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Israel and South Africa collaborated closely in rocket technology in the 1970's and 1980's. South Africa provided Israel with the uranium and test facilities it needed for its strategic weapons programmes. In exchange Israel provided aerospace technology. This included the capability of building the ten-tonne solid propellant rocket motors designed for the Israeli Jericho-2 missile. These motors were the basis of two space launchers for an indigenous 'R5b' space programme. It seems that South Africa also planned to use these motors in a series of missiles to provide a nuclear deterrent. Two shorter-range missiles (the RSA-1 and RSA-2) were intended for use on Cuban or Warsaw Pact troop concentrations should a massed attack be made from an adjacent country. The RSA-4 ICBM was also in long-term development, possibly to deter the United States or Soviet Union from sponsoring such an attack in the first place. The original intended payload for the missiles was said to be the uranium gun-type atomic bombs developed in South Africa between 1971 and 1989. Seven of these weapons were built, each with a mass of about one tonne, a diameter of 65 centimetres and a length of 1.8 meters. Each device contained 55 kilograms of highly enriched uranium, producing a fission yield of 10-18 kilotons. Five of the weapons were configured as air-launched bombs, but were said to be adaptable to missile launch. It was planned during the 1990's to lighten and modernise these warheads, and use tritium-boosting to increase the yield to 100 Kt. A missile using the original devices would have to be provided with a heat shield, implying a total warhead mass of around 1500 kg. This would not have permitted the RSA-4 to reach intercontinental range. Presumably the boosted, modernised warheads, that were to have been completed by 2000, would have been under 700 kg and allowed the missile to reach New York or Moscow. Substantial facilities for assembly, test, and launch of the rockets were built at the Overberg Test Range at the tip of Africa. Overberg was also used for Israeli Jericho-2 test flights. As a cover for and supplement to the missile development, the R5b indigenous space programme was funded. This would use the RSA-3 and RSA-4 launchers. Four South African space rockets were built. Three were launched into suborbital trajectories in the late 1980's in support of development of the RSA-3 launched Greensat Orbital Management System (for commercial satellite applications of vehicle tracking and regional planning). The range was also used for aerospace and system testing for British, Swedish and Czech programmes. Following the decision in 1989 to cancel the nuclear weapons program, the missile programs were allowed to continue until 1992, when military funding ended and South Africa officially ended its missile collaboration with Israel. All ballistic missile work was stopped by mid-1993. In order to join the Missile Technology Control Regime the government had to allow American supervision of the destruction of key facilities applicable to both the long range missile and the space launch programmes. The RSA-3 and RSA-4 space launchers were therefore cancelled in 1994. Prime Contractor Houwteq had to dismantle its existing RSA rocket components, and retrieve and sequester technical data from its subcontractors. Propellant manufacturer Somchem eliminated the RSA solid propellants and rocket casings that remained in stock. Denel filled in its large engine casting pits at Somerset West and demolished its large-scale X-ray inspection equipment. The Hangklip static motor test facility at Rooi Els was converted to a nature reserve. The Overberg Test Range was allowed to remain for use by 'potential foreign partners'. Following these measures, South Africa joined the Missile Technology Control Regime in September 1995. Houwteq's staff at Overberg grew to a peak of 500 in 1992, before the cancellations began. By 1997 the staff was reduced to 28. Houwteq's Ian Farr continued to market the Overberg facility for commercial launchers until at least 1997. Nothing came of these efforts, and it seems that the book was closed on further indigenous African space activities. Since much remains undisclosed about the Israeli Jericho missiles and Shavit / Next space launchers, the material on the South African rockets provides some insights into Israeli programmes. The RSA-2 clearly corresponds closely to the Jericho-2, and the RSA-3 to the Shavit launcher. It is interesting that there has been no mention of an Israeli counterpart to the Peacekeeper-class first stage motor of the RSA-4. This may represent a 'reserve' Israeli capability to upgrade the Jericho-2 to ICBM range that has never been made public. However there have been reports of Israeli development of a MIRV capability for its missiles. The post-boost warhead dispenser for such a capability could correspond to the RSA-4 fourth stage. Manufacturer: IAI. Launches: 3. Success Rate: 100.00%. First Launch Date: 1989-06-01. Last Launch Date: 1990-11-19. Launch data is: incomplete.
Version: RSA-1.
