This article begins with a brief history of the Yuzhnoye Design Office, one of the leading design bureaus in the Soviet Union and Ukraine. The Yuzhnoye Design Office designed numerous strategic missile systems, launch vehicles, and spacecraft (1). One interesting aspect of this history involves the conversion of missiles into space launch vehicles by a team of developers led by the Yuzhnoye Design Office, well in advance of the officially announced USSR conversion effort.
Brief History
The M.K. Yangel’ Yuzhnoye State Design Office (initially known as Special Design Bureau 586, abbreviated OKB-586) was established on 10 April 1954 in Dniepropetrovsk, a city on the banks of the Dniepr River in central Ukraine. Mikhail Kuz’mich Yangel’ was named General Designer. Before this, he had been Director of the Scientific Research Institute 88 (NII-88) in Podlipki (now known as Korolev), a city near Moscow. In 1946, NII-88 became the USSR’s main center for missile development.
A group of missile specialists from the General Designer’s Department of All-Union State Plant 586 (the former Dniepropetrovsk Motor Vehicle Plant, now the State Enterprise Production Association Yuzhnyi Machine-Building Plant) formed the core of the newly established OKB-586. In 1951, Plant 586 started mass production of the R-1, R-2, and R-5 missiles developed by NII-88 and the NII-88 Special Design Bureau 1 (OKB-1), which was headed by General Designer Sergei Pavlovich Korolev. OKB-586 and Plant 586 joined forces to establish a missile design and production center where everything was under one roof. A nationwide network of developers and manufacturers for the components, systems, intermediate products, and hardware specified for use in OKB-586 missile production and development activities was also set up along with OKB-586 itself.
The Soviet Government’s intention was to have this newly established system of production and development facilities headed by OKB-586 become (relative to the existing system) a stronger, more productive scientific production cooperative for developing future USSR strategic missiles. The intent was to promote missile development and also to increase substantially the military effectiveness of the missiles themselves. Despite the success of NII-88 in developing the R-1, R-2, and R-5 missiles, high-level military and government personnel understood that these missiles and any others that could be developed using the same principles could not meet future strategic requirements. Each of these missiles had substantial deficiencies that prevented them from being used in real combat conditions.
The new team would eventually eliminate these deficiencies by developing future missiles within a completely new conceptual framework. This conceptual framework was based on several principles developed by NII-88 personnel in 1952 under the leadership of M.K. Yangel’. These were the most important of the principles:
* development of a series of missiles having ranges consistent with strategic requirements;
* deployment of these missiles in hardened silo launchers constructed for concealment from the enemy;
* development of these missiles to use high-boiling-point propellants that could be stored long term and avoid using the low-boiling-point propellants previously used that had poor storage qualities;
* the use of autonomous onboard control systems protected against enemy electronic countermeasures and avoiding subsystems that might be vulnerable to noise.
The design concepts based on the high-boiling-point AK-27I and TM-185 rocket propellants developed by OKB-1 and Plant 586 General Designer’s Department personnel for the R-11 and R-12 missiles in 1952 confirmed that these principles were valid and sensible. The establishment of OKB-586 and the appointment of M.K. Yangel’ as its General Designer were evidence that the Government supported this new conceptual framework for developing new missile systems. This conceptual framework guided OKB-586 in all of its activities and was continually developed and enhanced, as various new development projects were implemented. In the process, OKB-586′s main mission became developing strategic missile systems capable of inflicting a highly effective second strike against any aggressor if the Soviet Union were the target of a nuclear attack.
Despite the organizational issues that arose during the establishment of OKB-586, a lack of essential equipment and experienced personnel, insufficient research on high-boiling-point propellants, delays in developing the rocket engine and autonomous onboard control system, and delays in constructing the silo launcher, the first medium-range missile system, the 8K63, was placed into service less than 5 years after OKB-586 was established. Two additional missile systems, the 8K65 medium-range missile and the 8K64 intercontinental ballistic missile, were placed into service at 2-year intervals. For some time, these missiles served as the primary weaponry in the USSR Strategic Missile Forces arsenal.
Within this brief period, under the leadership of M.K. Yangel’, OKB-586 had become the USSR’s leading design bureau for developing the strategic missile systems that were most important to the USSR’s defense capability.
From 1954 to 1991, a total of 29 strategic missile systems were developed by OKB-586 (from 1966 on, known as the Yuzhnoye Design Office) and the team of engineers under the leadership of Academicians M.K. Yangel and V.F. Utkin. Thirteen of these systems were accepted for military service by the Strategic Missile Forces and became the backbone of their forces (Fig. 1).
Some of these systems have no peers in missile technology. Examples include the 8K69, 15A14, 15A18, the 15A18M, the 15Zh60 fixed solid propellant missile system 15Zh60, and the 15Zh61 rail-mobile missile system, all of which played an important role in enabling the Soviet Union to reach strategic parity with the United States and in negotiations of the Strategic Arms Limitations Treaties between the Soviet Union and the United States.
The four generations of missiles accepted into Missile Forces armaments are shown in Fig. 2, together with the dates when they were placed into service. Each of these missiles had a unique purpose and unique characteristics, and there were substantial differences in specifications. They were all developed at different times, and each embodies the scientific, technical, and economic capabilities of the country at the time they were developed. In spite of these differences, however, it is possible to identify some trends typical of all four generations of missiles. The service life of the missiles tends to increase, the capabilities of the missile-defense countermeasures improve, the total energy output increases, the range increases and the operational specifications of the missiles all improve from one generation to the next.
Figure 1. General Designers: S.P. Korolev (1907-1966), M.K. Yangel’ (1911-1971), V.F. Utkin (1923-2000), S.N. Konyukhov (1937-).
Figure 2. Military missiles developed by the Yuzhnoye State Design Office.
The missiles of the fourth generation possess the highest military effectiveness. Whether in silos or during active flight, these missiles are able to preserve their performance in the face of any countermeasures. They are equipped with a very effective multifunctional antimissile system, which in combination with high survivability during active flight allows them to overcome with a high probability of success even a future adversary missile defense system. As of now the four types of the most highly developed military missiles 15A18, 15A18M, 15Zh60 and 15Zh61 are still deployed in the Russian Federation.
In addition to the missiles described, from 1957 on, OKB-586 also developed a variety of space launch vehicles. The most predominant idea involved developing a missile-based launch vehicle. This approach would lead to a substantial reduction in one-time and recurring costs, as well as a reduction in launch-vehicle development time due to the reduced amount of design and development work required and the ability to use the existing manufacturing infrastructure, the existing basic missile components available at the various manufacturing plants, and existing ground-based launch facilities.
