Category Energiya-Buran

In the course of Energiya’s history several studies were made of configurations in which the core stage was flanked by just two strap-ons, providing payload capacities of between about 30 and 60 tons. The first version, called RLA-125, was proposed in 1976 and another one known as Groza (“Thunderstorm”) appeared in the mid-1980s. Groza, using a standard core stage with four RD-0120 engines and two strap-ons, had a reported payload capacity to low orbit of up to 63 tons. The Cargo Transport Container strapped to the side would be a downsized version of that developed for Buran-T. Groza required virtually no modifications to the existing Energiya pads. All that needed to be done was to bolt the strap-ons more firmly to the pad because the rocket would be more susceptible to high winds. Because of this, launch weather rules were also tightened.

On 25 December 1984 the Soviet government released a major decree on rocket and space systems to be developed in the period 1986-1995. One of these was planned to be a series of rockets with payload capacities between 30 and 60 tons, although it is not clear what payloads exactly were being considered. Three systems were adopted for parallel studies: a modernized version of the Proton rocket, several heavier versions of the Zenit (11K37), and Groza. As mentioned earlier, these boosters were to use a standardized series of cryogenic upper stages, Shtorm for Proton and Vikhr for the 11K37 and Groza.

The preliminary design for Groza was completed in December 1985. However, on 18 August 1988 the Ministry of General Machine Building ordered NPO Energiya to modify the rocket in order to make it compatible with more realistic payloads of between 25 and 40 tons. This made it necessary to reduce the number of RD-0120 engines to one or two and hence make the core stage smaller. The first idea was to reduce the diameter of the core stage to 4.1 m or 5.5 m and lower the propellant mass to between 200 and 450 tons. However, since this would have required different manufacturing techniques, it was decided to retain the standard Energiya core stage diameter of 7.7 m. By late 1989 engineers were focusing on a version with one RD-0120 engine and a propellant mass of around 240 tons. With the core stage (called Blok-V) only about half as high as that of Energiya, the payload had to be stacked on top rather than strapped to the side. At the intersection between the core stage and the payload bay the rocket would taper off to a diameter of 6.7 m, the same as that of the 14S70 Cargo Transport Container of Buran-T. With a length of 25 m,

the payload bay was probably almost identical in dimensions to the Cargo Transport Container for Groza. The concept was approved by the Council of Chief Designers on 19 July 1990. Initially called Neytron (“Neutron”), the new rocket eventually became known as Energiya-M.

Four configurations were considered for the payload section, one in which the satellite would occupy the entire bay and have its own engine system (N-11) and three where the satellite would be attached to various upper stages (N-12, N-14, and N-15).

The N-12 was a Blok-DM modification with an engine known as the 11D58MF and was also planned for use on Zenit, Proton, and the original Angara. It allowed the rocket to place 29 tons into low Earth orbit or up to 3 tons into geostationary orbit. The N-14 was a Blok-DM modification with the standard 11D58M engine and was identical to the second stage of the 204GK upper-stage combination planned for Buran-T. It was capable of delivering a 5.5-ton payload to geostationary orbit. The N-15, able to launch 6.5 tons into geostationary orbit, was a LOX/LH2 upper stage but no further information on this is available. It is known that in 1992 work got underway on a LOX/LH2 upper stage known as Yastreb carrying the RO-97 engine of KBKhA. This stage was primarily intended for Proton, Zenit, and Angara, but with slight modifications could also be mounted on Energiya-M. However, it was smaller than the N-15 and also had its propellant tanks configured differently.

As early as 1990 a mock-up of Energiya-M was ready for tests at Baykonur. It was placed both on the UKSS pad and Energiya-Buran pad 37. It was only afterwards, on 8 April 1991, that the government issued a decree ordering NPO Energiya, KB Yuzhnoye, and KB Salyut to come up with competing proposals for boosters in the 25 to 40-ton payload range. This basically was a repeat of the order given in the 25 December 1984 decree, although in a somewhat lighter payload class. KB Salyut and KB Yuzhnoye had apparently also been optimizing their Proton and 11K37 designs. Eventually, on 6 July 1991 the Ministry of General Machine Building opted for Energiya-M. Between 1991 and 1993 preparations were made for starting production of flight models.

During that period, NPO Energiya worked out plans to launch a 30-ton space – plane (OK-M2) atop the rocket and also to turn the two strap-ons into reusable flyback boosters, something which appears to have been studied as early as 1989. Another idea was to launch the rocket from an ocean-based platform near the equator. This would not only allow Energiya-M to loft heavier payloads, but would also resolve the political problems associated with flying it from Baykonur, which became foreign territory after the collapse of the Soviet Union. One exotic mission considered for the ocean-launched Energiya-M was to deposit radioactive waste into heliocentric orbits, eliminating the risks involved in launching such dangerous payloads over populated territories. These studies formed the basis for the creation of the international Sea Launch venture, which would eventually use the three-stage Zenit rocket.

