AUSROC III PROPULSION SYSTEM Author Mark Blair (B.E.Mech) AUSROC Program Coordinator Australian Space Research Institute First Issue 1992 INTRODUCTION Essentially all rocket system designs begin with the determination of the mission objectives. In this case the mission objective has been determined "to carry a useful payload of 100 kg mass to an altitude of 500 km on a predetermined and controlled suborbital trajectory and return it intact". Several custom particle trajectory analysis programs were written(ref1-2) to size the propellant requirements to achieve the mission objectives with the constraints mentioned in table #4. From these simulations, it was determined that 1200 kg of propellant would be required to achieve the objective. The next step is to determine the propellant combination to use. Once this is known a detailed analysis of the motor geometry and materials of contruction can be presented. PROPELLANT SELECTION There is a large variety of liquid propellant combinations which have been analytically and experimentally investigated. Unfortunately, it has not been possible to discover an ideal liquid propellant combination which will have only desireable characteristics. Almost every liquid propellant, especially every liquid oxidizing agent, has at least one or more undesireable properties, and no standard liquid propellant has yet been developed. A monopropellant contains an oxidizing agent and combustible matter in a single substance. It may be a mixture of several compounds, or it may be a homogeneous chemical agent. A bipropellant rocket has two separate propellants which are mixed inside the combustion chamber. The majority of successful liquid propellant rockets have used bi-propellants. Occasionally rockets with three or more liquid propellants have been used, but never very extensively. Due to the fact that no single propellant has all desirable properties, the selection of the propellant combination is a compromise of many parameters as discussed in Sutton8 and listed below: Economic: - availability in large quantities - low cost - logistics of production - simplicity of production process Performance: - specific impulse - effective exhaust velocity - specific propellant consumption - high energy content - low molecular mass Hazards: - corrosive effects - explosive hazards - fire hazards - toxicity Desirable Properties: - low freezing point - high specific gravity - chemical stability - high specific heat - high thermal conductivity - high boiling point - low viscosity - low vapour pressure Propellants can be classified into two groups according to the type of ignition used. The first group is termed hypergolic due to the fact that these combinations ignite spontaneously upon contact with each other. Nonspontaneously ignitable propellants have to be heated by external means before ignition can begin. The propellants considered for use in AUSROC III are described below: LIQUID OXIDIZERS Liquid Oxygen: Chemical Formula - O2 Boiling point - 90 K Specific gravity - 1.14 Liquid oxygen is currently used in conjunction with alcohols, jet fuels (kerosene types), gasoline and hydrogen. It burns with a bright white-yellow flame with most hydrocarbons and usually does not burn spontaneously. It is a noncorrosive and nontoxic liquid. Because liquid oxygen evaporates rapidly, it cannot be stored readily for any great length of time. It is neccesary to insulate all lines, tanks, valves, etc. Liquid Fluorine: Chemical Formula - F2 Boiling point - 54 K Specific gravity - 1.5 In combination with most fuels, liquid fluorine affords higher values of performance and energy than other oxidizers. It is extremely toxic, corrosive and reactive. Special passivation techniques and insulation have to be used on containers, pipelines and valves to permit handling of liquid fluorine in common construction metals. The production of liquid fluorine is an expensive process and commercial consumption is low. Nitric Acid: Chemical Formula - HNO3 Boiling point - 411 K Specific gravity - 1.5 Red fuming nitric acid is the most common type, consisting of concentrated nitric acid and between 5-20% nitrogen dioxide. It is corrosive, toxic and requires special handling precautions. Nitric acid affords a lower specific impulse than most other oxidizers. Nitrogen Tetroxide: Chemical Formula - N2O4 Boiling point - 294 K Specific gravity - 1.44 Nitrogen Tetroxide is the most common storable oxidizer used today but its liquid temperature range is narrow and it is easily frozen or vapourized. It is hypergolic with many fuels. It is toxic and has a high vapour pressure, necessitating heavy