This idea was implemented via the development of several launch vehicles based on the 8K63, 8K64, 8K66, 8K67, 8K68, 8K69, and 15A18 missiles; five of these launch vehicles—the 11K63 (Kosmos), 11K65 (Kosmos-2), 11K69 (Tsiklon-2), 11K68 (Tsiklon-3), and Dnepr—were in actual use. These five launch vehicles are the subject of this paper. Two additional launch vehicles, the Zenit-2 and Zenit-3SL, were developed without using military prototypes. By contrast with the missile-based launch vehicles, all stages of these launch vehicles used liquid oxygen as the oxidizer and RG-1 kerosene as the fuel. A modified Zenit-2 first stage was used as a module in the Energia vehicle (Fig. 3).
The space ambitions of the Yuzhnoye Design Office were also embodied in the design and successful development of Module E, the lunar-lander portion of the lunar spacecraft developed as part of the lunar program. Since 1960, OKB-586 and the Yuzhnoye Design Office produced a series of research, commercial, applied scientific, and military spacecraft (more than 70 different types). This count includes the following Kosmos and Interkosmos spacecraft: AUOS, Okean, Taifun, Tselina, etc. Approximately 400 spacecraft designed by the Yuzhnoye Design Office and manufactured by the Yuzhnoye Machine-Building Plant have been launched into space, many of them aboard launch vehicles designed in-house (Fig. 4).
All design work at Yuzhnoye Design Office was performed under the direction of General Designer M.K. Yangel’ before October 1971 and under General Designer V.F. Utkin between October 1971 and November 1990. M.K. Yangel’ served as General Designer for the development of first-, second-, and third-generation missiles, but did not live to see the 15A14, 15A15, and 15Zh60 missiles certified for military use; however, these latter missiles were also based on his ideas, which he had to defend at many levels, up to and including the USSR Defense Council. M.K. Yangel’ also attached a great deal of importance to space research. The 11K63 (Kosmos), 11K69 (Tsiklon-2), and 11K68 (Tsiklon-3) launch vehicles and Module E of the lunar spacecraft were developed under his leadership. At his initiative, a spacecraft design bureau was established at OKB-586 in 1960, spacecraft were produced at the Yuzhnoye Machine-Building Plant. Approximately three dozen types of spacecraft and several hundred spacecraft were launched under his leadership.
As Yuzhnoye Design Bureau Chief Designer, Vladimir Fedorovich Utkin had enormous influence on the development and delivery of the third- and fourth-generation missiles, as well as the Tsiklon-3 and Zenit-2 launch vehicles. The Energia launch-vehicle module unit and approximately 40 other types of spacecraft were developed under his leadership. The highly efficient and reliable Zenit-2 launch vehicle served as the basis for developing the Zenit-3SL Integrated Launch Vehicle core of the offshore launch platform developed under the Sea Launch program. Work on the Sea Launch project began on 25 November 1993 when an agreement was executed between the American aircraft and missile company Boeing, the Russian Rocket and Space Corporation Energia, the Norwegian company Kvarner, and two Ukrainian enterprises—the Yuzhnoye State Design Office and the State Enterprise Production Association Yuzhnyi Machine-Building Plant. Geosynchronous satellite launches via Sea Launch began on 28 March 1999 with a demonstration launch of the Zenit-3SL launch vehicle.
Figure 3. Launch vehicles developed by the Yuzhnoye State Design Office.
Figure 4. Spacecraft developed by Yuzhnoye State Design Office.
As before, development of future launch vehicles remains at the center of attention. A more powerful launch vehicle, the Tsiklon-4, has been developed on the basis of the Tsiklon-3. Work is currently underway on the Air Launch and Mayak projects, and various approaches for modernization of the Tsiklon-2, Zenit-2, and Dnepr launch vehicles are also being explored.
From Missiles to Launch Vehicles—the First Missile (8K63) and Launch Vehicle (11K63) 8K63 (SS-4) Missile (Fig. 5). The first strategic missile to embody the new concept developed under the leadership of M.K. Yangel’ was the 8K63 (SS-4). The Government task order for developing the missile was issued to coincide with establishment of the OKB-586 design bureau.
The major responsibilities for development of the 8K63 missile/missile system were allocated as follows:
• OKB-586, Chief Designer M.K. Yangel’—systems engineering of missile and missile system as a whole;
Figure 5. 8K63 missile.
* KB-11, Chief Design Engineer S.G. Kocharyants—design of warhead and related equipment;
* NII-885, Chief Design Engineer N.A. Pilyugin—design of the autonomous onboard control system;
* NII-944, Chief Design Engineer V.I. Kuznetsov—design of gyroscopic instruments;
* OKB-456, Chief Design Engineer V.P. Glushko—design of RD-214 engines;
* Spetsmash State Special Design Bureau, Chief Design Engineer V.P. Barmin—design of aboveground and silo-based launch facilities.
These chief design engineers became the most active proponents of Yangel’s approach to missile system development.
The 8K63 consisted of a monocoque single stage that had a nose cone section for the nuclear warhead, cylindrical fuel tanks, an instrument section that had an autonomous onboard control system, and a conical tail compartment containing a fixed RD-214 four-chamber engine using TM-85/AK-27I high-boiling-point propellants. The AK-27I oxidizer is an iodine-inhibited mixture of nitric acid (70%) and nitrogen tetroxide (27%), and TM-185 is a modified-kerosene hydrocarbon fuel. At the time, these were the best-known high-boiling-point propellants that had an adequate production infrastructure.
The missile had a launch weight of41.7 metric tons, a length of22.1 m, and a body diameter of 1.652m. Its RD-214 engine produced a thrust of 648/744kN and a specific impulse of 2300/2640 N • s/kg (sea level/vacuum). TG-02, a xylidine/ triethylamine mixture, was used as an ignition propellant to ignite the fuel in the RD-214 combustion chamber. The engine turbopumps were operated using a gas-generator mixture obtained by decomposing hydrogen peroxide in the presence of potassium permanganate. The oxidizer and fuel tanks were pressurized by compressed air and compressed nitrogen, respectively, stored in high-strength cylinders. Four adjustable graphite control vanes were used; one vane was placed in the exhaust of each engine chamber. The main body of the missile was constructed from a lightweight, high-strength aluminum alloy.
The fuel tanks were made from nonreinforced cylindrical shells mounted between two bottom plates that were segments of spheres. The oxidizer tank was mounted forward of the fuel tank to control the center of gravity during flight. Moreover, an intermediate bottom plate was also mounted in the oxidizer tank for the same purpose, so that additional oxidizer would flow from the top to the bottom portion of the tank as the oxidizer was consumed from the bottom portion. The oxidizer feed line ran through a tunnel pipe built into the fuel tank. A riveted instrument compartment was located between the two tanks (Fig. 5). To shift the aerodynamic center of force closer to the center of mass, four fixed aerodynamic stabilizers were placed on the tail section of the missile.