It is conjectured that this designation was assigned to an intermediate range single-stage ballistic missile consisting of the first stage of the RSA-3. Purported mission was to strike Cuban military concentrations from mobile launchers on South African territory. The rocket motor closely followed the design of the Israeli Jericho-2 first stage.
to a: 1100 km range trajectory. Liftoff Thrust: 412.700 kN (92,779 lbf). Total Mass: 12,000 kg (26,000 lb). Core Diameter: 1.30 m (4.20 ft). Total Length: 8.00 m (26.20 ft). Standard warhead mass: 1,500 kg (3,300 lb).
Version: RSA-2.
It is conjectured that this designation was assigned to an intermediate range ballistic missile consisting of the first and second stages of the RSA-3. Probably very similar to, or a licensed copy of the Israeli Jericho-2 missile. A third stage apogee kick motor was added to produce the RSA-3 space launcher.
to a: 1900 km range trajectory. Liftoff Thrust: 412.700 kN (92,779 lbf). Total Mass: 23,000 kg (50,000 lb). Core Diameter: 1.30 m (4.20 ft). Total Length: 15.00 m (49.00 ft). Standard warhead mass: 1,500 kg (3,300 lb).
Version: RSA-3.
The RSA-3 satellite launcher began development as an IRBM in the 1980's because of the perceived Soviet threat and isolation of South Africa. It was developed with the assistance of Israel and was believed to be essentially identical to the Israeli Jericho missile/Shavit launch vehicle. The objective of the satellite launcher was to place a small surveillance satellite of 330 kg mass into a 41 degree, 212 x 460 km orbit around the earth. Development continued even after South African renunciation of its nuclear weapons. However the launcher was found not to be viable commercially and so was cancelled in mid-1994. The RSA-3 was developed by the Houwteq organization at Grabouw, 30 km east of Cape Town. The Overberg Test Range near Bredasdorp, 200 km east of Cape Town, was used for test flights. The engine test facility was at Rooi Els. At the peak of development in 1992 50 - 70 companies in the public and private sector were involved, employing 1300 -1500 people. As in the Shavit, the first and second stages used the same rocket motor loaded with 9 metric tons of propellant. The first stage used vanes in the exhaust for steering during the first 16 to 20 seconds of flight, after which the fins at the base of the vehicle provided aerodynamic control. The second stage had a higher expansion nozzle and may have been equipped with TVC for steering. Atop the second stage was a guidance / orientation / spin-up bus for the third stage and payload. Total mass of this bus and the payload shroud was 583 kg. After second stage burnout, the upper stage package entered a 148 second ballistic coast. A sideways trajectory deflection was made and the shroud was jettisoned. Then the third stage and payload were spun up, following by separation of the bus. The spin-stabilized third stage then made the 4,555 m/s burn to place the payload into orbit. The third stage was evidently similar to a 5 metric ton thrust spherical motor displayed by the Israelis for their Shavit launch vehicle. The composite payload fairing for the RSA-3 was 4.5 m long, 1.3 m in diameter, and had a mass of 57 kg. The solar array for the satellite had a mass of less than 7 kg and with three panels could supply 295 W of power. The fairing could resist temperatures of up to + 100 degrees C during ascent and the thermal satellite blanket insulated the payload from temperatures ranging from -80 degrees C to + 100 degrees C. As an ICBM, it is estimated that the three-stage version of the RSA-3 could have delivered a 340 kg warhead on Washington DC or a 400 kg warhead on Moscow. However such lightweight warheads were beyond declared South African technology. Therefore the RSA-3 was most likely purely a space launch adaptation of the RSA-2 IRBM, with the Peacekeeper-class RSA-4 fulfilling the ICBM role. The RSA-3 and its mobile erector-launcher were in an advanced stage of test at the time the program was cancelled. It is not known what happened to the hardware that was built. Warheads of the size and type required for use on the RSA-3 were not in the inventory according to the declarations made by South Africa at the time of its nuclear disarmament and signature of the nuclear non-proliferation treaty. The RSA-4, was in the design stage. The following is the detailed launch trajectory of the RSA-3:
Launches: 3. First Launch Date: 1989-06-01. Last Launch Date: 1990-11-19. LEO Payload: 330 kg (720 lb). to: 210 km Orbit. at: 41.00 degrees. Apogee: 300 km (180 mi). Associated Spacecraft: Greensat. Liftoff Thrust: 412.700 kN (92,779 lbf). Total Mass: 23,630 kg (52,090 lb). Core Diameter: 1.30 m (4.20 ft). Total Length: 17.65 m (57.90 ft).
RSA-4.
