Subject: Space-tech Digest #110 Contents: P2 Launcher (long) (1 msg) ------------------------------------------------------------ Date: Sun, 29 Mar 92 16:54 PST To: space-tech@cs.cmu.edu Subject: P2 Launcher (long) From: Bruce_Dunn@mindlink.bc.ca (Bruce Dunn) I have recently rejoined the space-tech mailing list after having been absent for a year or so. I read with interest George Herbert's proposals regarding a big dumb hybrid booster. I too have been working on a large launcher concept. The following rather long post describes the current state of the concept. I would appreciate any and all comments, particularly if they address the detailed technical aspects of the proposal. I have tried to do my homework before posting - if I am wrong about technical issues, it is because of lack of adequate sources of information rather than because of wishful thinking. GENERAL DESIGN OF P2 LAUNCH VEHICLE Stage 1 consists of two parallel burning pressure fed liquid fueled boosters with a total mass of approximately 600 metric tons each, burning propane and hydrogen peroxide. These boosters are attached to either side of the second stage of the vehicle. Propellant is contained in high strength steel tanks, and is pressurized to 7.5 MPa by hydrogen gas. Propellants are routed through hydraulically actuated throttling valves, are burnt in a regeneratively cooled combustion chamber at a pressure of 6 MPa (about 900 psi). Gases are exhausted through an ablatively-cooled steerable nozzle derived from the space shuttle Solid Rocket Booster (SRB). Take-off thrust is approximately 12,000 kN (approximately 2,700,000 pounds force) for each of the two first stage engines, while vacuum thrust is approximately 15,000 kN. Total take-off mass is approximately 2000 metric tons, while initial take-off acceleration is 1.2 G At the end of the first stage burn, the two first stage boosters are separated from the second stage by separation motors adapted from the SRB. The separated boosters are recovered by parachute, using the same parachutes as are used for SRB recovery. Stage 2 is essentially the same as the two first stage boosters, but uses a slightly modified engine with a lower thrust and a larger expansion ratio. This stage is expended. Each individual first stage booster is used and recovered two times, then is converted into a second stage and is expended. This semi-reusable system achieves multiple re-use of hardware, while minimizing the need to design for long component lifetimes. Stage 3 has a mass of 200 tons, and utilizes a propulsion system adapted from the second stage of the Ariane 4 launcher, coupled with a large designed-to-cost lightweight tank. Its single pump fed restartable Viking IV engine produces a thrust of 768 kN, and burns UDMH and nitrogen tetroxide. Like the second stage, this stage is expended. This stage can be used to boost payloads of approximately 50 metric tons (110,000 lbs mass) into low earth orbit, or inject smaller payloads into GTO or interplanetary trajectories. PROPELLANT SELECTION The first and second stages burn propane/peroxide at a mixture ratio of 7.5 to 1. The launcher is named the "P2" in recognition of this propellant combination. Propane is used rather than RP-1 as it is less expensive, gives a higher Isp than RP-1, and has a lower viscosity (lowering pressure drops in fuel lines and injectors). Its only disadvantage is its lower density (0.5 vs. 0.8 for RP-1). In practice, the Isp advantage of propane outweighs the density advantage of RP-1, and in launch simulations the propane/peroxide combination slightly outperformed the RP-1/peroxide combination (approximately 2% higher payload). Hydrogen peroxide is used as an oxidizer as it gives good performance, is not cryogenic, is relatively inexpensive, and gives propellant combinations with a very high bulk density (important for pressure fed rockets). 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 and runaway decomposition can only happen at elevated temperatures. In practice, aside from slow decomposition, hydrogen peroxide is probably as stable as the various hydrazine fuels which are already in use in the aerospace field. 90% peroxide is a readily obtainable commercial chemical, and refining peroxide to 100% is a relatively simple process involving distillation. 