This missile was flight-tested at the Kapustin Yar facility from July 1957 to December 1958. During this time, there were 24 launches of this missile, which confirmed that it was highly reliable. Some final design modifications were made, and the 8K63 missile was accepted into armaments for use with aboveground and silo launchers. The silo-launched version of this missile was assigned the code number 8K63U.
The requirement to develop both aboveground and silo-based launch facilities for this missile was largely dictated by the need to reduce the amount of time required for final development of the missile. Constructing a silo launcher would have required a large amount of time, so most test launches of this missile were done from a hastily constructed aboveground launch facility.
The tactical, engineering, and operational characteristics of the 8K63 represented a considerable advance over previous missiles. The 8K63 had a reaction time of 20min and could deliver a 2.3-MT warhead to a maximum range of 2080 km. For some time, this was the main missile used by the Strategic Missile Forces (established December 1959). At the same time the 8K63 was on operational duty in the Strategic Missile Forces, it also served for 25 years as the primary launch vehicle for testing new technology and designs of warheads and antimissile defense systems. The 8K63 was decommissioned in July 1988 pursuant to the Treaty on the Elimination of Intermediate-Range and Shorter-Range Missiles (INF Treaty).
11K63 (SL-7) Launch Vehicle (Fig. 6). Even the first few launches of spacecraft into low Earth orbit had demonstrated a wide variety of new opportunities to study Earth and near-Earth space using space-based instrumentation. These opportunities stimulated wide interest among scientists, businesspeople, and the military in obtaining information on Earth’s surface, Earth’s upper atmosphere, Earth’s magnetic field and cosmic rays, the interaction between these particles and Earth’s magnetic field, and the effects of space environment on objects launched into space.
This generated a requirement to launch a large number of spacecraft into low Earth orbit for various purposes. A need for low-cost launch vehicles therefore arose. The three-stage launch vehicle then (late 1950s) in use in the Soviet Union, the 8K72 Vostok, was not appropriate for frequent use as a launch vehicle due to the relatively high cost of launch, the relatively large amount of time required to prepare it for launch, and the fact that it was generally used to address more prestigious problems. Thus, in late 1959, OKB-586 embarked on an initiative to develop a two-stage launch vehicle based on the mass-produced 8K63 missile. This proposal was supported by the USSR Academy of Sciences and Ministry of Armaments, each of which were interested in the development of an inexpensive launch vehicle to address their specific needs using small spacecraft placed in low Earth orbit.
OKB-586 received the Government order to develop this launch vehicle, which came to be called the 11K63, in August 1960. The Government authorized the production of ten 11K63 launch vehicles and use of these vehicles to launch 10 spacecraft; each was for a different purpose and carried different instrumentation. Two of these spacecraft, designated the MS series, were developed by OKB-1, and the remaining eight spacecraft, designated the DS series, were developed by OKB-586.
The main tasks required for developing the 11K63 launch vehicle involved developing a second stage and aerodynamic fairing and adapting these components to a first stage that was virtually identical to the 8K63 missile. The second stage and aerodynamic fairing weighed ~7.7 metric tons, giving the 11K63 launch vehicle a launch weight of 49.4 metric tons, and a length of 30 m.
The second-stage fuel and dry compartments of the launch vehicle were similar to the corresponding first-stage compartments. However, there were also several differences due to the fact that the second-stage RD-119 engine required liquid oxygen and unsymmetrical dimethylhydrazine (UDMH) as propellants. This engine had been developed by OKB-456 for use on the Vostok launch vehicle but ended up not being used for a variety of reasons. The RD-119 engine was fairly well developed and also had relatively good energy performance characteristics (thrust and specific impulse in vacuum 106 kN and 3454 N • s/kg, respectively). The existence of these additional propellants undoubtedly made operation of the launch vehicle more complicated, but the availability of a fully developed engine reduced the effort and time required to develop the launch vehicle. Therefore, OKB-586 decided to use the RD-119 engine in the second stage of its first launch vehicle.
Figure 6. 11K63 launch vehicle.
The RD-119 engine is a fixed, single-chamber, liquid-fueled engine installed on the second stage in combination with several movable low-thrust nozzles used to control the second stage in pitch, yaw, and roll. The engine is started using a pyrotechnic device. Positive pressure in the oxidizer tank was maintained by evaporating oxygen in a heat exchanger mounted on the engine’s turbine exhaust pipe. Positive pressure in the fuel tank was maintained by using a mixture of producer gas and UDMH vapor. The RD-119 engine was only capable of single-use operation; as a result, spacecraft were launched to place them directly into orbit—primarily low-level highly elliptical orbits. To increase the amount of UDMH that could be stored in the second-stage fuel tank, it was initially cooled to — 45°C. The second stage was mated to the first stage using a tubular beam that had a conical heat shield attached to the lower chord to protect the first stage from the exhaust of the RD-119, as it pushed away the first stage during the stage separation process.
The spacecraft was initially housed under a conical/cylindrical aerodynamic fairing (jettisoned during the boost phase of the flight after passing through the dense layers of the atmosphere). The spacecraft was separated by using pusher springs. Like the first stage, the second stage of the launch vehicle had an autonomous onboard control system developed by the newly established OKB-692 in Kharkov (now NPO Khartron-Arkos) under the direction of Chief Designer B.M. Konoplev). Initially, the 11K63 was to be launched from the 8K63U launcher at the Kapustin Yar Test Site, and an appropriate operational scenario was developed for the launch vehicle under this assumption (including use of a silo launcher that was shorter than the launch vehicle).
The first launch of the 11K63 from a silo took place on 27 October 1961. Both the first and the second launches were unsuccessful. Nearly 5 additional months were required to eliminate all of the problems. The third launch of the 11K63 took place on 16 March 1962 and was successful. The first spacecraft designed and built by OKB-586 personnel, the DS-2, had been placed into Earth orbit. After 37 standard 11K63 silo launches from the Kapustin Yar Test Site, all further 11K63 launches were from a new, aboveground facility at the Plesetsk Test Site, which was developed by the Design Bureau for Transportation Machinery directed by Chief Designer V.N. Sobolev.