The RSA-4 ICBM / satellite launcher was a planned follow-on to the RSA-3. A large new first stage optimised the vehicle and more than doubled the payload in comparison to the RSA-3. It is not known if the project reached the point of testing of the large motor, which was equivalent to the US Peacekeeper first stage. The second and third stages were essentially those of the RSA-3. The fourth stage was clearly adapted from an ICBM MIRV post-boost bus platform. As an ICBM or orbital nuclear system the RSA-4 would have been capable of delivering a single 700 kg warhead anywhere on earth. Work on the RSA-4 was cancelled in 1994. An attempt was made by Houwteq to market the RSA-4 as a launcher for MEO earth satellite constellations. It was not to be...but the sales brochure from that effort survives and is reproduced below.
Introduction to RSA-4 Launch VehicleHouwteq The South African Space Industry is spearheaded by Houwteq. Houwteq is the prime contractor in the space industry for space systems and services and is supported by twenty local subcontractors. Most of the subcontractors had established proven capabilities in the high technology field before joining the space industry. Houwteq is the systems house responsible for the space system design, assembly, integration, launch preparation and execution and project management of space activities. The company is located near Grabouw in the Cape Province. Houwteq's subcontractors are responsible for the design, development,, qualification, production, as well as technical and logistic support, at configuration item level. Houwteq is offering a comprehensive launch service with its RSA-4 Launch Vehicle from the Overberg Test Range situated in South Africa. Chapter 1 1.1 Purpose of Document This document is intended as an introduction to the RSA-4 launch vehicle and the launch service offered to prospective clients for placing small to medium sized payloads into low earth orbit (LEO). The document covers the following aspects;
Houwteq offers a comprehensive service for LEO launches including the launch vehicle, the launch facility and associated services. The RSA-4 launch vehicle comprises three solid propellant boost stages and a hydrazine powered fourth stage for accurate orbit injection and positioning, It can lift a satellite with a mass of 550 kg into a circular orbit at a height of 1400 km and a inclination of 55 degrees. Provision is made for a payload volume of 10.4 m3 with a maximum diameter of 2.2 m. II is possible to launch two satellites into different orbits which are in the same orbital plane. The launch vehicle is described in more detail in Chapter 2. Launches are conducted from the newly established facility at the Overberg Test Range at the southernmost tip of Africa on the south-eastern coast of the western Cape at Lat 34 deg 35 min S and Long 20 deg 19 min E. The facility (OTR), which extends over a total area of 43,000 hectares, is situated close to the villages of Waenhuiskrans and Bredasdorp. Cape Town, one of the major cities in South Africa, is located approximately 200 km from OTR and has a commercial airport capable of handling large airliners. A good quality highway exists between Cape Town and OTR. A modern air base adjacent to OTR can accommodate all types of aircraft, The use of this facility could be negotiated for specific transport arrangement if required. Chapter 2 2.1 System Description Figure 2.1 shows the overall dimensions and layout of the RSA-4 launch vehicle. It comprises four stages. The first three have solid propellant motors to lift the satellite into orbit. Orbit raising is then performed by means of the fourth hydrazine-propelled stage, which is also used to make fine orbital adjustments to place the satellite accurately into the required orbit. The masses for the stages are respectively 66 tonne for the first stage, 10 tonne for the second, 3 tonne for the third, and approximately 300 kg for the final stage, which gives an all-up mass of 80 tonne for the complete system. The overall length is 23.5 m with a diameter of 2.4 m. Provision is made for a payload with a maximum diameter of 2.2 m and a maximum height of 3.74 m. The length of the payload volume at maximum diameter is 1.5 m (Figure 5.4). Two satellites, which fit into the available volume, can be placed into the same orbital plane. 2.2 First Stage The first stage (Figure 2.2) is propelled by means of a solid propellant motor weighing 62.6 tonne, of which 58 tonne is propellant. It delivers an impulse of 139,000 kNs at sea level and burns for 73 s with an average thrust of just under 200 tonne. The expansion ratio of the carbon-phenolic expansion cone is 14. A graphite throat insert is used. The composite casing is made of Kevlar and covered with cork for thermal insulation. The base and the interstage structures are made of aluminium 2219. Control is done by means of LITVC, as well as air vanes manufactured from honeycomb material. Also situated at the base is a PCM unit for telemetry acquisition and an S-band telemetry transmitter. Power is supplied by 28 volt silver-zinc triggered batteries. In the interstage section on top of the motor is the pyrotechnic activation unit for motor ignition, separation from the launch platform, and detonation of the cutting cords for motor destruction. 