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. The propane/peroxide combination gives good performance, is dense, and uses inexpensive and relatively non-toxic propellants. While hydrocarbon/peroxide engines have not been widely used in space launchers to date, there is a reasonable amount of 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. UDMH and N2O4 could be used as an alternate non-cryogenic propellant combination, and in simulations yielded essentially the same performance as propane/peroxide (having a marginally higher Isp and a marginally lower density, and yielding a payload about 1 % higher). Both the UDMH and N2O4 components however are very toxic in the event of an accidental spill or leak, and the hydrazine fuel is relatively expensive. Hydrogen pressurization of N2O4 may not be feasible, due the high vapor pressure of N2O4 and the possibility of the formation of explosive mixtures in the tank ullage space (see below for details on tank pressurization). In spite of these disadvantages, the UDMH/N2O4 combination might in the end be preferable to propane/peroxide, due to the large amount of accumulated experience with this combination. The choice of propellants does not strongly influence the design or performance of the P2 launcher. The standard pump-fed rocket propellant combination of RP-1 and LOX, although having a higher Isp than propane/peroxide, is not as easy to employ in pressure fed booster designs. The low bulk density of the RP-1/LOX combination leads to large and heavy tanks. The cryogenic oxidizer is difficult to pressurize, and is not very compatible with high strength steel tanks due to steel embrittlement at low temperatures. Performance calculations were done assuming helium pressurization, and either RP-1/LOX or propane/LOX. To give the benefit of the doubt to the LOX based combinations, it was assumed that a LOX tank would weigh no more than an identically sized maraging steel tank holding a room-temperature oxidizer. Relative to propane/peroxide, calculated payloads were 6% higher for RP-1/LOX and 3.6% higher for propane/LOX. This performance advantage may well disappear once realistic masses for LOX tanks are used. Each flight uses approximately 1400 metric tons of hydrogen peroxide or 1150 tons of LOX. LOX would be less expensive than peroxide as an oxidizer - however using LOX would mean using approximately 20 metric tons of extremely expensive helium rather than about 9 tons of less expensive hydrogen for pressurization. Finally, using a cryogenic oxidizer complicates propellant valve design, fueling, and countdown operations. The third stage burns UDMH and N2O4, and uses a conventional pump-fed engine and lightweight tanks. This propellant conbination is hypergolic. If the third stage is used to deliver payloads to a space station, unused third stage propellants could be salvaged and stored for station reboost or other purposes. TANK DESIGN First and second stage tankage is welded ultra-high-strength steel. The tank diameter is 6 meters, and the fuel and oxidizer tanks together have a length of 20.4 meters. Tanks are designed with 5% ullage. Separate fuel and oxidizer tanks with no common bulkhead are used. The propane tank is on top of the peroxide tank, and an internal fuel line descends through the oxidizer tank from the propane tank. A starting slug of hydrazine is maintained in the propane feed line - this hydrazine is hypergolic with the peroxide oxidizer and initiates combustion once valves are opened to allow propellants to enter the combustion chamber. Tank mass is almost totally dominated by the requirements of pressurization, and 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 strengthing in critical areas. Pressurized propellant tanks are designed with a safety factor of 1.2 (calculated yield stress divided by normal operating stress). Before each use, tanks will be tested hydrostatically to 10% over their maximum operating pressure. For comparison, the SRB casings are designed with a safety margin of 40% but are man-rated and are used in a system which does not have inherent protection against over-pressurization (as does the P2 design with tank pressure determined by pressure regulators). The P2 preliminary design assumes the use of maraging steel, 18Ni(250), heat treated to a minimum yield strength of 1700 MPa. This gives tank walls ranging from 8 to 16 mm in thickness. Maraging steel is a high nickel steel with excellent weldability and toughness. It has previously been used for motor casings for solid rocket motors. Although maraging steel is more expensive than other high strength steels, components fabricated of maraging steel often cost less because of the ease with which it can be fabricated. The ultimate choice of steel for the tanks is not critical to the design - the use of 18Ni(250)steel in the preliminary design is intended to demonstrate that steels already in use by the aerospace industry can meet the strength requirements. Third stage tankage is of conventional light weight construction. No detailed tank design has been done, and performance estimates have been made assuming that the third stage has a mass fraction of 0.95, a value consistent with current practice in upper stage design. TANK PRESSURIZATION SYSTEM Tank pressures are high, so masses of pressurization gases are substantial and it is highly desirable to use a pressurization gas with a low molecular weight. In the baseline P2 design, stage 1 and 2 tanks are pressurized with room temperature hydrogen gas. Liquid hydrogen is stored in the