Despite the fact that Energiya-M used existing hardware and infrastructure and outperformed rockets like the Titan-4 and Ariane-5, it was ahead of its time. At the time there was simply no demand for the types of satellites that the rocket could place into orbit. On 15 September 1992 the Russian government started yet another com­petition to develop a family of even lighter rockets, which would eventually evolve into the Angara series. By late 1993 government funding for Energiya-M was stopped, with Russian Space Agency officials stating there was no demand for the rocket on the market. The following year NPO Energiya made an ultimate attempt to attract Western customers to Energiya-M and other Energiya variants, but to no avail. The prototype Energiya-M still stands today inside the Dynamic Test Stand at Baykonur [66].

The plans underwent further changes with the inception in the mid-1980s of the more capable An-225 Mriya carrier aircraft. Although conceived in the first place to transport Buran and elements of the Energiya rocket from the manufacturers to the Baykonur cosmodrome, designers may have had air-launch capability in the back of their minds from the outset.

The Mriya-based system was dubbed the Multipurpose Aerospace System (Mnogotselevaya aviatsionno-kosmicheskaya sistema or MAKS). The rocket was now replaced by an expendable external fuel tank (VTB), perched on top of which was either a reusable spaceplane (MAKS-OS) or an expendable unmanned cargo canister (MAKS-T). Also envisaged was a fully reusable unmanned winged cargo carrier with integrated propellant tanks (MAKS-M).

The OS was a 26-ton, two-man spaceplane with a length of 19.3 m, a height of 8.6 m, and a wingspan of 8.6 m. As on Spiral and BOR, the wings could be folded back for re-entry. The thermal protection system was the same as that of Buran, although a different material was needed for the much thinner wing leading edges. Behind the crew compartment was a 2.8 x 6.8 m payload bay. The original plan was for the spaceplane to have three Kuznetsov NK-45 LOX/LH2 main engines with a vacuum thrust of 90 tons each. That idea was turned down in favor of a tripropellant LOX/LH2/kerosene engine called RD-701, developed at NPO Energomash on the basis of the RD-170. Although this lowered the mean specific impulse, it still resulted in better performance because the external tank became much lighter by reducing the amount of liquid hydrogen, which is a low-density fuel taking up a lot of volume.

The RD-701 is a twin-chambered, staged, combustion cycle engine. Each chamber has a pair of turbopumps. One pump processes liquid oxygen and kerosene, which is turned into an oxygen-rich gas at 700 atmospheres after passing through a preburner. The other pump feeds liquid hydrogen to the main combustion chamber at ambient temperatures. The RD-701 has two modes of operation, combining first and second-stage engine characteristics in one package. During the initial phase it burns 81.4 percent liquid oxygen, 12.6 percent kerosene, and 6 percent liquid hydro-

gen, producing a total thrust of 400 tons with a specific impulse of 415 s. Then it switches to a combination of just liquid oxygen and hydrogen, with the thrust decreasing to 162 tons, but the specific impulse climbing to 460 s, helped by the deployment of a nozzle extension.

A typical MAKS-OS launch profile would see Mriya climb to an altitude of 9 km and assume the proper pre-launch attitude. The spaceplane would then ignite its main engine while still riding piggyback on the aircraft, making it possible to check its performance before separation. Some ten seconds later the 275-ton combination of spaceplane and external tank would be released from the Mriya to begin the trip to orbit. The engines would shut down before the spaceplane reached orbital velocity, allowing the external tank to burn up over the ocean across the world some 19,000 km from the launch point. The OS would then perform two burns of its two hydrogen peroxide/kerosene orbital maneuvering engines to place itself into orbit.

The basic version of the OS was designed to launch and retrieve small and medium- size satellites. Payload capacity was 8.3 tons to a 200 km orbit with a 51° inclination and 4.6 tons back to Earth. For space station missions there were two configurations. In one of them (TTO-1) the payload bay would house a small pressurized module capable of carrying four passengers plus cargo. This would be used for crew rotation or rescue missions, although the latter required additional fuel supplies for quick maneuvering. In the other (TTO-2) the payload bay would remain unpressurized and carry structures such as solar panels, antennas, or propellant tanks for refueling a space station. In both configurations a docking adapter was installed just behind the crew compartment. Also considered was an unmanned OS without a crew compart­ment and with a slightly enlarged cargo bay to fly heavier payloads (9.5 tons into a 200 km, 51° inclination orbit). Before committing MAKS-OS to flight, NPO Molniya planned to fly a suborbital unmanned demonstrator (MAKS-D). This would have the same size and shape as the OS, but would be equipped with a single RD-120 Zenit second-stage engine fed by propellant tanks in the payload bay.