tanks. LIQUID FUELS Hydrocarbon Fuels: Chemical Formula - CxHy Boiling point - varies Specific gravity - varies These include gasoline, kerosene, diesel oil and turbojet fuel. Their physical properties and chemical composition vary widely. They are relatively easy to handle, and there is an ample supply of these fuels at low cost. Liquid Hydrogen: Chemical Formula - H2 Boiling point - 20 K Specific gravity - 0.07 Liquid hydrogen gives high performance when burned with liquid oxygen or liquid fluorine and is an excellent regenerative coolant. Of all known fuels it is the lightest and the coldest. Special insulation provision must be used and care taken to select materials for the storage tanks. Hydrazine: Chemical Formula - N2H4 Boiling point - 386 K Specific gravity - 1.01 Hydrazine is a toxic and colorless liquid. It is spontaneously ignitable with nitric acid and nitrogen tetroxide. It is an excellent monopropellant when decomposed by a suitable catalyst. It generally gives good performance when compared with many common fuels. Unsymmetric Dimethylhydrazine: Formula - (CH3)2NNH2 Boiling point - 336 K Specific gravity - 0.61 This is a more stable derivative of hydrazine. It gives slightly lower performance than pure hydrazine and is usually used in a mixture with hydrazine itself. Monomethylhydrazine: Chemical Formula - CH3NHNH2 Boiling point - 361 K Specific gravity - 0.88 This is another derivative of hydrazine. It has better shock resistance to blast waves, better heat transfer properties and a better liquid temperature range than pure hydrazine. The selection of our propellant combination was based on 4 considerations: cost, availability, performance and ease of handling and storage. The following selection discussion has been taken from Blair and Kantzos4. Liquid flourine gives the highest performance but was eliminated due to difficulties involved in handling and lack of availability. Nitric acid is a good storable oxidizer and is readily available at low cost but it was eliminated due to its relatively low performance when compared to other oxidisers. Nitrogen tetroxide is a storable high performance oxidizer but it is not readily available in the required concentrations or quantities that we require. Liquid Oxygen is readily available in large quantities from air liquification plants at very low costs (approximately 50c/Lt) and gives high performance with various fuels. While liquid oxygen presents some storage problems due to its cryogenic nature, its other properties make it the most favourable oxidizer for this particular project. With liquid oxygen chosen as the oxidizer, we can eliminate Hydrazine fuel and its derivatives due to their lower performance with liquid oxygen and also due to their low availability in high concentrations. Liquid hydrogen gives a higher performance with liquid oxygen than any other fuel. However, liquid hydrogen has a very low density and requires large and hence heavy propellant storage tanks. It boils at around 20 K and requires very special materials to prevent hydrogen embrittlement and special insulation to prevent boil-off during storage. Liquid hydrogen has a low to nil availability in this country and was eliminated for these reasons. Alcohols are readily available at low cost, are storable and non toxic but afford lower performance than the hydrocarbon fuels and thus were eliminated. Hydrocarbon fuels give good performance in combination with liquid oxygen. They are storable, quite easy to handle and available in large quantities at low cost. Kerosene type hydrocarbons are most commonly used as rocket fuels so we chose to use readily available JA-1 Jet fuel, which is a kerosene derivative, as the fuel for this project. The LOX/Kero propellant combination is not a hypergolic mixture so an ignition system will be required to ignite the propellant at startup and from then on the combustion will be self sustaining provided that instabilities and other phenomenon do not extinguish the flame or force it outside the chamber. Table 1 shows the chemical properties of liquid oxygen (now on refered to as LOX) and JA-1 Jet fuel as obtained from the Mobil Oil company. TABLE 1 CHEMICAL PROPERTIES OF LIQIUD OXYGEN AND JA-1 JET FUEL Property Liquid Oxygen JA-1 Jet Fuel Chemical Formula O2 C7.135H14.187 Molecular Mass 32 175 (av.) Melting Point (K) 54.36 226 Boiling Point (K) 90.19 477-561 (1atm.) Density (kg/cu.m) 1141.1 (at b.p.) 800 (at 298 K) 600 (at 400 K) Specific Heat (J/molK) 1738.14 (sat.liq.) 2090 (273-373 K) Hv (kJ/kg) 218.2 246 Hc (MJ/kg) 42.8 (s.t.p.) Conductivity (W/mK) 0.150 0.16 (273 K) 0.12 (373 K) Viscosity (Centipoise) 0.87 (54 K) 0.75 (290 K) 0.19 (90 K) 0.21 (366 K) Availability Good Good Storability Poor Good Toxicity Low Low Corosiveness Low Low THERMOCHEMICAL REACTION CALCULATIONS To analyse the performance of a particular propellant combination we have to first calculate the combustion chamber gas conditions and the conditions of the exhaust products. These conditions include: - combustion and exhaust temperature - average molecular weight of products - specific heat (Cp) - enthalpy changes - entropy These can be determined, through an iteration procedure using Gibbs Free Energy, from the chemical composition of the initial propellant mixture, the prereaction temperatures of the propellants and the predetermined combustion chamber pressure (Pc). The thermochemical calculations were performed using a software package3 developed by the NASA-Lewis Research Centre in America. This package calculates the specific impulse, thrust coefficient and expansion ratio of a rocket motor in addition to the thermochemical parameters. Ausroc III is to utilize a pressure feed system to deliver the propellants to the combustion chamber. Thus the propellant tanks must hold a pressure in excess of the chamber pressure. High combustion pressures yield high Specific Impulses but, in a pressure fed rocket, this would result in heavy propellant tanks. A combustion pressure of 2 MPa, however, is a good compromise between tank weight and specific impulse and was chosen as the combustion pressure for Ausroc III. Fig. 1 shows the variation of specific impulse with propellant mixture ratio (Mox/Mf) based on a nozzle optimised for sea-level operation. Based on these results a mixture ratio of 2.4 has been chosen for the Ausroc III propulsion system. NOZZLE GEOMETRY DESIGN As a rocket climbs in altitude, the atmospheric pressure (Pa) drops. At high altitudes, where ambient pressure is low, large nozzle expansion ratios can be used to gain higher specific impulses. However, if large expansion ratio nozzles are used at sea-level, flow separation occurs in the nozzle and causes thrust vector control problems. Thus, there is a trade-off to be made. In general, flow separation will not occur until the nozzle exit pressure (Pe) falls below 0.4 x ambient pressure8. To be safe the Ausroc III propulsion system will be designed to expand the combustion gases to 0.55 x ambient pressure at sea-level. This corresponds to an expansion ratio (Ae/At) of 6 where Ae is the nozzle exit area and At is the nozzle throat area. Hence, the rocket motor exhaust gases will be overexpanded from sea-level to around the 5 km altitude mark where the ambient pressure is 0.55 times sea- level (55kPa). At this point the exhaust gases will be optimally expanded. As the rocket continues to climb in altitude, the exhaust flow becomes underexpanded. As this process occurs, the overall thrust generated by rocket motor will increase. This increase in thrust corresponds to an increase in specific impulse and the thrust coefficient (Cf). The thrust coefficient is given by: Cf = Cf(opt) + ((Pe - Pa)/Pc)(Ae/At) Where Cf(opt) is the value for optimum expansion, ie. where the ambient pressure equals the nozzle exit pressure. For an expansion ratio of 6 using Liquid oxygen and kerosene at a combustion pressure of 2MPa, Cf(opt) = 1.533. The 2 extremes of thrust coefficient can now be calculated for sea-level and vacuum conditions as: Cf = 1.394 (s.l.) = 1.698 (vac.) To calculate the propellant specific impulse (Isp) for the 2 extremes of operation we use the following relationship: Isp = c* Cf / 9.81 {Ideal} The parameter c* is the characteristic exhaust velocity and is a figure of merit of the propellant combination, independant of the nozzle expansion ratio. The Lox / Kero system at 2 MPa combustion pressure produces a c* value of 1804 m/s. Thus, at the 2 extremes of operation: Isp = 256 sec (s.l.) = 312 sec (vac.) {Ideal} These values, as described, are ideal values assuming no system losses. In reality losses do exist and can be accounted for by the inclusion of a thrust correction factor (zt). For the Ausroc III motor a correction factor of 0.94 has been assumed. Hence: Isp = 241 sec (s.l.) = 293 sec (vac.) {Corrected} To achieve resonable acceleration figures for the vehicle, it was found, from the trajectory simulations (1-2), that an initial thrust of 35 kN would be required. The nozzle throat area required to achieve this thrust at sea-level is obtained using the thrust equation: F = zt x Cf x Pc x At At = 0.013355 m2 (130 mm diameter throat) In a vacuum, with this nozzle size, the thrust obtained is 42.6 kN, so the thrust range