booster in a light weight non-pressurized tank. It is pumped to high pressure using a turbopump similar to the hydrogen pump for a liquid hydrogen/liquid oxygen rocket engine. The turbopump is powered by decomposition of hydrogen peroxide (approximately 2 kg/sec), which is stored in a small tank which is pressurized to eliminate the need for a peroxide pump. The turbopump must pump about 30 kg/sec of liquid hydrogen - this may be compared to the 37 kg/sec output of the hydrogen pump of the J2 oxygen/hydrogen engine. The stream of liquid hydrogen is vaporized and warmed to room temperature in a gas generator which burns a small proportion of the hydrogen with hydrogen peroxide from the same pressurized supply. The hydrogen pressurization gas is non-reactive with the propane fuel. It may at first glance seem unsafe to pressurize the oxidizer tank with hydrogen, but this should be perfectly feasible. The pressurization gas is at room temperature, and hydrogen peroxide has negligible vapor pressure at this temperature. The gas over the peroxide will therefore not have sufficient oxidizer in it to achieve an explosive mixture. The tank pressurization system and its fluid supply tanks are located in the hour-glass shaped 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 water tight bay for the tank pressurization equipment, avionics etc. The booster diameter of 6 meters is determined in part by the need to have this intertank bay large enough to contain the tank pressurization system and other equipment. The current design relies on a gas generator to vaporize the hydrogen. Engine heat from the combustion chamber could also be used to evaporate the hydrogen. 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 hydrogen and to conduct the vaporized hydrogen 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 throat and nozzle are not available as a heat source as they are ablatively cooled). Finally, pressurization gas is not available until the engine is running, which complicates engine startup. In the event that hydrogen pressurization is deemed unsuitable, little performance would be lost by switching to helium. The gas generator would be more complex, as it would need a source of fuel in addition to the peroxide oxidizer. The main problem would be the high costs involved in providing the large amounts of liquid helium used for each flight. ENGINE AND NOZZLE The P2 launch vehicle uses a new pressure fed engine created by joining a designed-to-cost regeneratively cooled pressure fed combustion chamber to the ablatively cooled steerable throat and nozzle assembly from the shuttle SRB. This yields an extremely high thrust engine (approximately twice the thrust of an F1 engine) which should have low development costs as it has no pumps, burns dense storable propellants, and uses an existing nozzle and thrust vector control system. The major component needing development is a very large combustion chamber to burn hydrogen peroxide and propane. The chamber is regeneratively cooled with propane, which is a low viscosity fuel with good cooling properties. On an SRB, the nozzle is bolted to the bottom of the motor casing and has an ablatively cooled throat which protrudes into the solid motor chamber. The casing to nozzle joint is protected by insulation. The P2 combustion chamber would need to be designed to accommodate the standard attachment fittings of the nozzle, and direct the gas flow into the nozzle throat. Propellant flow for the combustion chamber is approximately twice that for the combustion chamber of an F1 engine while the chamber pressure is about 80% of the F1 pressure. In mass calculations, this combustion chamber and its associated propellant flow control valves has been assigned a mass of 5 metric tons - this figure may need to be revised once a specific design for the chamber is executed. Together with a nozzle assembly with a mass of 10 tons, the engine would mass 15 tons in total. This gives a sea level thrust/mass ratio of 800 kN/ton, similar to that of the F1 engine. In comparing the two engines, 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 a regeneratively cooled equivalent. The standard SRB nozzle has a throat diameter of 1.37 m, an exit cone diameter of 3.76 m, and an area expansion ratio of 7.53. It is designed for sea level operation with a combustion gas pressure of approximately 6 MPa. This nozzle will be used with little or no modification for the first stage P2 boosters (a minor adjustment to the expansion ratio may be desirable in the light of the different chemical nature of the exhaust gases). For second stage use, the nozzle will be modified to give a higher expansion ratio and a lower thrust. A throat insert will be used to narrow the throat to 1.1 meters, while an extended nozzle with an exit diameter of 5 meters will be used. This will achieve an area expansion ratio of approximately 20. The current SRB nozzle design is