AUSROC IV ORBITAL CAPABILITY December 1991 Author Mark Blair (B.E.Mech) Rocket Technology Defence Science and Technology Organisation ABSTRACT This paper discusses some preliminary aspects of an amateur micro-satellite launch vehicle based on Ausroc III hardware. The requirements for the placement of a satellite into orbit are mentioned and the results of some basic orbital trajectory simulations are presented based on the performance estimates of the Ausroc III vehicle. INTRODUCTION The ultimate goal of the Ausroc Program is to develop the capability to place a micro-satellite (20-40kg) into a polar orbit. The Australian Space Engineering Research Association (ASERA) is a not for profit organisation based in Sydney with 2 major projects currently in the works. The first is to design and build a micro- satellite and the second is the Ausroc Program, aimed at developing rocket hardware to, eventually, place a micro-satellite into orbit. One target of these 2 projects is for them to eventually unite such that an amateur satellite can be launched into orbit on an amateur rocket. A project of this type and magnitude has not yet been undertaken anywhere in the world. AUSROC III PERFORMANCE CONSTRAINTS In order to determine if a satellite launcher can be developed from Ausroc III hardware it has been necessary to perform some simplified trajectory simulations based on several staging configurations. The data that has been assumed from the Ausroc III specifications1 for the purpose of these simulations is: Vehicle Mass Ratio (Mprop/Mtotal) = 0.85 Propellant Mass per Module = 1200 kg Specific Impulse = 260 sec liquid(stage 1) = 320 sec liquid (stages 2-4) = 290 sec solid (stages 2-4) The 2 values of liquid specific impulse are for the normal Ausroc III nozzle with an expansion ratio of 6 and for a modified high altitude nozzle with an expansion ratio of 20 to increase the specific impulse. An expansion ratio of 20 could not be used at sea level because of separation of the gases from the nozzle wall. The nozzle cone extension does not represent a major redesign and is very easy to implement. A possibility exists for using solid propellant upper stages. If this is the case then the solid specific impulse, at altitude, will be assumed to be 290 sec. ORBITAL PARAMETERS For an object to be in orbit, 2 conditions must be achieved. Firstly, the object must have sufficient height to clear the atmosphere and prevent aerodynamic drag from decaying the orbit and causing re-entry. Some spacecraft have orbited as low as 150 km but the orbital stay times have been low. The target altitude for the Ausrov IV payload is 300 km which should give an orbital stay time of several years at least. The second condition is that the object has sufficient tangential velocity (v) to counter the effects of the earth's gravity (g). The following relationship can be used to determine the orbital velocity required to maintain an orbit: g = K me / r2 = v2 / r Where: K = 6.673e-11 me = 5.976e24 (kg) r = orbit radius (m) The average radius of the earth is 6.371e6 m. For a 300 km orbit the radius r=6.671e6 m. This gives an orbital velocity of 7732 m/s. This value assumes a spherical non-rotating earth. Satellite orbits can have inclination angles from 0-90 degrees to the equator. Equatorial orbits are those with low inclination angles and can take advantage of the velocity at the surface of the rotating earth. This velocity advantage is 465 m/s for launch sites at the equator and drops off to 0 at orbital inclinations of 90 degrees or from launch sites at the poles. Figure 1 shows these various orbits. It is envisaged, at this stage, that the satellite to be orbited will carry a remote sensing instrument. To get complete global coverage for such an experiment, it will be necessary to place the satellite into a polar orbit. In this case there is no advantage from the earth's rotation. For the purpose of the simulations a velocity loss of 1500 m/s has been assumed2 for gravity and drag losses. Thus the total velocity to be gained is 9232 m/s. ROCKET MODULE STAGING For a rocket utilizing liquid oxygen and kerosene propellants to attain orbit with only one stage, a vehicle mass ratio of 0.96 must be achieved. This mass fraction would be very difficult to achieve even with state of the art materials and turbopumps. For this reason Ausroc IV will need multiple stages to reach the desired velocity. Minimizing the number of stages, reduces the system complexity but results in larger growth factors. The Growth Factor is a ratio of launch weight to payload weight. The velocity equation used to obtain the theoretical maximum velocity that a rocket will reach in a drag and gravity free environment is given by3: Velocity = Isp g ln(Minitial/Mfinal) This equation can also be used for multi staged vehicles to optimise the stage masses. A short program4 was written to perform this optimisation procedure. The results from this analysis are shown in the following list for a 3 and 