In the MAKS-T configuration the OS was replaced by an unmanned cargo canister equipped with an RD-701 tripropellant engine and an upper stage for inserting payloads into the proper orbit. Maximum payload capacity was 18 tons to a 200 km, 51° inclination orbit, and 4.8 tons to geostationary orbit. For geo­stationary missions the Mriya would fly to the equator to make maximum use of the Earth’s eastward rotation and be refueled in flight.

MAKS-M was a fully reusable, unmanned, winged cargo container with integrated propellant tanks, designed to deliver payloads to low orbits (5.5 tons to a 200 km, 51° orbit). Situated in between the propellant tanks was a cargo bay slightly larger than that of the OS. An earlier version of this (VKS-D) had the cargo bay on top of the propellant tanks. NPO Molniya designers even floated the idea of transforming VKS-D into a suborbital intercontinental passenger plane capable of carrying 52 passengers to any point on the globe within 3 hours at a price of $40,000 per ticket.

NPO Molniya had plans to further upgrade the MAKS system by fitting the Mriya with more powerful NK-44 engines and eventually by replacing Mriya with a giant twin-fuselage triplane called Gerakl (“Heracles’’) with a phenomenal 450-ton cargo capacity. A similar plane had already been studied for air launches in the early 1980s under the name System 49M. In the even more distant future the hope was to finally realize the old Spiral dream by developing an air-launched system based on a hypersonic carrier aircraft (VKS-G) [6].

The RLA concept fitted well in a new philosophy to replace the existing fleet of Soviet launch vehicles by a new generation of rockets. By the early 1970s the Soviet Union was operating five basic types of fundamentally different launch vehicles each derived from a specific intermediate or intercontinental ballistic missile: the Kosmos and Tsiklon rockets (based on the R-12, R-14, and R-36 missiles of the Yangel bureau), the Vostok/Soyuz family (based on the R-7 missile of the Korolyov bureau) and the Proton (based on the UR-500 missile of the Chelomey bureau). While the R-7 based rockets used LOX/kerosene as propellants, all the others relied on storable hypergolic propellants.

Around the turn of the decade plans were being drawn up for new generations of satellites with more complex on-board equipment and longer lifetimes, requiring the use of heavier, more capable launch vehicles. By early 1973 a research program called Poisk (“Search”), conducted by the Ministry of Defense’s main space R&D institute (TsNII-50), had concluded that future satellites should be divided into four classes: “light” satellites up to 3 tons, “medium-weight” satellites up to 10-12 tons, “heavy” satellites (up to 30-35 tons), and “super-heavy” satellites (not specified). The last three classes were not served by the existing launch vehicles.

The new family of launch vehicles was to have two key characteristics. First, in order to cut costs to the maximum extent possible, it would use unified rocket stages and engines. Second, it would rely on non-toxic, ecologically clean propellants, with preference being given to liquid oxygen and kerosene. The reasoning behind this reportedly was that “the number of launches of [space rockets] would be much higher than the number of test flights of nuclear missiles with [storable] propellants.” What also may have played a role were a series of low-altitude Proton failures that had contaminated wide stretches of land at or near the Baykonur cosmodrome. The basic conclusions of the study were approved on 3 November 1973 at a meeting of GUKOS [45]. Although not stated specifically, the long-range goal of this effort seems to have been to phase out all or most of the existing missile-derived launch vehicles.

Initially, the primary focus apparently was on unifying the light to heavy class of rockets, because at the time these decisions were made super-heavy rockets and reusable shuttles were still a distant and vague goal. It would seem that three design bureaus were ordered to come up with proposals for such a family of launch vehicles under a competition called Podyom (“Lift”). Both Branch nr. 3 of TsKBEM in Kuybyshev (which became the independent TsSKB in July 1974) and Chelomey’s TsKBM put forward plans to refit their respective Soyuz and Proton launch vehicles with the Kuznetsov bureau’s LOX/kerosene NK engines originally built for the N-1 rocket. The former Yangel bureau in Dnepropetrovsk (now called KB Yuzhnoye and headed by Vladimir Utkin) also weighed the idea of using NK engines (which after all were around and had been tested), but in the end favored the new generation of LOX/ hydrocarbon engines being designed by Energomash [46].