already built with the lower part of the nozzle as a replaceable item, as the bottom of the nozzle is sheared off by a linear charge before water landing in order to reduce shock to the nozzle as it hits the water. Thrust vector control is achieved by two hydraulic actuators powered by hydrazine-burning auxiliary power units - this is the same re-usable system employed on the SRBs. Since the nozzle assembly flexes for thrust vector control, the combustion chamber can be rigidly mounted to thrust structures and to its propellant feed lines and valves. Propellant for the engine is routed through hydraulically controlled throttling valves, which take their power from the thrust vector control hydraulic system. Propellant flow and mixture ratio is adjusted on a feedback basis by the throttling valves. With good propellant level sensing, this will ensure that propane and peroxide run out at the same time. It may prove advantageous to vary the mixture ratio of the propellants during flight. This is potentially possible by programming the propellant throttling valves, but has not been considered in preliminary performance calculations. It is also possible to throttle the engines partway through the flight if necessary to lower aerodynamic forces. In performance calculations, the burn-out mass of the lower stages included the mass of tank pressurization gas, which will however be propulsively vented before staging. No allowance is made in calculations for residual unburned propellants. To counterbalance the assumption of 100% propellant utilization, no credit is given for propulsive venting of tank pressurization hydrogen. The theoretical Isp of the propane/peroxide combination was estimated from published values for the theoretical equilibrium expansion of the products of combusting RP-1 with LOX (Isp 300), propane with LOX (Isp 306) and RP-1 with H2O2 (Isp 278). By using ratios based on these values, propane/peroxide is estimated to give a theoretical Isp of 284 under standard conditions (1000 psi chamber pressure, expansion to atmospheric pressure). The theoretical performance of the P2 engines (6 MPa chamber pressure, expansion ratio = 7.53 for stage 1, = 20 for stage 2) was estimated as 284/300 times the theoretical values for RP-1/LOX (obtained from published tables for different chamber pressures and expansion ratios). Finally, theoretical values were multiplied by 0.95 to estimate the actual performance of the engines. The first stage engine is predicted to have an Isp of 261 at sea level, and 291 at altitude, while the second stage engine is predicted to have an Isp of 311 in a vacuum. For comparison, the SRB has a vacuum Isp of 262. The third stage engine is a slightly modified Viking IV engine, used in a single engine configuration in the Ariane IV second stage. A nozzle extension is added to the Viking IV engine to give an Isp of 320. This engine was chosen as it is a modern UDMH/N2O4 engine which in its various variants is being produced in productions runs numbering in the hundreds (each Ariane IV launch uses four Viking VI engines as strap-on boosters, four Viking V engines for the first stage, and one Viking IV engine for the second stage). Other similar engines could be used. The Viking IV engine normally burns 34 tons of propellant in 123 seconds. In the P2 design, it is assumed that longer burn times are feasible, and the third stage burns 190 tons of propellant in a period of about 750 seconds. Thrust vector control and roll control during the first stage burn is via co-ordinated steering by the first stage engines. Thrust vector control for the second stage is via engine steering, and roll control is achieved by small thrusters utilizing hydrogen peroxide as a monopropellant. The thrusters and control equipment for this purpose are mounted in the inter-tank area, and use peroxide from the same supply tank as is used for the tank pressurization system. Thrust vector control for the third stage is via gimbaling of the Viking IV engine, while roll control is via secondary thrusters which are part of the standard Ariane second stage propulsion system. OPERATIONS ISSUES The P2 boosters are roughly the same mass as empty SRBs but are somewhat shorter and have a larger diameter (6 meters vs. 3.7 meters). Handling and erection of the booster prior to launch is considerably easier than handling for SRBs, as no stacking of segments is needed and the vehicle is erected empty at the launch pad. The boosters are fat and short, which increases aerodynamic drag but gives lighter tanks and easier ground access. Both main propellants are storable, and hydrogen pressurization gas is stored as a liquid in a heavily insulated tank with a low loss rate - booster fueling can thus precede launching by a convenient interval, propellant top-up systems are not needed during the launch countdown, and there is no need for propellant removal in the event of a launch delay. On