There were several differences between the operational configuration of the 11K63 for silo launches and that for surface launches. During a silo launch, final assembly of the launch vehicle occurred during placement in the silo. The first and second stages (including spacecraft) were tested in the launch support facility, transported to the launch facility separately in the horizontal position, and then placed in the silo in the correct order. During a surface launch, final assembly of the launch vehicle took place in the launch support facility, and the assembled launch vehicle was transported to the launch facility, where it was raised into a vertical position using special equipment rather than the gantry.
The 11K63 could place a payload of up to 450 kg into circular low Earth orbit (200 km altitude and inclination 82°). The 11K63 became the first Soviet launch vehicle to be mass-produced. It was accepted for operational use in 1965, along with the DS-P1-Yu spacecraft (developed by OKB-586) and the new aboveground launch facility at the Plesetsk test facility. The 11K63 was launched a total of 165 times, of which 143 were successful. Numerous Kosmos- and In-terkosmos-series spacecraft were launched using the 11K63. It was used for 16 years until the final launch on 18 June 1977.
8K65 (SS-5) Missile and 11K65 Launch Vehicle (Fig. 7). The 8K65 became the second missile developed by OKB-586 within its new conceptual framework. The government task order for developing this missile was received in early July 1958. The goal was to develop, within 2 years, a missile that had twice the range of the 8K63, which was then in the final stages of development. This promised to be a more difficult effort than the 8K63 had been and would be made even more difficult by the fact that OKB-586 had already been working for a year and a half on developing the 8K64 intercontinental ballistic missile, for which OKB-586 had received the task order in December 1956.
Figure 7. 8K65 missile.
However, gaining experience, OKB-586 enthusiastically began work on developing the 8K65. The only decision valid at that point was made: Use the engineering design solutions that had proved themselves during the 8K63 design effort, as well as several of the basic design solutions that lay at the core of the 8K64 design. It was decided to use a monocoque, single-stage design for the 8K65 similar to that for the 8K63, but using a higher energy propellant combination. The 8K63 used the same propellant combination used in the 8K64; unsymmetrical dimethylhydrazine (UDMH) was the fuel, and AK-27I was the oxidizer. This propellant combination enabled a 15% increase in the total energy output of the 8K65 over the TM-185/AK-27I propellant combination. The use of UDMH enabled the entire missile to operate on only two propellants, thereby eliminating one deficiency of the 8K63—the fact that several propellants were required. One important quality of the propellants selected was that they would ignite on contact and thereby eliminated the need for an ignition propellant. This also simplified development of the turbopump-boost gas generators, which could use the same propellant combination.
The opportunity to use UDMH in the 8K65 resulted from the success achieved by the State Institute for Applied Chemistry (director V.S. Shpak) in researching this fuel and setting up the appropriate manufacturing facilities. However, the decisive factor enabling its use in the 8K65 was that the engine designers at OKB-456 had developed a two-chamber, liquid-fueled rocket engine (that had a thrust of 735 kN) that used this propellant combination. Two of these engines together as a single engine module [given the code number RD-216, whose thrust and specific impulse were 1481.3/1741.3 kN and 2413.3/2835.1 N • s/kg, respectively (sea level/vacuum)], were used on the 8K65.
The increased range required a corresponding increase in the amount of fuel and oxidizer carried, and thus, an increase in the diameter of the missile. To standardize production, a diameter of 2.4 m was adopted for the 8K65 main body; this was identical to the diameter of the 8K64 second stage.
The 8K65 was designed by virtually the same team of developers as the 8K63, and there were significant similarities to the 8K63 in both design and external appearance. Like the 8K63, the 8K65 was a single-stage monocoque design that had a conical warhead compartment (containing the same nuclear warhead as was used on the 8K63); a conical warhead mounting adapter, cylindrical fuel and oxidizer tanks; an instrument compartment containing the autonomous onboard control system that was located between the tanks; and a conical tail section that had a fixed RD-216 engine module. The 8K65 used exactly the same tank pressurization system, actuator-control design, and structural materials as the 8K63. Just as in the 8K63, the oxidizer tank was located forward of the fuel tank, and four fixed stabilizers were mounted on the tail section. The major differences between the 8K65 and the 8K63 were the new fuel; the increased fuel capacity; the larger diameter of the missile; the new, more powerful engine; and the new control system.
The 8K65 engines were ignited by taking advantage of the hypergolic nature of the propellants as they mixed in the combustion chamber, thereby simplifying the engine design. Using the same basic fuel/oxidizer components also enabled simplifying the gas generators used in the turbopump-boost system.
However, there were also some additional differences. For example, the cylindrical fuel tank fairings on the missile were made from reinforced molded panels. A system for simultaneously emptying the tanks (not available on the 8K63) was used to reduce the amount of unused fuel and oxidizer by synchronizing the consumption rates. This missile also marked the first use of a gyro-stabilized platform, that gave the missile the same accuracy as the 8K63, despite the factor of 2 greater range. The warhead unit was separated by braking the missile body using solid-rocket motors mounted on the missile instrument compartment housing.
The 8K65 was twice as heavy as the 8K63, but was only 2.3 m longer. It could deliver a warhead identical to that used in the 8K63 to a maximum range of 4500 km. Like the 8K63, the 8K65 was developed and placed into military service for use from surface and silo-based launch facilities. The silo-launched version had the code number 8K65U. Silo launches of the 8K65U were similar to those of the 8K63U; they involved the use of a launch canister and the gas pressure generated by the missile engines after being started in the silo. The 8K65 was flight-tested at the Kapustin Yar Fourth State Central Test Site. Forty-four launches of the 8K65 or 8K65U missile were performed as part of the flight/ development testing program. The 8K65 became the second, longest range missile in the Strategic Missile Forces arsenal. The 8K65 enabled targeing strategic facilities maintained by potential adversaries in Europe, Asia, and even Africa that were out of range of the 8K63. The missile remained on operational duty for 15 years, starting in 1962, and was decommissioned in 1987 pursuant to the INF Treaty.
11K65 (SL-8) Launch Vehicle (Fig. 8). The second launch vehicle, designated the 11K65, was developed by OKB-586 pursuant to a USSR Government Decree issued on 30 October 1961. According to the stated requirements, the 11K65 launch vehicle would launch one (or more) spacecraft into Earth orbit. The orbits were required to be either elliptical or circular, the circular orbits had a maximum altitude of 2000 km. Stringent requirements were also imposed on the accuracy of the resulting orbital parameters.