Actually, by this time Yuzhnoye had been working for several years on a new medium-lift launch vehicle (11K77) burning storable propellants. In December 1969 it had been ordered by GUKOS to develop a new rocket capable of placing 8 tons into low polar orbits and 2 tons into highly elliptical Molniya-type orbits. The initial idea in 1970 was to build a three-stage rocket with 3.6 m diameter modules. The following year attention turned to a rocket derived from the bureau’s R-36M, a new ICBM that had been approved by a government decree in September 1969. By retaining the 3.0 m diameter of the R-36M’s rocket stages, no fundamentally new manufacturing techniques would be required. The 11K77 would now consist of a pair of R-36M first and second stages stacked on top of one another plus a newly developed third stage. In 1972 Yuzhnoye also designed a slightly less powerful rocket designated 11K66, essentially a two-stage R-36M with increased propellant load capable of orbiting 5.9 tons.

When Yuzhnoye made the switch to LOX/kerosene under the Podyom program in 1974, it opted for a vehicle maintaining the 3.0 m diameter of the rocket stages, but employing parallel staging. This version of the 11K77 would fly the single-chamber LOX/kerosene engines then being designed at Energomash. The second stage, acting as the core, would have a 130-ton thrust RD-125 engine, and the two first-stage modules flanking the core would each have three 113-ton thrust RD-124 engines. These engines had combustion chambers of roughly the same size and pressure as those of the R-36M first-stage engine. Payload capacity to low orbit was now 12 tons.

By early 1975 Energomash’s Department 728 had made enough progress on the powerful 600-ton thrust RD-123 for Yuzhnoye to incorporate it into the 11K77 design. This made it possible to replace the two modules of the 11K77’s first stage by a single module, although its diameter would have to be increased to 3.9 m, which was the maximum that could be transported by rail. The second stage would now be placed on top the first stage, turning the 11K77 into what the Russians call a “mono­block’’ booster. After a convoluted development path, the rocket had now acquired the configuration in which it later became known to the world as Zenit. It could also serve as the basis for a lighter rocket (11K55) and a heavier version (11K37), with the three covering the whole payload range from “light” to “heavy” (see Chapter 8) [47].

Being in the heavy to super-heavy class, Glushko’s RLA family was not part of the Podyom competition, but as plans for a large Soviet shuttle gained more support in 1975, so did the idea of unifying the first stage of the RLA rockets with that of the future medium-lift rocket. It is unclear if this consideration played a significant role early on in the Podyom competition. If it did, TsSKB and TsKBM had betted on the wrong horse by anticipating that whatever followed the N-1 would also carry NK engines, while Yuzhnoye had made the right choice by picking the new Energomash engines. However, the history of the 11K77/RLA unification is a complicated chicken-and-egg story, with sources differing on whether the initiative to unify the first stages came from Glushko or Utkin. At any rate, Yuzhnoye emerged as the winner of the Podyom competition in 1975, no doubt because its medium-lift rocket could now act as a test bed for the first stage of the rocket that would power the Soviet space shuttle to orbit.

The 11K77 is the only thing that ever came out of Podyom. Yuzhnoye’s proposed light-class and heavy-class rockets never flew and the whole idea of a standardized fleet of dedicated, environmentally clean space launch vehicles in the light to heavy class remained a distant dream, probably because it infringed too much on design bureau interests. In fact, most of the IRBM and ICBM-derived launch vehicles conceived in the 1960s continue to fly today, although another attempt is being made to develop a standardized rocket fleet under the Angara program (see Chapter 8).

The Energiya “Blok-A” strap-on boosters were largely designed and built by KB Yuzhnoye and its associated production facility in Dnepropetrovsk, although major elements were also supplied by NPO Energiya in Kaliningrad. Each strap-on booster was 39.4m high and had a maximum diameter of 3.9 m, dictated by railway trans­portation requirements. The wet mass was about 372 tons. The booster consisted of a nose section, an upper LOX tank, an intertank structure, a lower kerosene tank, and

a tail section with the gimbal-mounted LOX/kerosene RD-170 engine. The four boosters were numbered 10A, 20A, 30A, and 40A.

The nose section housed the avionics bay, which among other things contained an M4M computer that interacted with the core stage’s M6M central computer. Also installed in the nose section were flight data recorders, mounted in a special casing that protected them from the force of impact. Two of the four Blok-A boosters were equipped with radio beacons, enabling ground controllers to follow their trajectory after separation from the core stage.

The propellant tanks were made of an aluminum-magnesium alloy, with the walls being about 30 mm thick. The LOX and kerosene tanks had useful volumes of 208 m3 and 106 m3, respectively, holding approximately 220 tons of LOX and 80 tons of kerosene. The upper LOX tank fitted in a concave depression at the top of the kerosene tank and the LOX feed line passed right through the middle of the lower kerosene tank. There were vent and relief valves in the forward domes of both tanks. The kerosene fill and drain valves were in the aft dome of the kerosene tank, and the LOX fill and drain valves in the lower part of the LOX feed line. The LOX feed line also had a damper to suppress “pogo” oscillations.