4 stage vehicle, with and without solid fuelled upper stages, with the performance specifications mentioned above. The assumed payload for the simulations was 80 kg. This mass includes the satellite as well as the rocket guidance and control equipment 3 Stage Liquid Fuelled Rocket (Growth Factor=75.1) Stage Isp MR Prop DeltaV 1 260 0.85 3897 2668 2 320 0.85 924 3282 3 320 0.85 219 3282 Total 9232 m/s 2 Stages Liquid & 1 Stage Solid (Growth Factor=90.9) Stage Isp MR Prop DeltaV 1 260 0.85 4809 2759 2 320 0.85 1070 3396 3 290 0.85 238 3077 Total 9232 m/s 4 Stage Liquid Fuelled Rocket (Growth Factor=54.8) Stage Isp MR Prop DeltaV 1 260 0.85 2358 1966 2 320 0.85 867 2422 3 320 0.85 319 2422 4 320 0.85 117 2422 Total 9232 m/s 3 Stage Liquid & 1 Stage Solid (Growth Factor=61.6) Stage Isp MR Prop DeltaV 1 260 0.85 2691 2017 2 320 0.85 961 2483 3 320 0.85 343 2483 4 290 0.85 123 2249 Total 9232 m/s 2 Stage Liquid & 2 Stage Solid (Growth Factor=69.6) Stage Isp MR Prop DeltaV 1 260 0.85 3094 2069 2 320 0.85 1072 2547 3 290 0.85 371 2308 4 290 0.85 129 2308 Total 9232 m/s The propellant masses (kg) mentioned above are the optimum values given the mass ratios and specific impulses listed. As can be seen in these simulations, an increase in the number of stages results in a decrease in growth factor but increases the complexity of the system. Similarly, an increase in specific impulse decreases the growth factor. The stage mass ratio (MR) also has an enormous effect on the growth factor. As an example of this effect, the following table shows results for the 3 stage system with a third stage solid and a mass ratio of 0.8 for each stage. 2 Stages Liquid & 1 Stage Solid (Growth Factor=190.6) Stage Isp MR Prop DeltaV 1 260 0.8 10070 2759 2 320 0.8 1751 3396 3 290 0.8 305 3077 Total 9232 m/s These results reveal that a 6% reduction in stage mass ratio leads to a 109% increase in the growth factor. This example clearly shows why every gram counts on a satellite launch vehicle and that every effort must be made to trim off excess weight and still maintain safe operating margins. AUSROC III UPWARD COMPATABILITY One of the prime simplifications of the Ausroc IV vehicle is to use existing Ausroc III hardware wherever possible to reduce the amount of effort expended. Many of the systems being developed for the Ausroc III rocket are being done specifically for this future purpose. Ausroc III is, in fact, a test bed vehicle for the Ausroc IV satellite launcher mission. Technologies such as Guidance, Navigation and control (GN&C), composite structures and Thrust Vector Control (TVC) are being developed for Ausroc III and can be later refined for use in the Ausroc IV vehicle. For these reasons, it would be advantageous, in both time and resources, to utilize the Ausroc III propulsion modules in the Ausroc IV system with as little modification as is possible. The previous section derived the optimum stage ratios for various configurations. If Ausroc III 'modules' are used directly to form the Ausroc IV vehicle a sub-optimum staging ratio will result, since the propellant quantities in the Ausroc III units are fixed. For simplicity, the 3 stage configuration with the third stage being a custom solid fuelled motor has been chosen as the baseline for Ausroc IV. This choice was based on the similarity of the required propellant loadings with the loadings present in clustered Ausroc III modules. The staging configuration would consist of 4 modules in the first stage which are attached to a central second stage module atop of which is a custom third stage solid rocket motor as shown in figure 2. The stage optimisation program was modified to determine the velocity achievable with non-optimal propellant loadings. Using Ausroc III modules with 1200 kg of propellant and a mass ratio of 0.85, as previously stated, the first stage of 4 modules would contain a total of 4800 kg of propellant. The second stage of 1 module would contain 1200 kg of propellant. The third stage solid rocket motor propellant loading was optimised using the modified program and the above data. The final results obtained were: Ausroc IV Baseline Configuration 2 Stages Liquid & 1 Stage Solid (Growth Factor=92.9) Payload=80 kg Stage Isp MR Prop DeltaV 1 260 0.85 4800 2647 2 320 0.85 1200 3499 3 290 0.85 250 3139 Total 9285 m/s It can be seen from these results that the configuration achievable with the Ausroc III hardware and a custom solid propellant third stage is very close to the optimum configuration. The growth factor is only 2.2% larger than the optimum design. Figure 3 shows the graphical output of a computer program5 simulating the 3 stage launch vehicle utilsing Ausroc III hardware and a solid fuelled upper stage. The gross payload weight was found to be 50 kg. This includes the satellite as well as the guidance and control equipment. ADDITIONAL FLIGHT HARDWARE REQUIREMENTS Stage Separation Devices Each of the 4 first stage modules will be separated after 80 seconds. This