The spacecraft to be launched by the 11K65 were expected to remain active for long periods of time, and therefore would weigh more. In view of all these requirements, the 11K65 had to have approximately three times the total spacecraft launch capacity of the 11K63. Preliminary studies performed by OKB-586 indicated that such a launch vehicle could be built quickly and at minimal expense in terms of materiel, if the 8K65 missile were used as a first stage. Design optimization of the launch-vehicle parameters revealed that it would be desirable to design the second stage so that it would have the same diameter, use the same propellants as the first stage, and weigh ~ 24 metric tons. Placing such a stage on the 8K65 missile required partially modifying its frame configuration and developing a new interstage compartment that could bear the load of the second stage and also provide an outlet for the gas generated by the steering nozzles on the second-stage engine during stage separation. To reduce the length and weight of the second stage, it was decided to use a somewhat different design compared with that of the first stage. The second-stage fuel tanks were combined into a single cylindrical propellant compartment that had three spherical-segment bottom plates and a common intermediate bottom plate separating the propellant compartment into an upper cavity containing the oxidizer and a lower cavity containing the fuel. The instrument compartment was located above the propellant compartment, and a spacecraft adapter and clamshell nose fairing were attached to the top of the instrument compartment to protect the spacecraft against the air flow while the launch vehicle moved through the lower atmosphere and to protect the spacecraft during ground operations.
Figure 8. 11K65 launch vehicle.
The requirements that the 11K65 reach a high-altitude circular orbit and launch heavy spacecraft into such orbits led to the use of a two-burn orbital insertion procedure for placing spacecraft into such orbits. In this procedure, the launch vehicle second stage and spacecraft were placed into an elliptical transfer orbit whose apogee was equal to the desired altitude of the circular orbit, and were then transited into the desired circular orbit. The problem came in obtaining the appropriate magnitude and direction of the velocity vector required for the second stage to reach the perigee of the elliptical transfer orbit and then the circular orbit. Implementation of this procedure affected the design concepts for the control system, the second-stage engine, and the propellant feed system.
The second-stage engine, which came to be known as the S5.23, actually consisted of two subengines: a single-chamber main engine and a four-chamber steering engine. The main engine was fixed relative to the missile. The four steering-engine nozzles were combined with the four small thruster nozzles to form four combined, steerable units mounted on the main engine. The steering engine enabled control of the second stage while the main engine was in operation; the thruster system was used for stabilization and attitude control of the second stage during the passive portion of the trajectory. This engine was developed by OKB-2 (now known as the Chemical Machinery Design Bureau) under the direction of Chief Designer A.M. Isaev. Features of the S5.23 engine included three thrust modes (vernier, intermediate, and primary) and the capability of reusing the primary-thrust mode. In primary mode, the engine thrust was produced by the main engine and its four steering nozzles and was equal to 157 kN (specific impulse 2913.6 N • s/kg). In intermediate-thrust mode, only the steering nozzles were used (thrust 5.4 kN). This mode was used during engine start-up and shutdown. The low-thrust mode (thrust 0.1 kN) was supported by the passage of gas produced in a special gas generator through small nozzles. Between the two engine firings, the fuel and oxidizer for the second engine firing, and for the engine firing in the stabilization/attitude control mode during the passive portion of the trajectory, was stored in two small tanks located on the outside of the second-stage propellant compartment.
The oxidizer and fuel tanks were pressurized by using compressed air and compressed nitrogen, respectively, stored in high-pressure cylinders.
The control system for the second stage was developed by OKB-692 under the direction of Chief Designer V.G. Sergeev. To meet the requirements for increased accuracy of spacecraft placement into orbit, relatively high-speed computers and high-accuracy control devices (that had increased throughput and special software and algorithms) were used for the first time. The engineering design solutions implemented for the 11K65 gave it the ability to place up to eight spacecraft that had a maximum total weight of 1500 kg into Earth orbit (circular orbit, altitude 200 km, and inclination 51°) in a single flight. OKB-586′s role in developing the 11K65 was limited to preparing the preliminary design and design documentation, as well as fabrication and design/flight-testing of the first 10 launch vehicles. Due to the workload related to developing the new 8K67 missile, the 11K65 development task was transferred to OKB-10 (now known as the Academician M.F. Reshetnev Scientific Production Association for Applied Mechanics) in late 1962.
Development flight-testing of the 11K65 began on 18 August 1964 at the Baikonur Cosmodrome from a modified aboveground launch facility designed by the Novokramatorsk Machinery Plant Design Office. Routine 11K65 operations have been conducted at the Plesetsk Test Site since 1967 and the Kapustin Yar Test Site since 1973. Aboveground launch facilities that had movable gantries were built at each of these test sites. These launch facilities were designed by the Design Bureau for Transportation Machinery. Between 1965 and 1967, OKB-10 upgraded the 11K65 to improve its operational specifications, after which it was renamed the 11K65M (Kosmos-3). To date, there have been more than 500 launches of the 11K65 or 11K65M that carried more than 1000 Kosmos- and Interkosmos-series spacecraft into Earth orbit, including more than 130 DS-1P, Tselina, Tyul’pan, Taifun, and AUOS (”automated universal orbital stations”) spacecraft designed by OKB-586 (Yuzhnoye Design Office). The 11K65 remains one of the most reliable, highly productive launch vehicles in the Russian Federation inventory. For some time, a modified version of this launch vehicle, the K65M-R, has also been used for testing warhead units and missile-defense countermeasures.
8K69 (SS-9 mod 3) Missile—The Basis for the 11K69 and 11K68 Launch Vehicles (Fig. 9). The Government task order for developing the 8K69 missile was issued on 14 April 1962, simultaneously with the task order for developing the 8K67. The 8K67 was given an accelerated development schedule. Many of the requirements for the 8K67 and 8K69 missiles were either very similar or identical, so the decision was made to standardize these missiles in areas where the engineering design solutions were consistent. Thus, both stages of the 8K67 were used in the 8K69, after a few slight design modifications, mainly related to the fact that the 8K69 used a warhead different from the 8K67.
The 8K69 was developed as a countermeasure to U.S. deployment of the Safeguard missile defense system, which protected U.S. territory against missile attack from the north. This missile had several unique features. It had unlimited range and could reach targets from any direction; therefore, it could deliver a nuclear warhead to any area of the United States by evading the missile defense system (and rendering it useless). At the time, these features led to calling the 8K69 a ”global missile.” The 8K69 warhead could come in from various directions by being placed into Earth orbit at various azimuths and then reentering from orbit and hitting the target.
Like the 8K67, the 8K69 was a two-stage monocoque design in which the stages were on top of each other and stage separation occurred around the circumference of the missile (launch weight 181.3 metric tons and length 32.65 m). Both missile cases were cylinders 3 m in diameter. Both stages (and the warhead) used the same fuel/oxidizer pair: unsymmetrical dimethylhydrazine (fuel) and nitrogen tetroxide (oxidizer). This fuel/oxidizer pair gave the missile a high total energy output and also substantially increased the guaranteed military readiness lifetime (to 7 years).