Embedded in the lower part of the LOX tank were two rows of helium tanks for in-flight pressurization of both the LOX and kerosene tanks. The helium for the kerosene tank was supplied directly, while that for the LOX tank first passed through a heat exchanger in the lower engine compartment. Before launch the tanks were pressurized with ground-supplied helium. Electrical power for the boosters was provided by batteries.

There was a pair of strap-ons on either side of the core stage. Each pair was mechanically linked and jointly separated from the core stage, with the two boosters not splitting until about half a minute later. This was done in order to minimize the risk of any of the boosters hitting the orbiter after separation. The boosters were pyrotechnically separated from the core stage and subsequently activated 11 small solid-fuel motors (seven on the nose section and four on the tail section) to ensure safe separation from the core stage and payload. They came down some 425 km downrange from the launch site.

Although they were mechanically linked, the strap-ons operated independently from one another. The only electrical connections were with the core stage (one interface with 408 contacts for each booster). There were also twelve electrical, pneumatic, and hydraulic connections between each strap-on and the launch pad (eight for propellant, fluid, and gas supply and four electrical connections). The connections were between the nose section and the launch tower and the tail section and the launch table.

From the very beginning of the Energiya-Buran program the idea was that the strap-on boosters would be reusable. The degree of reusability mainly depended on the robustness of the RD-170 engine, which was certified to fly at least 10 missions. After having studied several schemes, designers opted for a horizontal landing system using parachutes, soft-landing engines, and a set of shock absorbers. Parachutes would be deployed from the forward and aft ends of each booster, orienting it to a horizontal attitude for descent. The plan had much in common with the landing

technique for the giant MTKVP lifting body that had been studied before the delta­wing concept of Buran was picked.

The strap-ons were designed from the beginning with special containers in their nose and tail sections to house the parachutes, other recovery systems, and control equipment. The containers were installed on the strap-ons flown on the two Energiya missions in 1987 and 1988 (6SL and 1L), but were loaded with instrumentation rather than recovery equipment. However, there were plans to demonstrate the recovery technique on the 2L mission with the GK-199 payload (see Chapter 8).

The boosters shared about 70 percent of their systems with the first stage of the Zenit rocket. The part of the booster that was largely similar to the Zenit first stage was known as the “modular part” (or 11S25) and included the propellant tanks, pressurization systems, and the engine compartment. The main differences with the Zenit first stage were in the gimbal axes of the engines and also in that the tank walls were slightly thicker because of the bending loads imposed by the strap-on config­uration. Elements unique to the strap-ons included the aft skirt surrounding the engine compartment (which needed to be compatible with the Energiya core stage), the entire nose section, and the parachute containers.

The original requirement was for the Zenit first stage to be reusable as well, but, if it was to use the same recovery systems and recovery zones as the Blok-A, the Zenit would need to have the same speed at the moment of first-stage separation as Energiya (1,800 m/s instead of 2,500 m/s). This would have made it necessary to make the second stage 38 tons heavier, reducing the rocket’s payload capacity to

an unacceptable 7 tons. A later idea to recover only the tail section with the engine was dropped as well because this was expected to produce a return on investment after only about 500 launches [3].

The Electric Power System (SEP) supplied power to Buran’s systems during the final countdown, the mission itself, and during initial post-landing servicing. As on the Shuttle Orbiter, electricity was to be generated with the help of fuel cells (“electro­chemical generators’’ in Russian terminology) using cryogenically stored oxygen and hydrogen reactants. Whereas NASA introduced fuel cells back in the Gemini program, the Russians had always used battery systems and/or solar panels on Vostok, Voskhod, and Soyuz. They did develop a fuel cell system called Volga-20 for the Soyuz-based LOK lunar orbiting ship to be used in the N-1/L-3 manned lunar-landing program, but the only LOK ever flown was lost in the fourth and final launch failure of the N-1.

The SEP consisted of the Oxygen/Hydrogen Cryostats, a Power Module, an Instrument Module, and the Distribution and Commutation System. The first three subsystems were situated in the mid fuselage under the front section of the payload bay, so that receding fuel levels in the cryogenic tanks would not affect Buran’s center of gravity. Although the oxygen and hydrogen were delivered to the fuel cells in gaseous form at a temperature of about 10° C, they were stored cryogenically to save mass. Buran could accommodate two oxygen and two hydrogen tanks, which needed to be filled in the final days before launch via 500 x 600 mm doors in the mid fuselage.