an SRB, the nozzle is surrounded by an after skirt which shields the thrust vector control equipment and acts as a support structure while the rocket is on the launch pad (the motor casing itself cannot be used for support, as the exit cone of the nozzle is wider than the casing). This skirt is not needed with the P2 first stage boosters, which have a larger diameter than the nozzle exit cone. The P2 vehicle will be directly supported by interfacing the bottom of the two first stage boosters with a support structure on the pad. The second stage will be supported on the pad by the first stages. Much of the hardware on the P2 booster is derived from that on the SRB, as this reduces development costs. The SRB parachute system is used for the P2 booster - the recovery of first stage hardware weighing the same as the empty P2 first stages has repeatedly been shown by routine SRB recoveries. Water recovery of the P2 boosters should be simpler than that of SRBs, as the booster is inherently buoyant and recovery operations would not need the nozzle plug and dewatering operations involved in SRB recovery. The SRB nozzle used in the P2 booster is designed to survive water landings and salt water immersion. The combustion chamber would need to be constructed so that it would not be damaged by salt water entering the throat. The combustion chamber, throttling valves for propellants and their associated hardware, as well as the hydraulic pumps for the thrust vector control system will be shielded by a water-tight structure at the base of the booster. Avionics and tank pressurization equipment (turbopump, gas generator, pressure regulators etc.) are in a watertight bay in the angle between the two propellant tanks. After recovery, the first stage boosters would need some refurbishment before relaunch. The tanks would be emptied of residual propellants and hydrostatically tested. The insulation on the ablatively cooled nozzle would have to be replaced, or the nozzle swapped for a refurbished one. The parachutes would have to be refitted, the separation motors replaced, and the TVC and tank pressurization systems checked. If the booster is to be converted to a second stage, the parachute mounting equipment would be removed, a third stage adapter fitted, and a high expansion ratio nozzle installed. COSTS Payload capabilities to low earth orbit are calculated to be approximately 50 metric tons (to the nearest ton), which is about twice the current limit of the shuttle and about 5 times that of the Ariane 44L. No estimates have been made of payload delivery costs, although these are expected to be substantially lower than that of current boosters. Studies have show that much of the cost of launch operations is dominated by the number of components and operations involved in a launch, rather than the size of the components. It is useful therefore to compare the expenditure of equipment between a P2 launch an Ariane 44L launch. Per flight: Ariane 44L P2 Payload delivered to LEO (tons) 9 50 Expenditures: Engines 10 2 Tank sets 7 2 DESIGN OPTIMIZATION The relative masses of the three stages was initially set to a ratio of 6 : 3 : 1. Information about stage masses, available thrust, mass fractions, and specific impulse were entered into a computer spreadsheet which calculated the total velocity which could be imparted to a payload of a given size. The "goal seeker" function of the spread sheet was then invoked to determine the payload which resulted in a nominal velocity of 9300 m/sec (low earth orbit typical requirement, including gravity and air resistance losses). The "solver" function of the spreadsheet was then used to systematically vary the stage masses in order to maximize the velocity given to the payload, subject to a series of constraints (ie. liftoff thrust must be greater than 1.2 times total mass etc.). Following this another "goal seeker" round was performed to determine the payload which could be taken to exactly 9300 m/sec. The overall design is governed by the thrust of the SRB nozzle derived engines (approximately 12,000 kN each at sea level, 15,000 kN each at altitude for sea level nozzles). Based on existing practice with large liquid fueled launch vehicles which do not have solid boosters, initial acceleration was required to be 1.2 G or greater. The use of 2 engines (total of 24,000 kN thrust) for lift off thus puts an upper limit of approximately 2000 metric tons on the overall vehicle mass. The second stage was also required to start with 1.2 G initial acceleration. Optimization of stage sizes always drove initial accelerations for stages 1 and 2 down to 1.2 G. Using a single Viking IV engine, third stages optimized at approximately 300 to 400 metric tons mass with burn times in excess of 1000 seconds. Payloads however were not strongly dependent on third stage mass, and to minimize costs associated with the expensive fuel and non-recoverable hardware of the third