operation is critical for a successful mission. The engine valves will be automatically shut off at a predetermined time to avoid a situation of uneven thrust levels. If the modules fail to separate simultaneously, the vehicle could enter a situation in which the control system becomes unstable. This would result in a terminated mission. A second critical area is that of 'stores clearance' safety. This requires that the disconnected burnt out modules do not become entrained in unsteady airflows and impact the remaining stages. This area can be analysed prior to flight with the use of wind tunnel facilities and scale models. The most common method of detaching stage modules that are not in-line is to use explosive bolts as shown in figure 4. The second/third stage separation would occur immediately after second stage burnout. The third stage, however, would not ignite until after a certain coast period in which the vehicle reaches the desired conditions for the final stage burn. Since the third stage is solid, a reduction in system complexity can be accomplished by eliminating the active thrust vector control elements and spinning both the third stage motor and attached satellite along its longitudinal axis to eliminate any thrust misalignments. The second and third stages are in-line and common separation methods include the use of explosive bolts or an explosive separation ring, as shown in figure 5, which doesn't produce any outgassing during operation. Both these methods are employed to physically detach the modules, but they do not provide the energy to separate the modules efficiently and safely. Small rocket motors are usually used to provide the separation force to clear the stage modules apart in a controlled and predictable manor. The satellite payload can either remain attached to the empty third stage casing in orbit or it can be separated at third stage burnout. Depending on the nature of the experimental payloads carried, the satellite may need to be de-spun prior to useful operation. This can be accomplished using a combination of small solid rocket motors and cold gas thrusters. If separation is required a combination of pyrotechnic or electro-mechanical disconnect latches and either spring push rods or solid rocket thrusters for separation may be used. Payload Fairing The Payload fairing encases the payload during the ascent through the atmosphere. At the end of second stage burn, the payload fairing would also be jettisoned, having completed its role. The nose cone and payload fairing on Ausroc IV will be subjected to a more severe thermal environment than that of Ausroc III and will require enhanced thermal protection. Control System The autopilot software is developed to control the vehicle within a certain operational envelope. This takes into account the dynamic and aerodynamic conditions existing within this envelope. When stage modules are jettisoned, both these conditions change and a new software algorith is required to control the vehicle. Thus, in the 3 stage Ausroc IV, there will be 3 autopilot control algorithms. The first stage has 4 engine modules which are fired simultaneously. To avoid an uncontrollable pitching or yawing moment from being produced, the thrust generated by each of these engines must be controlled to remain within fixed bounds. This is most easily achieved, with liquid propellant engines, by controlling the flow of propellants to the engine with active control valves. In an ablative motor, the throat dimension may increase due to erosion. This will result in a decrease in pressure and, hence, thrust. A chamber pressure sensor will be needed to provide feedback to the control system as to the extent of throat erosion. The Ausroc IV propulsion systems will need to be much more controllable than the inteded Ausroc III system. LAUNCH REQUIREMENTS Location The selection of a launch site is based on a combination of safety and performance considerations. Multi stage satellite launchers drop off stages once their propellant is consumed. These spent stages then impact the earth's surface at some further distance downrange. This impact zone must be located in a region where the probability of property damage and/or fatalities is very small. In the case of separated upper stages, they generally burn up on re-entry into the atmosphere and do not pose any safety problems. For the purpose of design, the Woomera Rocket Range5 in South Australia is being considered as the potential launch site for Ausroc IV. Woomera cannot be used for launches into the common easterly equatorial orbits since this would involve overflying the population centres on the east coast and present a safety problem. Northerly polar orbital launches can be undertaken from Woomera and have been in the past. The coastline of Papua New Guinea is approximately 2500 km to the north, as shown in figure 6, and the path across Australia is primarily desert and has a very low population density. Ground Use Facilities A