The oxygen and hydrogen were fed to the Power Module, which contained the actual fuel cells. There were a total of four fuel cell units (as compared with three on the Shuttle Orbiter), each consisting of eight 32-cell stacks connected in parallel and with an active electrode area of 176 m2. The alkaline fuel cells used a potassium hydroxide electrolyte immobilized in an asbestos matrix and had an oxygen electrode (cathode) and a hydrogen electrode (anode). Each fuel cell unit provided 10 kW continuous and 25 kW peak at between 29 and 34 volts of direct current. Only three

sets of fuel cells were needed for a nominal mission and two for an emergency landing.

The Instrument Module turned the fuel cells on and off and automatically controlled all processes taking place in the system. In case it detected an anomaly, the crew was notified of this on the control panels in the cockpit and with a master alarm. Although the fuel cells were designed to operate entirely automatically, they could also be controlled by the crew or from the ground. Power was distributed to all parts of the vehicle by the Distribution and Commutation System, which consisted of two redundant subsystems, one running along the starboard side, and the other along the port side.

The fuel cells produced water as a byproduct (more than 100 kg per day) for consumption by the crew and also for use in the flash evaporators of the Thermal Control System and the hydraulic system. The liquid oxygen stored in the SEP tanks could also be turned into gaseous oxygen for the crew compartment.

For extended missions, Buran could carry a “cryo kit” located near the middle of the payload bay and equipped with up to six liquid hydrogen tanks. During a long mission the fuel cells would first use the hydrogen supply from the cryo kit before switching to the standard LH2 tanks under the payload bay. Extra oxygen would be drawn from the LOX tank of Buran’s propulsion system situated in the aft fuselage. Buran’s cryo kit was comparable with that developed for the Shuttle’s Extended Duration Orbiter missions, although that was to be mounted in the aft payload bay and had both liquid hydrogen and liquid oxygen tanks (given the use of storable rather than cyrogenic propellants in the on-orbit propulsion system).

In addition to the fuel cells, Buran also had battery packs that were charged by the fuel cells and fed electricity to various power-hungry systems, mainly in the aft fuselage. For the first multi-day test flights it was also planned to fly battery packs operating independently from the fuel cells to give one day of back-up power in case of a fuel cell failure, enough to make an emergency return to Earth. Since Buran’s one and only mission lasted just several hours, the fuel cells were not installed, with the vehicle’s systems drawing power from batteries in the BDP payload stowed in the payload bay (see Chapter 7). Fuel cells were installed on the second vehicle and underwent loading tests at the launch pad.

Called Foton (“Photon’’), Buran’s fuel cells were developed jointly by NPO Energiya and the Ural Electrochemical Integrated Factory in Verkh-Neyvinsk (Sverdlovsk region), which had also developed the Volga-20 fuel cells for the LOK back in the early 1970s. Although never actually flown in space, the Buran fuel cells attracted the interest of the European Space Agency, which tested a Buran flight – model fuel cell in 1993 at the facilities of ESTEC in Noordwijk, Holland as part of studies to incorporate foreign technology into the Hermes spaceplane. A modified version of Foton powered the first Russian fuel cell car, the Niva, presented at a Moscow auto show in 2001. RKK Energiya is also considering a Foton-derived fuel cell system for its new Kliper spacecraft [20].

The single mission flown by Buran on 15 November 1988 was not a standard flight. It was flown without a crew on board and with the sole intention of testing the launch and re-entry procedures. No major on-orbit tasks were scheduled and Buran flew without many of the systems that would have been required for a multi-day manned mission. What will be described here are the standard launch and landing procedures and standard on-orbit operations for operational missions with a crew on board. Details of actually planned missions will be given in Chapter 8.

Launch

The launch began with the ignition of the core stage’s four RD-0120 engines at T — 9.9 seconds, followed at T — 3.7 seconds by the ignition of the strap-on rockets’ RD-170 engines. The interval was required to allow the core stage engines to slowly build up thrust and thereby ease the acoustic loads on the orbiter. If an anomaly was detected by the rocket’s flight control system, all engines could be shut down at any moment prior to T — zero.

As the stack cleared the tower, it performed a pitch and roll maneuver to place it in the proper attitude for the remainder of the ascent. About half a minute into the flight the core stage and strap-on engines were throttled back to minimize aero­dynamic pressures and longitudinal loads on the vehicle. After passing through the densest layers of the atmosphere, all engines were throttled back up to nominal thrust, although the Blok-A RD-170 engines were soon again throttled down in preparation for shutdown. The four strap-on boosters shut down in pairs with an interval of 0.15 seconds and were jettisoned about two seconds later at T + 2m26s. They continued to fly in pairs, separating from one another somewhat later to come down some 425 km downrange. As mentioned earlier, the strap-ons could land on parachutes for recovery but were not configured as such on the two Energiya launches that were flown.