Figure 9. 8K69 missile.
The first stage consisted of four sections: a stage transition section, an instrument section, a tail section, and propellant-tank section. The second stage had three sections: an instrument section, a propellant section, and a tail section. The second-stage propellant tanks were combined into a single section that had three spherical-segment bottom pieces. The second-stage instrument section was conical. All “dry” sections (stage transition section, instrument sections, and tail sections) in the missile and the warhead section were riveted, whereas the fuel tanks were welded. The fuel-tank design used molded panels and molded stock hollow frames, which led to a substantial reduction in the number of processes required, as well as an overall simplification of the tank fabrication process. To reduce the structural weight, light, high-strength aluminum alloy was used for all missile and warhead sections.
The first-stage propulsion system consisted of an RD-251 main engine and an RD-855 steering engine (total thrust 2651.6/2974.3 kN and specific impulse 2627.1/2945.9 N • s/kg at sea level and in vacuum, respectively). Structurally, the RD-251 main engine consisted of three identical engines affixed to a single frame. Each of these engines had two chambers, a turbopump unit, a gas generator, a solid-fuel starter, automatic control systems, and various other components. The RD-251 engine was attached to a frame, which was in turn attached to the fuel tank bottom frame, and was enclosed by the tail section. Solid-fuel starters were used to start all three engines simultaneously within a few seconds of steering engine start-up. The engines could also be turned off simultaneously a few seconds before the steering engine, in response to a command from the command system.
The RD-855 steering engine controlled the missile in roll, pitch, and yaw. It consisted of four steerable chambers and a stationary turbopump unit, gas generator, solid fuel starter, automatic control systems, and various other components. The steering engine chambers were steered by using hydraulic actuators: UDMH was obtained from the motor turbopump unit as the working fluid. The steering motor was started and turned off by the missile control system in accordance with a timed cycle schedule. The steering engine chambers were mounted in four fairings on the exterior surface of the tail section. Two of these fairings also contained solid-rocket motors to decelerate the first stage after stage separation.
The second-stage propulsion unit consisted of an RD-252 main engine and an RD-856 steering engine (total thrust 9955.7 kN and specific impulse 3093.1 N • s/kg in vacuum). The RD-252 design and systems were similar to those used in the first-stage, two-chamber, liquid-fuel engine but had larger nozzles for a higher maximum altitude. Like the RD-251, the RD-252 was affixed to an engine frame that was in turn attached to the fuel tank bottom frame, and was enclosed by the second-stage tail section.
The RD-856 steering engine was designed to control the second stage, which operated in a less-dense atmosphere. Therefore, the engine had much lower thrust than the RD-855 but was similar in design and was based on similar systems. The second-stage engine was mounted similarly to the first-stage engine. The steering chambers were likewise housed in four compartments mounted on the exterior surface of the second stage. Solid-propellant motors were also mounted in two of the compartments; these motors were used for second-stage braking after warhead separation. Second-stage engine start-up and operation were supported by using a control system similar in design to that used for first-stage engine start-up and operation. The main engines for both stages were developed by OKB-456, the steering engines were developed by OKB-586, and the retroengines were designed by the Iskra Machinery Plant in Moscow.
The fuel and oxidizer tanks for each stage were pressurized by using hot gases produced from the primary propellants using special gas generators (the steering-engine propellant feed system was a source).
The 8K69 autonomous onboard control system supported prelaunch missile preparation and silo-based launches. The control system also supported fairly high readiness and target accuracy. To improve the reaction time of the missile, the control system used gyro units that could be forced into operational mode. The orbital weapon unit (OWU) consisted of a relatively large single warhead, as well as a reentry stage and an interstage bay that connected the orbiting weapon unit to the second-stage instrument bay. The OWU reentry stage included an instrument bay that held the control system and toroidal fuel tanks; an RD-854 retro-engine and vernier thruster were located in the interior cavity between the tanks. The RD-854 is a fixed, single-chamber liquid-fueled engine (thrust 75.5 kN, specific impulses 3063 N • s/kg), that uses eight fixed exhaust nozzles to control the orbital weapon unit during braking maneuvers. This engine was started by using a pyrotechnic starter. The vernier thruster (8 nozzles, thrust 0.03 kN) was used for OWU damping after OWU separation from the second stage, for OWU trajectory stabilization during passive flight, and for OWU attitude stabilization before retroengine start-up. The RD-854 engine and thruster were designed by OKB-586. The control systems for the missile and orbital weapon were designed by OKB-692, and the gyro instruments were developed by NII-944.
The missile was silo-launched using the gas pressure generated by the first-stage missile engines after they were started in the silo. The missile was launched from a stationary launch platform located inside the launch silo. After the missile left the silo, it was turned toward the target azimuth in roll using the first-stage steering engines. Missile-mounted hooks slid along guides in the launch canister to assure the safety of missile launches. In the interests of launch safety while the engines were in operation, the gas outflow from the first-stage engines was directed away from the missile by special baffles.
Ground-based firing-stand tests of the OWU reentry stage and aircraft-based testing of the OWU reentry stage under weightless conditions were followed by development flight-testing of the 8K69 missile, which involved some 19 missile launches. Final design modifications of the first and second stages had already been made as part of the 8K67 missile development project, so flight-testing of the 8K69 was limited in scope and basically involved integration testing of the OWU and missile. The 8K69 was accepted into armaments in 1968 and remained in military service for 15 years.
11K69 (SL-11) and 11K68 (SL-14) Launch Vehicles (Figs. 10,11). OKB-586 began developing the 11K68 and 11K69 launch vehicles in August 1965 pursuant to a government decree. The USSR Ministry of Defense, the main party interested in developing these launch vehicles (LVs), hoped to use them as delivery systems for both long-term or tactical space-based reconnaissance systems and antisatellite defense systems. As a result, these launch vehicles had to be extremely flexible and rapidly launched. The 8K69 missile was selected as the basis for both LVs. Virtually identical versions of the first two stages of this missile were used in both the 11K68 and 11K69 LVs. This led to standardization of many launch-support-facility and launch-facility components for these LVs, as well as a reduction in development cost.
Figure 10. 11K69 launch vehicle.
Figure 11. 11K68 launch vehicle.
There were various reasons behind selecting the 8K69 missile as the basis for the 11K68 and 11K69 launch vehicles. The 8K69 missile, which could place orbital warheads into Earth orbit, could already, after some modification, place other spacecraft into orbit. Additionally, the 8K69 had a relatively high total energy output and could place payloads of up to 3000 kg into low Earth orbit. It could also be readied for launch fairly rapidly, thereby laying the groundwork for meeting the 11K68 and 11K69 requirements for flexibility and rapidity of launch.