2.3 Second Stage Stage 2 (Figure 2.3) is very similar to stage 1 with following differences: The motor with 9 tonne solid propellant burns for 52 s and delivers an impulse of 24,500 kNs with an average thrust of 40 tonne. Control is done by injecting strontium perchlorate into the motor flame. It is stored in a torus around the nozzle throat and pressurized by means of helium. The second stage also houses the receivers for destruct in case of malfunction. The destruct channels are redundant and destruct is initiated on receipt of a destruct signal or in case of loss of the carrier wave. An aluminium aerodynamic sleeve is utilized to provide a constant external diameter between the lower interstage of stage 1 and the avionic section of stage 3. The sleeve is jettisoned after burn-out of stage 1. 2.4 Third Stage Stage 3 (Figure 2.4) houses the autopilot for digital implementation of the guidance and control algorithms, onboard safety implementation, initiation of discrete events like stage separation, and power switching. Communication is via a 1553 bus. Navigation is performed by a strapdown platform and GPS. Power is supplied from 28V triggered batteries. Also housed in the third stage is the equipment for tracking and monitoring the launch vehicle, i.e. telemetry, television for monitoring stage separation, a Doppler beacon for measuring velocity and a radar transponder for measuring velocity and range. The structure is made of aluminium. The third stage motor weighs 2 tonne, with 1,9 tonne solid propellant in a titanium casing and a nozzle with an expansion ratio of 60. It burns for 92s and delivers a specific impulse of 292 s. 2.5 Fourth Stage The purpose of the fourth stage (Figure 2.5) is to raise the orbit and to make small orbital adjustments. For this purpose it is equipped with a reactive control system comprising four 50 liter hydrazine tanks, filled to the right level for the particular mission and pressurized by helium, and ten thrusters. Four 200 N thrusters are used for roll control, four 25 N thrusters for pitch and yaw control and two 200 N thrusters for orbit raising. Figure 2.5 shows the layout of the reactive control system. Navigation is done by means of GPS with a CA code receiver. Rate gyros and accelerometers measure the orientation and velocity increments needed by the control computer for control loop implementation. The fourth stage provides status information to the satellite during launch on the progression of events, and relays satellite telemetry during launch preparation and launch in the S-band via a stripline antenna to the ground. Power is supplied by 28 V lithium batteries. The structure is made of carbon composite and honeycomb material. The front adapter to the satellite makes provision for cryogenic cooling supply, two electrical connectors for telemetry relay and power supply, as well as for mechanical attachment and pyrotechnic separation of the satellite. The fairing which protects the satellite is made of a honeycomb composite material and is jettisoned sideways in two halves after stage 2 burn-out. Chapter 3 3.1 Introduction The RSA-4 launch vehicle is used to lift small to medium sized satellites into elliptical or circular low earth orbits at inclinations between 37 deg and 90 deg and in sun synchronous. Two satellites can be launched simultaneously, if necessary into different orbits, which must however be in the same orbital plane. Since many different missions are possible, the performance figures and flight profiles are presented for selected cases. 3.2 Lifting Capability The payload mass which can be placed into a circular orbit for different orbital orbits is given in Figure 3.1 for 55 deg and 90 deg inclinations. Typical values are
3.3 Injection Accuracy The 2-sigma values for the accuracy of injection into a circular orbit are as follows:
As an example of typical values for the main trajectory parameters, their time variation is presented in Figures 3.2 to 3.6 for a 550 kg spacecraft launched at a 55 deg inclination. 3.5 Spacecraft Deployment In the standard mode the spacecraft will be spun up to >1 rev/s and will be separated with a relative velocity > 0.5 m/s with an angle between the spacecraft longitudinal axis and the velocity vector <5 deg. Requests for other deployment conditions will be handled on an individual basis. 3.6 Launch Sequence The launch sequence is schematically illustrated in Figure 3.7. The trajectory depends on the particular mission. As an example the flight parameters and sequence of events are presented for the injection of two satellites into a 1400 km circular orbit. During the first few seconds after ignition the launcher is rotated from the vertical and flies a gravity turn during the first powered stage. At a height of 40 km, 55 km downrange, the first stage separates, the sleeve is jettisoned, and the second stage motor is ignited. It burns out at a height of 97 km, 230 km downrange, and is jettisoned 25 s later. Several events are scheduled during the following cruise phase before the third stage motor is ignited. The stage is despun, the satellite protective fairing is jettisoned and then the stage is aligned in the correct orientation for