Moving on downrange, the core stage again began throttling down its four liquid oxygen/liquid hydrogen engines less than a minute before shutdown, which occurred at T + 7m47s. The engines were shut down in pairs with an interval of 0.2 seconds. Fifteen seconds later the orbiter separated from the core stage and safely maneuvered itself away with gentle burns of its primary thrusters. The core stage then continued on a ballistic trajectory to burn up over the Pacific Ocean. Not having required orbital velocity yet, Buran then needed two burns of one of its DOM engines about 11 and 40 minutes into the flight to place itself into an initial orbit. The required burn duration was calculated by the on-board computers on the basis of the launch vehicle’s performance. The maximum acceleration forces for the crew during launch would not have exceeded 3g.

The crew had no active role to play during the launch phase and merely had to monitor the operation of on-board systems on their cockpit displays. The orbiter’s computers automatically controlled the operation of the life support, thermal con­trol, power, and monitoring systems as well as that of the hydraulic systems and Auxiliary Power Units, which might be needed in a launch abort to perform an emergency landing. They also opened and closed the vehicle’s vent doors at the required moments [28].

The Energiya core stage’s RD-0120 engines were designed by the Design Bureau of Chemical Automatics (KB Khimavtomatiki or KBKhA) in Voronezh. This bureau was founded in 1941 by Semyon A. Kosberg in the city of Berdsk and was transferred to Voronezh as OKB-154 in 1946. Kosberg headed the bureau until he was killed in an automobile accident in 1965 and replaced by Aleksandr D. Konopatov, who remained in charge of the bureau until 1993. The bureau developed engines for surface-to-air missiles, submarine-launched and intercontinental ballistic missiles, and entered the space business in the late 1950s with the development of upper-stage engines for R-7 derived launch vehicles. It also designed the second and third-stage engines for the Proton rocket. KBKhA was a newcomer to the development of cryogenic engines when it was assigned to develop the RD-0120. Chief designer of the RD-0120 was Vladimir S. Rachuk, who would go on to become the general designer of KBKhA in 1993.

Actual manufacturing of the RD-0120 engines took place at the Voronezh Machine Building Factory (VMZ), located on the same premises as KBKhA.

Founded in 1928, VMZ switched to the production of rocket engines in 1957, building all the engines designed at KBKhA [5].

Describing the history of the Buran cosmonaut team is not as straightforward as it may seem at first glance. Unlike the situation in the US, where NASA has always been in charge of selecting and training the (career) astronauts that make up Space Shuttle crews, the Soviet Union’s space program lacked a central coordinating NASA-type organization. Several organizations involved in test pilot and cosmonaut training felt they should all independently select their own cosmonaut teams. From these groups, Buran crew members representing those organizations would be assigned.

Three organizations selected cosmonauts specifically for Buran:

• The Cosmonaut Training Center named after Yu. A. Gagarin (TsPK for Tsentr Podgotovki Kosmonavtov) based in Star City (Zvyozdnyy Gorodok) near Moscow.

This Soviet Air Force unit, set up in I960, had been in charge of selecting and training cosmonauts for flights on Vostok, Voskhod, Soyuz, and Salyut.

• The Flight Research Institute named after M. M. Gromov (LII for Lyotno – Issledovatelskiy Institut) in Zhukovskiy, some 35 km southeast of Moscow.

The Flight Research Institute, a civilian research and development entity subordinate to the Ministry of the Aviation Industry (MAP), was founded in 1941 as the leading test center for experimental and production aircraft. LII had a Test Pilot School (ShLI).

• The State Red Banner Scientific Test Institute named after V. P. Chkalov (GKNII for Gosudarstvennyy Krasnoznamennyy Nauchno-Ispytatelnyy Institut) in Akhtubinsk, some 130 km west of Volgograd and about 50 km south of the Kapustin Yar cosmodrome in the Volga delta.

This Air Force unit was set up in 1960 at the same site that had served since 1947 for testing various unmanned flying apparatuses, such as surface-to-surface missiles, air-to-surface missiles, and the Burya intercontinental cruise missile.

The site was sometimes referred to as Vladimirovka, after a nearby railway station. With the establishment of GKNII its role was expanded to testing various aircraft for the Air Force and it also included an Air Force test pilot school known as the Test Pilot Training Center (TsPLI).

In March 1979 MOM, MAP, and the Ministry of Defense jointly decided that a pool of 17 pilots would be required for the Buran test flight program: six from LII, six from GKNII, and five from TsPK [1]. Part of the reason for assigning pilots from three different organizations was no doubt the departmentalism typical of the Soviet space program. However, there appear to have been more rational considerations as well. Since Buran was far more complex than any Soviet spacecraft flown before, the unmatched flying skills of the LII pilots were probably considered necessary to safely guide the vehicle through its initial atmospheric and orbital flight tests, with GKNII becoming involved in the test flights at a somewhat later stage. It was not uncommon in the former Soviet Union for the Air Force team in Akhtubinsk to further test new aircraft once they had been declared airworthy by the LII pilots, and in this respect Buran was no exception [2]. Presumably, the LII pilots were to fly test flights with civilian payloads, and the GKNII pilots test flights with military payloads.

