes energetically to oxygen and water, it is impossible to get bulk hydrogen peroxide to detonate unless a strong booster charge is used. Runaway decomposition can only happen at elevated temperatures or with grossly impure material. Anecdotal reports of difficulties with peroxide instability probably date from wartime German experience with highly impure peroxide in improperly prepared tanks. Peroxide decomposition is a surface catalyzed phenomenon, and is less troublesome when volumes are large. In practice, aside from slow decomposition, hydrogen peroxide is probably as least as stable as the various hydrazine fuels which are already in use in the aerospace industry. 50 to 90% peroxide solutions are readily obtainable commercial chemicals, and refining peroxide to 100% can be done distillation or freezing methods. For use in a launching system, peroxide could be bought and shipped to the launch site at moderate strengths, then refined there to 100%. The propane/peroxide combination gives good performance, is dense, and uses relatively non-toxic propellants. While hydrocarbon/peroxide engines have not been widely used in space launchers to date, there is a considerable information on combustion with this combination. Considerable development work was done in the 1950s on using peroxide and jet fuel for "super-performance" rockets for short term bursts of thrust in military aircraft. A useful reference is "Advantages of Hydrogen Peroxide as a Rocket Oxidant" by David Andrews (J. Brit. Interplanetary Society, 43:319-328, 1990). The helium used to pressurize propellants is stored as a liquid, and vaporized as needed. The amount of helium needed is impractical to store as a room temperature pressurized gas because of the required tank weight. Using gaseous helium would approximately double the vehicle tank mass. When stored as a liquid at its normal boiling point, the approximately 9 tons of helium needed by a P2-E minimum booster requires approximately 70 cubic meters of tank space. Since booster liquid hydrogen tanks are easily produced which weigh 15% of their content, and liquid helium is over 50% denser than liquid hydrogen, it seems reasonable that a first estimate for the mass of the needed helium tank might be 10% of the helium mass, or less than 1 metric ton. This tank would be contained within the inter-tank area, and insulated to provide a holding time of many hours so that the helium could be pumped into the tank well before launch. It is not easy to avoid the liquid helium pumping system. Storing liquid helium in a cryogenic pressure vessel is possible, but such a vessel would be heavy and pressurization would present difficulties. The P2-E design therefore relies on pressurizing the helium with a turbopump similar to the hydrogen pump for a liquid hydrogen/liquid oxygen rocket engine. The turbopump must pump about 0.5 cubic meters/sec of liquid helium - this may be compared to the 0.4 cubic meters/sec output of the hydrogen pump of the J2 oxygen/hydrogen engine used in the upper stage of the Saturn. The pump end of the system is thus technology which has been mastered 3 decades ago. The J2 pump turbine consumed approximately 3 kg/sec of non-stoichiometric propellant in order to provide the power to pump both fuel and oxidizer. It seems reasonable to estimate the peroxide composition of the P2-E pump at less than 3 kg/sec, or under 500 kg for a typical vehicle. The P2-E design relies on a gas generator to vaporize and warm the helium used for tank pressurization. The gas generator requires approximately 200 kg of propane+peroxide for every ton of helium vaporized - most energy is used for warming, not vaporization. Engine heat from the combustion chamber could also be used to evaporate and warm the helium. Adding this requirement to the engine however complicates the design of the combustion chamber and involves considerable piping to connect the combustion chamber to a source of liquid helium and to conduct the vaporized helium back to the tops of the propellant tanks. Furthermore, there may be problems in getting enough heat from the combustion chamber to vaporize and heat the needed volume of pressurizing fluid (the chamber walls, throat and nozzle are not available as a heat source as they are ablatively cooled). If the helium vaporization system were integrated with the engine, it could not be tested independently of full scale engine tests. Finally, pressurization gas would not be available until the engine is running, complicating engine startup. Nitrogen could be used as a less expensive alternative to helium as an inert pressurization gas, but would reduce vehicle payloads by approximately 25% because of its high mass relative to helium. Other potential pressurization gases are either even heavier that nitrogen, or are potentially reactive with either the fuel or oxidizer. For the P2 booster, the expansion ratio of the SRB nozzle will be increased by extending the exit cone. In the current calculations, an area ratio of 12:1 is assumed, although the actual ratio used would obviously rely on more sophisticated modeling of vehicle performance. The P2-E engine represents a way to generate an extremely high thrust engine with a minimum of development costs. The P2-E engine has essentially double the thrust of the LOX/kerosene F1 engine used in the Apollo program or the ex-Soviet RD-170 engine. The chamber pressure of the P2-E engine is about 80% of the F1 pressure, while the expansion ratio is 12:1 compared with 16:1 for the F1. The sea level thrust/mass ratio of the 15 metric ton P2-E engine is 800 N/kg, similar to that of the F1 engine. In comparing the two engines, it may be noted that the F1 engine mass includes extensive turbomachinery not present on the P2 booster engine, while the P2 engine has an ablatively cooled nozzle which would be heavier than its regeneratively cooled equivalent on the F1. The ablative material used in the SRB nozzle must withstand solid propellant exhaust gases at 3500 K and the same pressures used in the P2-E. A number of ablatively cooled engines (including the lunar lander engine) have been developed which burn hydrazine derivatives and N2O4 at approximately 3400 K. Propane and peroxide have a combusion temperature of only 3000 K, and except for the absence of nitrogen the exhaust products are similar to those of hydrazine + N204. It seems unlikely therefore that there will be any major problems in adapting the SRB-derived ablative engine materials to the P2-E, or in substituting other materials which would be resistant to the peroxide+propane exhaust. CALCULATION ISSUES: The spread sheet used to determine payloads to LEO does not actually "fly" an ascent, and gravity and air resistance losses are only estimated. Nevertheless, the spreadsheet gives reasonably accurate results with test cases. The spread sheet estimates a gross payload to low earth orbit of 7.8 metric tons for the Soyuz launch vehicle. When carrying a shroud and escape tower for the first 160 seconds of flight, the literature value for the nominal payload for this vehicle is 6.8 metric tons. For the Ariane 44L launcher, the spreadsheet predicts a gross payload of 10.9 metric tons, against a literature value of 9.6 metric tons net. The calculated performance of a vehicle depends on only a few key numbers. For the minimum P2-E launcher, these are: Sea Level Isp of Core Stage estimate 250 Vacuum Isp of Core Stage estimate 290 Mass Fraction of Core Stage estimate 0.875 Vacuum Isp of Upper Stage estimate 315 Mass Fraction of Upper Stage estimate 0.925 I am confident that the upper stage numbers are achievable and are in fact probably quite conservative, as they represent numbers taken from a real-world vehicle designed several decades ago. I feel that the mass fraction of the core stage is a reasonable estimate and is well within current technology - the estimate is based on materials, construction practices and component masses already in use for the shuttle SRB. I am somewhat less confident about the sea level and vacuum Isp of the core stage - I would really like to get someone with experience to calculate or estimate these. This is not simply a problem of calculating the theoretical values of the Isp for propane and peroxide - the main problem is what fraction of the theoretical Isp is actually realized in practice. Any and all comments are welcome. I found the feedback on the original P2 design very perceptive and most helpful in pinpointing issues of concern. -- Bruce Dunn Vancouver, Canada Bruce_Dunn@mindlink.bc.ca ------------------------------ End of Space-tech Digest #131 *******************