stage and to minimize burn time, the stage was arbitrarily designed to have a total mass of 200 metric tons. The first stage engines were baselined to a chamber pressure of 6 MPa (approximately 900 psi) which is the chamber pressure produced in an SRB at full thrust. With a fixed engine size, lower chamber pressures resulted in lower available take-off thrust, and thus a smaller vehicle. Losses in vehicle size were not made up for by lower tank masses, so first stage chamber pressures lower than 6 MPa result in a loss of payload. Chamber pressures higher than 6 MPa would allow a larger vehicle and higher payloads - this option is discussed in the "enhancements" section below. Second stages were investigated at chamber pressures from 2 MPa to 6 MPa. Optimum results were obtained at a chamber pressure of 6 MPa using a nozzle extension and throat insert to increase the engine expansion ratio. The throat insert was sized to provide sufficient thrust to achieve an initial acceleration of 1.2 G for the second stage. Optimization indicated little improvement in performance by using different sized first and second stage rockets. The basic hardware for the first and second stage were therefore made identical to lower development and production costs. A given rocket can be flown either as a first stage (with parachutes for recovery) or as a second stage (with an increase expansion ratio nozzle, and with a third stage adapter taking the place of the parachute equipment). In practice this means that after a given rocket has flown twice as a first stage, it can be easily modified to a second stage and expended. All tank and engine sets are thus used three time before being discarded. MASS BUDGET (METRIC TONS), STAGES 1 AND 2 Masses for SRB-derived equipment marked are from the literature except for the thrust vector control system (which is a guestimate), and are rounded upwards to the nearest metric ton. Frustrum, parachutes, forward skirt (stage 1) or third stage adapter (stage 2) 6 Fuel tank 8.93 Oxidizer tank 27.3 Pressurizing gas 2.89 Pressurization equipment (pump, peroxide supply etc.) 2 Inter-tank fairing 3 Separation motors (stage 1) or nozzle extension (stage 2) 1 Combustion chamber and valves 5 Nozzle 10 Thrust vector control pumps, actuators 2 Avionics 1 Miscellaneous Structure and Equipment 5 Propellant 521 Total Ignition Mass 595.12 Total Burnout mass 74.12 Mass fraction .875 SENSITIVITY ANALYSIS The performance of the P2 vehicle has been computed using best estimates of the relevant parameters. Real values could be better or worse than those assumed. However, because the P2 launcher is a three stage design the delivered payload is generally relatively insensitive to the assumptions made for individual stages. This is demonstrated by the fact that even if all mass and Isp estimates are substantially optimistic, the vehicle will still deliver over 75% of the originally calculated payload. Condition Payload P2 launcher preliminary design 100% stage 1 and 2: burnout mass 10% heavier 95% stage 3: burnout mass 10% heavier 98% stage 1,2, and 3: burnout mass 10% heavier 92% stage 1 and 2: Isp 5% lower 90% stage 3: Isp 5% lower 91% stage 1, 2 and 3: Isp 5% lower 81% all of above simultaneously 75% It is seems unlikely that the burnout mass of stages 1 and 2 would be more than 10% higher than estimated. A rough demonstration that the P2 burnout mass is approximately correct may be obtained by noting that the mass is comparable to that of current-version shuttle SRB. The P2 booster has components not found in the SRB (combustion chamber, tank pressurization system), while the SRB has components not found in the P2 booster (case insulation, after skirt). The P2 tanks enclose a somewhat larger pressurized volume than the SRB motor casing, but are designed to a lower safety factor with a higher strength steel and do not have the penalty of requiring elaborate pinned and sealed joints for assembly. No detailed design has been done of the 200 ton third stage - it is assumed to have a mass fraction of 0.95 and thus a burnout mass of 10 tons. Based on current aerospace practice, this is probably conservative for a single engine UDMH/N2O4 upper stage with a low thrust/mass ratio. For comparison, the slightly larger high thrust/mass ratio Ariane IV first stage has four Viking engines. This stage holds 226 tons of propellant and has a dry mass of 17 tons, giving a mass fraction of .93. Isp values are only estimated, and could easily be in error. For the stage 1 and 2 engines, efficiency (realized Isp as a fraction of theoretical Isp) has the potential to be high. A realized/theoretical Isp ratio of 0.95 has been assumed in designing the booster - from limited information in the literature, this figure seems to be reasonable for an engine where no main propellants are used for secondary uses such as turbopump operation. I have to date been unable to obtain information about the