generic launch site layout is shown in figure 7. This shows the requirements for a large and frequently used launching facility. Much of the hard infrastructure shown will not be required for the Ausroc IV system. The Ausroc III launch pad will need to be modified to handle the larger and heavier vehicle. The erector unit can be made to be integral with the transportation truck or a mobile crane could be brought in for the job. The rocket stages would be assembled and attached at a factory prior to being transported out to the site. Depending on the payload requirements, it to could be mated to the vehicle before being transported. If this is not possible an assembly building will be required on-site. Launcher area 5 at Woomera was used in the past as the launch site for the Black Knight and Black Arrow vehicles and, while most of the infrastructure has been scrapped, the control centre block house still exists and could be used after a basic refit. The main instrumentation building, which is several kilometers from the launch area, can be used as the central base of operations. This would include telemetry receiving equipment, impact prediction and range safety equipment as well as the transmit base for the flight termination system. A test and assemly building of sufficient size is also available on-site. Propellant & Hazardous Storage Propellants for use in the vehicle will be brought in by truck specifically for the trial and will not require special storage facitities to be constructed on-site. Modern vacuum insulated Lox tankers can store their cryogen for long periods of time without excessive boil-off, even in hot conditions as displayed at Woomera. The kerosene fuel has a high flash point which makes it relatively safe and easy to store in most cool dry locations such as one of the magazines spread around the Woomera site area. The third stage solid motor will need to be stored in a safe magazine area as well. This fact alone may require the final assembly of the vehicle to be done on-site in the hazardous test shop facilities. The complete Ausroc IV will contain several pyrotechnic devices which are not installed into the vehicle until the final preparation. These devices will include pyrotechnic motor igniters, explosive bolt detonators, squibbs and explosive cord. These will also need to be stored in a safe magazine until required. Meteorological Facilities A meteorological station is located in the Woomera tech area. Immediately prior to a launch, a series of weather balloons are released to determine the wind conditions at various altitudes. This data will determine whether or not it is safe to proceed with a launch. Radar & Tracking Equipment Woomera currently has 2 french Adour Radars and while they are probably adequate for use during the Ausroc III flight trial, they would definately be inadequate for the Ausroc IV satellite launch mission. The FPS-16 radars used during the satellite launching activities of the Joint Project days have long since been removed. The purchase and commisioning of a new radar system capable of being used for orbital shots from Woomera would cost many millions of dollars. The radars are used, primarily, to provide an independant position and velocity reference for range safety purposes. The Global Positioning System (GPS) satellites could be used as an alternative to the radar, to provide the position and velocity data. New GPS systems on the market today can provide accurate velocity and position data over a large speed range. These systems have not yet been used to replace radars but have been used in conjunction with them. By the time Ausroc IV is ready for launch, the GPS system may be proven reliable enough to use for range safety and would be substatially cheaper than a Radar. Telemetry Equipment A nominal 1.4 GHz transmitter on board the rocket will be transmitting data to a ground station. This data will include position information as well as valuable flight status and environmental data. Environmental data includes aspects such as vibrations, temperatures, pressures etc. Several steerable dish antennas will be required, for redundancy, to track and recieve the data stream from the vehicle. A second powerful and highly reliable transmitter will be based on the ground and be used as part of the flight termination system. In a typical system, recievers on-board the launcher will continuously receive a 'no fire' signal from this ground transmitter. In the case of a malfunction, which causes the vehicle to cross the safe range boundaries, the 'no fire' signal is cut off and the flight terminates. This also implies that a transmitter or reciever failure will cause an unwarranted termination. Several other variants of this system are currently used. The type to be implemented in Ausroc's III & IV has not yet been decided. Environmental Concerns The first 2 stages of the preliminary Ausroc IV design use liquid oxygen and kerosene as propellants. This combination produces