Another consideration was the fact that the first and second stages of the 8K69 were already in final development as part of the 8K67 missile and the necessary series production facilities had been established for them. This enabled reducing the one-time development costs for the 11K68 and 11K69 LVs. These considerations served as the basis for the decision to base the 11K68 and 11K69 LVs (which later came to be called, respectively, the Tsiklon-3 and Tsiklon-2) on the 8K69 missile.
The 11K69 LV is an 8K69 missile in which the orbital warhead has been replaced by one of two nose units containing either a reconnaissance spacecraft or an anti-satellite system. The anti-satellite system consisted of a guidance stage [developed by the Central Machinery Design Bureau (TsKBM), Chief Designer V.N. Chelomei], and an antisatellite weapons system (developed by Comet Design Bureau, Chief Designer A.I. Savin). Each of these standard 11K69 pay-load types can be placed under a standard nose fairing.
The launch vehicle is assembled in the Launch-Support-Facility Test and Assembly Building, with the launch vehicle and nose units in horizontal position. The 11K69 LV is transported to the launch pad in assembled form using a custom transporter-erector. Various assemblies installed on the transporter-erector ensure proper mating (or demating) of all pneumatic lines, hydraulic lines, electrical lines, and mechanical joints, the appropriate lines and connections to the launcher are made as the transporter-erector moves over them. The LV is placed into vertical position using the transporter-erector, which is also used for all further servicing (no gantry tower is used).
The 11K69 LV marked the first use of a safe, automated, crewless launch system that completely eliminates any need for operations personnel in the vicinity of the launch facility during the most hazardous launch operations. This system of operations also led to a substantial reduction in the amount of time required to prepare the LV for launch.
The 11K69 space launch system was developed in cooperation with the Design Bureau for Transportation Machinery, the Elektropribor Design Bureau, and various other organizations. The space launch system for the 11K69 LV was located at the Baikonur Cosmodrome (Scientific Research Test Site 5). Standard
11K69 launches have been performed since 6 August 1969. This LV holds a unique record in rocketry and space history. All 103 launches of this launch vehicle have been successful, and the 11K69 is still in regular use.
The 11K68 LV differs from the 11K69 LV in that the former has a newly developed third-stage, nose fairing, and interstage-skirt section. This interstage-skirt section consists of an inverted frustrum of a cone because the mating diameter of the nose cone was greater than that of the second-stage instrument bay. The nose cone is intended to protect the spacecraft and third stage from external effects during ground and flight operations. The nose cone consists of a conical cylindrical “clamshell” whose two halves are held together by a special locking device embedded in the longitudinal surface of the nose cone. Once the dense layers of the atmosphere have been traversed, the mechanical linkage between the two halves is severed, and they are jettisoned from the launch vehicle.
The third stage of the 11K68 launch vehicle includes a spacecraft adapter, an instrument section, fuel/tail section, RD-861 main engine, 11D75 vernier thruster, a control system, and a telemetry system. To reduce the weight of the third stage, the instrument compartment uses a space-frame design. In addition to the control system devices, the instrument compartment also houses a spherical high-pressure tank containing helium used by the propellant-tank pressur-ization system. The telemetry system instruments are mounted on the exterior of the instrument section, along with the all-riveted spacecraft adapter structure. The third stage uses a number of engineering design solutions tested during basic design of the orbital-weapon-unit (OWU) braking engine assembly on the 8K69 missile. Like the OWU, the third stage was sealed. The third stage also used the same propellants as the OWU, nitrogen tetroxide and unsymmetrical dimethylhydrazine. The third stage can be stored for several years in the fueled and gas-filled state and transported to the launch pad in this state as part of a launch vehicle. This engineering design solution led to a significant reduction in launch preparation time. Both the third stage and the OWU used a toroidal fuel compartment that had a fixed main engine mounted in the interior cavity. The thruster was also fixed and was mounted in the third-stage tail compartment.
By contrast with the OWU, the third stage toroidal fuel section was cylindrical rather than conical and had one and a half times the volume. A common intermediate bottom plate divides the section into an upper cavity for the ox-idizer and a lower cavity for the fuel. The cavities include baffles to prevent propellant oscillation, intake assemblies, and other fittings. Fine-mesh mist extractors were used to support propellant feed into the main engine during startup under weightless conditions.
The RD-861 main engine is an improved version of the RD-854 engine used in the OWU and has a higher total energy output (thrust 78.1 kN, specific impulse 3374 N • s/kg) than the RD-854. However, its main features are its ability to be fired twice under weightless conditions and extended operational life. These features enabled using a two-impulse spacecraft launch trajectory that had an elliptical transfer orbit and resulted in the ability to increase the altitude of the orbits that could be reached, as well as the mass of the spacecraft that could be placed in such orbits.
The RD-861 main engine is a single-chamber, liquid-fueled engine that has an open-loop turbopump-based propellant feed system; the generator gas used to operate the turbopumps is discharged through eight nozzles operated by the third-stage flight-control actuator system while the main engine is in operation.
The 11D75 vernier thruster is an improved version of the OWU thruster on the 8K69. Just as for the main engine, both the number of times the thruster can be operated and the operating lifetime have been improved; the functions performed by the thruster were also expanded. It was used for attitude control and stabilization of the third stage and also to support main-engine restart in weightless conditions. To this end, the thruster was equipped with two special nozzles to create a micro-overload in the longitudinal direction for approximately 100 seconds until the third-stage main engine restarted. The 11D75 thruster has an independent fuel system that is filled from the primary third-stage fuel tanks while the first two stages of the 11K68 launch vehicle are in flight. The RD-861 and 11D75 engines were developed by the Yuzhnoye Design Office.
The 11K68 control system consists of two interconnected systems; one has equipment installed in the first- and second-stage instrument sections, and one has equipment installed in the third-stage instrument sections. The control system for the first two stages of the launch vehicle supports prelaunch preparations, launch, and flight-control functions until third-stage separation, and the second control system controls the third stage during subsequent spacecraft orbital insertion phases. The control system for the first and second stages of the LV was developed by the Elektropribor Design Bureau; that for the third stage was developed by the Kiev Radio Plant (now the Kiev Radio Plant Production Association) Design Bureau. Like the 11K69, the 11K68 was operated in horizontal mode. Like the 11K69, preparation of the 11K68 for launch was automated using standardized assemblies, command lines, actuator lines, and data lines.