motor firing. Now the stage is spun up and the avionic section separated. At a height of 240 km, 1050 km downrange, the motor is ignited and burns for 92 s. Nutation control is performed during the propelled phase by firing the fourth stage 25 N thrusters intermittently to limit the amplitude of any coning motion which may develop due to thrust or mass asymmetries. The third stage injects the satellite into an elliptical LEO with a perigee of 250 km and burns out 1600 km downrange from the launch site. A period of three minutes is allowed to ensure that the residual thrust drops off to zero before the motor is separated. The satellite plus third stage is now in an elliptical transition orbit with the nominal apogee at the circular orbit height. Immediately after separation a velocity correction is done if the state vector at burnout deviates from the nominal. The velocity increment required at apogee to circularize the orbit and the time at which it needs to be given is calculated and after half a revolution the thrusters are fired to make the orbit circular. After one revelation the fourth stage, still carrying the two satellites, passes within view of the Overberg ground station. The fourth stage is despun, if required, the two satellites separated pyrotechnically and deployed by means of springs. Half a revolution later when the separation between the fourth stage and the satellites is large enough, the thrusters are fired for the last time to deorbit the stage. If the satellites are given a velocity increment of 0.5 m/s relative to one another and relative to the fourth stage, the two satellites are separated from one another by 5 km along the trajectory and just over 2 km in height after half a revolution and the fourth stage is separated from the satellite nearest to it by the same margin (Figure 3.8). The fourth stage starts off in front of the satellites at separation, but goes into a slightly higher orbit and is in fact overtaken by the satellites. The remaining problem is to synchronize the phase angle of the satellite in its orbital plane with that of its allocated slot in the constellation. In the launch sequence outlined here, it is assumed that this will he done by the satellite. If the maneuver is done over a few days, it requires little hydrazine. Should it be a requirement, this function could be performed using the fourth stage. In this case the feasibility will be investigated with the client, the limiting factor being the power required by the spacecraft up to deployment. It will be possible to inject the two spacecraft into different elliptical or circular orbits provided that they are in the same orbital plane and that the energy requirements are within bounds. Provision is made to give a payload of 550 kg a velocity increment of 300 m/s with the fourth stage. 3.7 Launch Execution If the spacecraft is to be launched into an orbit with a specified nodal angle and if the direction of motion in the orbit is prescribed, then energy considerations and the range safety volume restrict the opportunity for launching to a window of fifteen minutes around a particular time once a day. It will be possible to monitor the launch from the Overberg ground station up to burnout of the third stage motor. This is the most intensive part of the launch. Separation of the third stage motor and the velocity correction directly afterwards take place after it has disappeared over the horizon During development flight test these events will be monitored from a downrange telemetry station onboard a ship. Orbit determination necessary for the calculation of the velocity increments needed for orbit raising will be done by the fourth stage using GPS, with the ground station only used for verification. Circularization of the orbit will be performed after half a revolution. It will preferable to involve a ground station in the Northern Hemisphere to monitor this event. The ground stations at Hawaii, Fairbanks and Guam have line-of-sight communication at this point (Figure 3.11). Figures 3.9 and 3.10 show the ground track of the satellite for a 55 deg inclination. At an orbital height of 1400 km it will pass within view of the Overberg ground station for the first six revolution. LEO Payload: 770 kg (1,690 lb). to: 400 km Orbit. at: 55.00 degrees. Payload: 550 kg (1,210 lb). to a: 1400 km, 55 degree circular orbital trajectory. Liftoff Thrust: 2,000.000 kN (449,600 lbf). Total Mass: 80,000 kg (176,000 lb). Core Diameter: 2.40 m (7.80 ft). Total Length: 23.50 m (77.00 ft).
RSA Chronology 1989 June 1 - Overberg -. RSA-3-d 1 Test mission Agency: Armsco. Apogee: 100 km (60 mi). 1989 July 6 - Overberg -. RSA-3 2 Test mission Agency: Armsco. Apogee: 300 km (180 mi). 1990 November 19 - Overberg -. RSA-3 3 Test mission Agency: Armsco. Apogee: 300 km (180 mi). 1994 June - RSA-3 / RSA-4 South African satellite launchers cancelled The RSA-3 satellite launcher began development as an IRBM in the 1980's. It was developed with the assistance of Israel. The satellite launcher was found not to be viable commercially and so was cancelled in mid-1994. The Overberg Test Range near Bredasdorp, 200 km east of Cape Town, was used for test flights. Bibliography:
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