is: F = 35,000 N (s.l.) = 42,600 N (vac.) The expansion ratio has already been discussed and set at 6. Therefore the nozzle exit area is 6 times the throat area: Ae = 0.08013 m2 (320 mm diameter exit) The Combustion chamber Contraction Ratio (Ac/At) is used in conjunction with the characteristic chamber length: L*=Chamber Volume / Throat Area to determine the combustion chamber size. Both these parameters are, basically, determined by experiment and have been characterised with years of experience5-6. The L* value for a liquid oxygen and kerosene propellant combination is 1.0 m. The contraction ratio has been set at 3 for the Ausroc III motor as this value is typical for pressure fed, low thrust rocket motors. Thus: Ac = 0.04 m2 (226 mm diameter chamber) Chamber Volume = 0.013355 m3 Chamber length = 334 mm To simplify the manufacture of the motor a conical nozzle will be used with a cone half angle (a) of 15 degrees. Low half angle nozzles are higher performing, but are longer and, hence, heavier than the large angle nozzles. Thus a trade-off was required. The performance correction factor using a conical nozzle is given by: l = 0.5(1+cosa) For a 15 degree half angle, the correction factor is 0.983. This factor is incorporated in the thrust correction factor mentioned in a previous section. The remaining geometrical features to determine are the contraction half angle (ac), the throat longitudinal radius (ru) and the contraction approach radius (rc). Small radii and large contraction half angles reduce the motor length but increase the aerodynamic losses. In the contraction section the flow is subsonic and losses are not as great as the supersonic exit cone. However, the shallower angles and larger radii have been found to be effective in boundary layer film cooling through the throat. The 2 radii are usually expressed as ratios with the throat cross-sectional radius (rt)5-6. The preliminary Ausroc III motor design will have values of: ru/rt = 1.0 rc/rt = 1.0 ac = 30 degrees Fig.2 shows the internal geometry of the Ausroc III rocket motor and provides the nomenclature for the various parameters used in its design. Table 2 shows the motor specifications Table 2 AUSROC III Motor Specifications Fuel: Kerosene Oxidiser: Liquid Oxygen Mixture Ratio (Ox/F): 2.4 Thrust Coefficient: 1.394 s.l. - 1.698 vac. Specific Impulse (sec): 241 s.l. - 293 vac. Thrust (N): 35,000 s.l. - 42,600 vac. Nozzle Throat Diameter: 130 mm Nozzle Exit Diameter: 320 mm Nozzle Expansion Ratio: 6 Expansion Cone Half Angle: 15 degrees Chamber Diameter: 226 mm Characteristic Length (L*): 1.0 m Chamber Length: 340 mm Contraction Ratio: 3 Contraction Cone Half Angle: 30 degrees Longitudinal Throat Radius: 65 mm Longitudinal Contraction Radius: 65 mm MOTOR COOLANT SCHEME The Ausroc III motor could be manufactured in a number of different ways depending on the materials used and cooling methods employed. There are 4 primary methods for cooling rocket motors: Coolant Jacket (Regenerative) Ablative Cooling (Sacraficial) Radiation Cooling Film Cooling (Boundary Layer) In the coolant jacket method, propellant (fuel or oxidizer) is circulated through passages along the motor wall to absorb the heat transfered through the wall. These motors are usually manufactured by brazing pre-formed steel tubes to form the motor geometry and adding manifolds for coolant inlet and outlet. Theoretically, these motors can be fired indefinately. They are relatively light weight but can be time consuming to manufacture. Ablative motors are one shot devices used, primarily, in short burn liquid motors or solid propellant motors where a liquid coolant is not available. These motors use endothermic materials which decompose and absorb large quantities of heat in the process to keep wall temperatures low. Composite materials utilising phenolic resins and either carbon, graphite or silica fibres are usually used and the motors are manufactured by a filament or tape wrapping process on a mandrel. In Radiation cooled motors, external radiation losses from the wall material balance the convective heating from the combustion products, thereby allowing the chamber wall to operate in thermal equilibrium. This requires the use of rare and expensive high temperature refractory metals and ceramics. This method has been used for small low thrust motors and for nozzle extensions where the temperatures are lower. Film cooling can be incorporated into any of the above types of motors and involves injecting a coolant fluid along the motor wall to generate a 'cool' gas boundary layer to slow the rate of heat transfer. Of all the cooling methods