Ariane IV engine. Somewhat arbitrarily it has been assigned an Isp of 320. This value is that of an UDMH/N2O4 engine with a realized/theoretical ratio of 0.95 and an expansion ratio of approximately 40. If this expansion ratio is not present on the stock Viking IV engine it could probably be easily realized with a nozzle extension. ENHANCEMENTS The 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. To this end, the booster diameter has been chosen to be 6 meters to allow sufficient intertank volume for more pressurization gas. Both the oxidizer and fuel tank have a cylindrical barrel section which would be trivial to make longer in a stretched booster. The combustion chamber and nozzle would possibly require modifications to operate at higher pressures and heat transfer loads. An alternate approach to larger payloads would be to derive a clustered vehicle in which four boosters would lift a high pressure central booster with enlarged tanks. A substantially larger payload could be orbited by using a hydrogen/oxygen upper stage. If the 200 ton storable propellant third stage were replaced with a similar mass hydrogen/oxygen stage with a mass fraction of 0.9, then delivered payloads to LEO would rise to about 75 tons (from 50 tons). This stage would need a thrust of approximately 1000 kN, and could be derived by stretching the slightly smaller Ariane V hydrogen/oxygen main stage, which will use the 1025 kN HM60 engine and 155 tons of propellant. Stages much larger than about 200 tons would be too bulky for the lower stages. Given the trouble that a cryogenic stage would add, the added payload may not be worth it. For occasional very large payloads, it would probably be easier to perform on-orbit assembly or use a P2-derived clustered vehicle. It might be possible to recover the second stage booster by parachute. The velocity and range achieved by the second stage normally make this impossible. However, the following scenario may bear investigation: 1) The engine of the second stage booster is shut off before all propellant is burned, and the third stage is separated. 2) The second stage (at this point in free-fall in a vacuum) re-orients itself, pointing roughly back along its ascent trajectory. 3) The second stage re-ignites, kills its forward velocity, and gains enough velocity to travel ballistically back to the prime recovery area. 4) The falling second stage is recovered by parachute. The third stage tankage is probably impossible to recover in any cost effective manner (although as mentioned above, residual propellants could be salvaged at a space station). If the third stage were docking at a space station, it might be economically feasible to detach the propulsion and avionics from the tankage, and return them to earth by shuttle for refurbishment and reuse. The basic P2 booster is designed for reliability via simplicity. For very high value payloads, some of the high payload capability of the launch vehicle could be used for even further safety. High value payloads could be enclosed in a reinforced re-entry capsule with emergency escape rockets. As was done in the pre-shuttle manned spaceflight era, this would give assured payload recovery for almost any potential in-flight malfunction. UNRESOLVED ISSUES I would like to eventually write this launcher concept up for publication. Before this happens, I feel that considerable more work is needed. I would like to get more information on the following: - Cost of hydrogen peroxide, assuming long term supply contracts in the multi-thousand ton per year range. - Cost of liquid hydrogen and helium in lots of tens of tons. - Raw material costs for maraging steel, 18Ni(250), for tanks. Note that there are at least three types of 18Ni(250). The original alloy contains 8.5% cobalt, but when cobalt got expensive in the late 1970s, a 2% cobalt alternative and a cobalt free alternative 18Ni(250) were developed. The latter seem to have very similar mechanically properties, but are presumably less expensive. - Any cost information at all on SRB components (does anyone have access to financial details for NASA contracts for SRB nozzles, parachutes etc.) ? - More solidly grounded estimates for the working Isp of the various engines involved. - Any detailed suggestions on propane/peroxide combustion chamber design (NB. older designs buring kerosene and peroxide did not use 100% peroxide, and were more concerned with easy ignition than ultimate Isp; many older designs therefore decomposed the peroxide catalytically before chamber injection, a procedure which is probably not needed, although it should not necessarily be ruled out). - Anything else which would have a bearing on either the technical feasibility or the costs of such a launcher. Thank you everyone in advance for your feedback! -- Bruce Dunn Vancouver, Canada Bruce_Dunn@mindlink.bc.ca ------------------------------ End of Space-tech Digest #110 *******************