relatively safe and benign exhaust products consisting mainly of water, carbon dioxide and carbon monoxide. The liquid oxygen will readily disperse itself into the atmosphere if spilled but the kerosene will soak into the ground and only disperse over a long time period. The anticipated third stage will use a solid propellant propulsion system. Common high performance propellants consist of ammonium perchlorate, aluminium powder and butadiene rubber. A major constituent of the exhaust products is hydrochloric acid. An analysis of the effects of this product, when dispersed at high altitude (>100 km), has not yet been undertaken. LEGAL IMPLICATIONS In order to place a satellite into orbit a number of national and international issues6 must be addressed and solved before the launch occurs. A number of United Nations Treaties have been formulated regarding satellite launching. Three of the more important of these treaties are listed below: 1. The Outer Space Treaty The outer space treaty is generally regarded as creating a general framework for activities in outer space, with more detailed provisions left to be covered in subsequent conventions and agreements. 2. The Liability Convention The liability convention makes the launching state or entity liable for any damage or fatalities arising from the operations of that state or entity. In the case of a launch by a private group, such as the AUSROC Projects Group, ultimate liability will rest with the Australian Government. Thus the government will, no doubt, request indemnification against the loss or damage caused by the launch vehicle. Being utlimately liable for the launch, the government may also impose severe restrictions on the activities undertaken during the manufacture, test and launch of the Ausroc IV vehicle. 3. The Registration Convention The registration convention has as its objective, the recording and making available basic information concerning space objects. It does so by requiring launching states to record information for their own records and for those of the United Nations. The intended flight path of Ausroc IV passes over Papua New Guinea. Permission for such a crossing of national boundaries would need to be forthcoming from their government. This liason would need to be undertaken by the Australian government, hence another requirement for their intervention in the activity. INSURANCE ISSUES It will be a requirement for the Ausroc Projects Group to obtain liability insurance to indemnify the Australian Government against claims from both foreign nationals, according to the Liability convention, and nationals according to common law. The amount of liability insurance required is not yet known. Most satellites launched these days are insured against launch failures and depending on the track record of the launcher used, the premiums for the insurance usually range from 15-20% of the launch contract value. Due to a nil track record and its amateur design nature, the Ausroc IV vehicle would be unable to get payload or reflight insurance. CONCLUSIONS A general discussion of the possibilities of developing a satellite launcher vehicle using Ausroc III hardware has been presented. It has been shown that clusters of Ausroc III modules and solid rocket motors have the required amount of energy to place a micro satellite into orbit. It was also shown that with modifications, the Woomera Rocket Range could be used as the launch site. A brief mention of the legal and insurance issues was also presented for completeness. The following table lists a number of projects, derived from this discussion, that will need to be undertaken in addition to those listed for Ausroc III. AUSROC IV SUBPROJECT LISTING Propulsion System (PS) PS:9 High Altitude Nozzle PS:10 Fourth Stage Engine PS:11 Fourth Stage Motor Gimbal System Composite Structures (CS) CS:4 Fourth Stage Propellant Tanks CS:5 Satellite Shroud / Nose Separation Systems (SS) SS:1 Stage Separation System (1-2) SS:2 Stage Separation System (3) SS:3 Shroud Separation System SS:4 Satellite Deployment System Systems Analysis (SA) SA:5 Dynamic Analysis (IV) SA:6 Aerodynamic Analysis (IV) SA:7 Staging Analysis SA:8 Orbital Trajectory Determination SA:9 AutoPilot Control System (IV) General Issues (GI) GI:1 Legal Implications GI:2 Public Relations Satellite Project (SP) REFERENCES No. Author Title 1. Blair M "AUSROC III Propulsion System" Ausroc Conference 1991 2. Brown K. "Ballistic Missile & Space Vehicle Seifert H. Systems" Wiley & Sons 1961 3. Sutton G. "Rocket Propulsion Elements" Wiley & Sons 1986 4. Blair M. "Stage Optimisation Simulator" Computer Program (c) 1991 5. Cheers A. "A Spherical Earth Model Particle Trajectory Simulator Utilizing a 4th Order Runge-Kutta Method" Computer Program (c) Ardebil 1991 6 D.S.T.O. "General Guidelines and Information for Defence Activities in the Woomera Prohibited Area" July 1989 7. I.E.Aust "Cape York International Spaceport Scoping Study - Legal Issues Report" I.E.Aust July 1987