listed above, the boundary layer cooling is the easiest to implement. The ablative motor is the easiest and cheapest to manufacture but the regenerative motor gives the highest performance and longest duration. The Ausroc III program will require numerous static firings before the launch can be approved. The regeneratively cooled motor can be reused for many firings and there are no erosion problems. Thus we decided to use this type of motor for Ausroc III. The manufacture method has not been determined yet but we are considering a tube wall method as well as a machined slotted wall method. PROPELLANT REQUIREMENTS The propellant requirements can be calculated with a knowledge of the specific impulse, thrust level, mixture ratio and burn time. These valuse have been determined as: Isp = 241 sec (sl) - 293 sec (vac) T = 35 kN (sl) - 42.6 kN (vac) Mr = 2.4 (Lox/Kero) tb = 80 sec The propellant mass flow is calculated using the following relationship: dm/dt = T / Isp g = 14.8 kg/s (at sea level & in vacuum) The total propellant mass is just the burn time multiplied by the propellant mass flow rate and is equal to: Mass of Propellant = 1184 kg Using the propellant mass ratio we can determine the required masses of the Lox and Kerosene. These are found to be: Mass Lox = 836 kg Density of Lox = 1142 kg/m3 Volume of Lox = 732 lt Mass Kerosene = 348 kg Density of Kerosene = 800 kg/m3 Volume of Kerosene = 435 lt At this stage, it seems as though boundary layer film cooling will be employed to assist in the chamber wall cooling. The Kerosene will be used as the film coolant and will constitute approximately 5% of the total kerosene flow or 17.4kg (22lt). The chamber gimballing actuators will use the pressurised kerosene as the actuating fluid and thus further increases the requirement. Taking these volumes into account and allowing for a tank ullage of around 10%, the tank volumes have been set at: Lox Tank Volume = 800 lt Kerosene Tank Volume = 500 lt A nominal inside diameter for the tanks and rocket body has been set at 700mm. The rocket, as mentioned in previous sections, is pressure fed. There are 2 gases that have been identified as being applicable to this application; Nitrogen and Helium. Nitrogen has a boiling point only 10 degrees K less than oxygen. This results in a percentage of the nitrogen compressing and being absorbed into the liquid oxygen. It has been found that almost double the amount of nitrogen is used when compared to helium as a result of this compression and absorption. Nitrogen has a molecular weight of 28 as opposed to helium which has a molecular weight of 4, ie nitrogen has a 7 fold increase in weight as opposed to an equivalent requirement of helium. Nitrogen, however, is very cheap and readily available in large quantities. Helium is expensive but for the above reasons, it has been chosen as the pressurant for the flight vehicle. For the static firings and ground tests, nitrogen will be used as the pressurant since weight is not of concern in these trials. The pressurant tank will store the gas at high pressure (30MPa). This high pressure gas will then be regulated down to the tank pressures of 3MPa. Thus, a pressure tank volume of 150 lt, which includes a safety excess of 20 lt, will be required. CONCLUSION In this paper, the initial objective has been used to determine the propulsion system requirements. In conjunction with the simulations performed we were able to size the rocket motor geometry and propellant requirements. Arguments were forwarded to determine the type of rocket motor to use and an ablative type construction was selected. A winding procedure was selected as the manufacturing method and a carbon fibre / phenolic resin material was chosen as both the ablative and structural member of the motor. REFERENCES No. Author Title 1. Coleman J. "A Basic Particle Trajectory Simulator" Computer Program (c) Ausroc 1990 2. Cheers A. "A Spherical Earth Model Particle Trajectory Simulator Utilizing a 4th Order Runge-Kutta Method" Computer Program (c) Ardebil 1991 3. Gordon S. and "Computer Program for Calculation of McBride B. Complex Chemical Equilibrium Compositions, Rocket Performance, Incident and Reflected Shocks and Chapman-Jouguet Detonations" NASA SP-273 1967 4. Blair M. and "AUSROC II Final Report" Kantzos P. Monash Uni. October 1989 5. Huang D. and "Design of Liquid Propellant Rocket Huzel D. Engines" NASA SP-125 1971 6. Gill G. "Liquid Rocket Engine Nozzles" Hyde J. NASA SP-8120 7. Evensen H. "Liquid Rocket Engine Self Cooled Ewen R. Combustion Chambers" NASA SP-8124 8. Sutton G. "Rocket Propulsion Elements" John Wiley & Sons 1986