The ground facility for these launch vehicles was developed by the Design Bureau for Transportation Machinery and located at the Plesetsk Test Site. To date, there have been 119 launches of the 11K68 launch vehicle; 114 were successful. Several launches have supported the deployment of up to six spacecraft in one flight. More than 10 different types of spacecraft have been deployed into a variety of orbits using the 11K68.
15A18 (SS-18 mod 4) Missile (Fig. 12). The 15A18 is one of the most powerful ballistic missiles ever developed by the Yuzhnoye Design Office; it is designed to deliver 10 warheads weighing approximately 8.5 metric tons to a range of up to 11,000 km. The missile had a launch weight of 211.1 metric tons, a length of 34.3 m, and a case diameter of 3 m. One such missile can destroy up to 10 arbitrarily located targets within an area several hundred kilometers across using the nuclear warheads carried by the missile. The 15A18 has the capability of high-accuracy strikes on enemy facilities even after multiple enemy nuclear strikes against 15A18 deployment areas.
The government task order for developing the 15A18 missile system was issued on 16 August 1976. The new missile was to be based on the 15A14 missile system, and the main reason for developing the new system was to modernize the 15A14 for increased military effectiveness and increased launch-facility hardening. The improved 15A14 missile was designated the 15A18. The task order called for increasing the yield of the nuclear warheads, improving the effectiveness of the missile-defense countermeasures used, increasing the target accuracy, increasing the range, and increasing the dimensions of the area tar-getable by the reentry vehicles to meet certain specifications. This task was successfully accomplished.
Figure 12. 15A18 missile.
These requirements were met through the following developments:
• high-strength reentry vehicles carrying higher-yield nuclear warheads;
• multipurpose missile-defense countermeasures;
• two-level weapons compartment that has clamshell fairing;
• high-energy output liquid-fueled reentry-vehicle stage separation units and missile-defense countermeasures;
• missile control systems that have gyroscopes of enhanced accuracy and an improved onboard digital computer; and
• targeting systems that remain functional and ensure highly accurate targeting, even in the event of multiple enemy nuclear attacks on the 15A18 deployment area.
A launch silo that had improved hardening against nuclear explosions was also developed.
The 15A14 missile was used for the first and second stages of the 15A18 after several modifications. These missile stages have a highly compact layout to maximize the amount of propellant carried within the limited dimensions of the missile. The first-stage propulsion unit is a module consisting of four single-chamber RD-264 engines developed by the Design Bureau for Power Machinery (total thrust 4167.3/4524.4kN and specific impulse 2877.3/3123.5 N • s/kg at sea level and in vacuum, respectively). The second stage uses an RD-229 single-chamber, fixed, main engine and an RD-0230 four-chamber steering engine (total thrust 760.3 kN and specific impulse 3193.5 N • s/kg in vacuum). Both of the second-stage engines were developed by the Design Bureau for Automated Chemical Equipment. To increase the specific impulse, the RD-264 and RD-0229 engines had a closed-loop design in which the gas-generator output used to operate the turbopump assemblies was also burned in the combustion chambers.
The RD-864 engine on the RV bus was developed by the Yuzhnoye Design Office (thrust 19.6 kN and specific impulse 2848 N • s/kg). It is a four-chamber, liquid-fueled engine built to a new design; before engine start-up, the chambers are moved outward from the body of the bus and are stopped at a certain specific angle to the longitudinal axis of the bus so that the stage appears to be drawn forward. This simplified the system for separating the warhead unit and missile-defense countermeasures system components from the RV bus.
The first stage is controlled during flight by gimbaling the RD-264 engine, and the second stage and reentry stage are controlled by gimbaling the RD-0230 and RD-864 steering chambers. The chambers are gimbaled using hydraulic actuators where UDMH is the working fluid (the UDMH is fed by the turbopumps for each of these respective engines).
The tanks were prepressurized before engine start-up using a chemical pressurization system based on injecting a metered amount of oxidizer into the fuel tank and fuel into the oxidizer tank. In-flight pressurization of the first two stages of the missile used hot gas produced in special gas generators. The reentry stage used a gas-cylinder-based pressurization system.
The missile stages and warhead were separated by using braking systems on the first and second stages based on the impulse generated by blow off of pressure from the fuel tanks in each missile stage.
The missile was launched mortar-style from a launch canister placed in the silo. Launching the missile mortar-style provided the following advantages: substantial simplification of both the shock-absorption system for the surface test and launch equipment and the launch-silo design; improved protection against nuclear blast effects; reduction in launcher cost; and a reduction in the time required for launch silo construction and for placing the missiles into military service. The missile was flight tested at the Baikonur Cosmodrome. The flight-testing program for the 15A18 included 19 launches, of which 17 were successful. Dnepr Launch Vehicle (Fig. 13). The SALT-I and SALT-II treaties authorized the use of decommissioned strategic missiles for spacecraft launches. In this context, the Yuzhnoye State Design Office proposed a design for a missile/space system, called the Dnepr, based on the 15A18 (RS-20B) missile, which was then being decommissioned. The project was supported by the National Space Agency of Ukraine and the Russian Space Agency.
An international corporation, Kosmotras, including the following Ukrainian and Russian enterprises, was established to pursue this project: the Yuzhnoye State Design Office, the State Enterprise Production Association Yuzhnyi Machine-Building Plant, the Khartron Joint-Stock Corporation, and the Design Bureau for Special Machinery. The resulting company took responsibility for developing, operating, and marketing the system. The project was based on the availability of a significant base of materiel that could be used to address the purpose at hand: more than 150 15A18 missiles on duty in launch silos and stored in arsenals, as well as the experimental 15A18 ground facility at Baikonur Cosmodrome. Implementation of this project required some slight modifications to the missiles, including the onboard control systems and the launch support facility/launch facility at Baikonur Cosmodrome.
The Dnepr launch vehicle is essentially a modified 15A18 in which the weapon section is replaced by a spacecraft payload. The high energy output and high orbital inclinations reachable by using the Dnepr launch vehicle will enable it to be used for placing communications, remote-sensing, and scientific satellites into low Earth orbit. The first launch of the Dnepr took place on 21 April 1999, placing the British UoSAT-12 spacecraft into orbit. The second launch of the Dnepr launch vehicle took place on 26 September 2000, placing the following five spacecraft into orbit: UniSat and MegSat-1 (Italy), SaudiSat-1A and SaudiSat-1B (Saudi Arabia), and TiunSat-1 (Malaysia).
Treating space research as a highly valuable engine of scientific, technical, and economic progress, the Yuzhnoye Design Office, which began production of rocketry and space hardware more than 45 years ago, is continuing its space activities in developing new spacecraft launch systems for the Ukrainian national space program and other international space programs.
Figure 13. Dnepr launch vehicle.