Once their job was completed, these career test pilots would then return to their usual line of work, passing the torch to the “regular” TsPK pilots to finish the test program, and ultimately fly Buran’s operational missions. This, at least, appears to have been the original intention when the first pilot teams were selected in the 1970s and Buran was expected to begin flying in the first half of the 1980s. As the orbital test flights slipped into the late 1980s and were spread out over many years, the opera­tional phase became a distant and vague goal. Therefore, in the end the only pilots seriously considered to fly on Buran were from LII and GKNII, and further TsPK selections were solely aimed at the mainstream Salyut/Mir space station program.

In addition to the pilots, engineers from both NPO Energiya and the Air Force were assigned to Buran as well, although none of them ever appear to have been specifically selected for the program. Because of all this, it is not possible to really give one single founding date for the Buran cosmonaut team.

Before the beginning of the Energiya-Buran program the only liquid oxygen/liquid hydrogen engines developed in the Soviet Union had been the 7.5-ton thrust 11D56 of KB Khimmash and the 40-ton thrust 11D54 and 11D57 of KB Saturn, intended to be used on upper stages of the N-1 rocket. Therefore, the development of the 190-ton thrust RD-0120 for Energiya’s core stage, assigned to KBKhA in Voronezh, was a major challenge, partly because much of the infrastructure needed for testing the engine was not yet in place. This required a step-by-step approach to certifying the engine for flight. The first step was to test individual components of the engine, followed by test firings of experimental engines at increasing rates of thrust, and ultimately test firings of flightworthy engines both individually and mounted in clusters of four on Energiya’s core stage at Baykonur’s UKSS pad.

Since the RD-0120 was the first powerful Soviet cryogenic engine, many com­ponents that would usually be installed straightaway in experimental engines now had to be tested individually first. Particular attention was paid to the ignition system in the combustion chamber and the gas generator as well as to the operation of the low-pressure turbopumps. The components for these tests were built both by KBKhA’s own Experimental Factory and by the Voronezh Machine Building

Factory and tested at various locations in the Soviet Union, including KBKhA’s own experimental base in Voronezh.

While KBKhA had its own rocket engine test stands near Voronezh, none of these could be converted for test firings of this radically new engine, nor was it deemed expedient to build the new facilities there. Instead, test firing stands for the RD-0120 were built by two other organizations. One was the Scientific Research Institute of Machine Building (NII Mashinostroyeniya or NIIMash) in Nizhnyaya Salda north of Yekaterinburg. This organization had become independent in 1981 after having been a branch of the Scientific Research Institute of Thermal Processes (NII TP), which was how the former NII-1 was renamed in 1965 (in 1995 it was again renamed as the Keldysh Research Centre). NIIMash mainly specialized in reaction control thrusters for satellites and manned spacecraft (including Buran).

NIIMash had two test stands (nrs. 201 and 301) for firing the RD-0120 in vertical position. Construction got underway in 1977 and 1981, respectively. The test stands allowed the engines to be gimbaled and had diffusers to simulate operating conditions at higher altitudes. Stand nr. 301 was capable of handling longer duration test firings (over 1,000 seconds) and maximum gimbal angles. RD-0120 engines were delivered to Nizhnyaya Salda from the manufacturing plant in Voronezh by Antonov-8 cargo planes.

The other organization was NIIkhimmash near Zagorsk. It had a cryogenic complex called KVKS-106, already used before the Energiya-Buran program for testing the 11D56 and 11D57 cryogenic engines. KVKS-106 comprised five test facilities, one of which had two test stands (V-2A and V-2B) for full-scale test firings of the RD-0120 in horizontal position. Others were used to test individual components of the engine and the core stage. The engines were transported from Voronezh to NIIkhimmash by road.

The bulk of the testing was to be carried out at Nizhnyaya Salda, but, as the construction of the facilities there ran into delays, it was decided to test the first engines at the existing cryogenic complex of NIIkhimmash. The first test engine (with

a shorter-than-nominal nozzle) was delivered to Nllkhimmash in the autumn of 1978 and underwent a first brief test firing (4.58 seconds) at the V-2B stand in March 1979. The initial tests at Nllkhimmash were conducted at just 25 percent of rated thrust, but allowed to test such things as the ignition and shutdown sequence. The adjacent V-2A, capable of supporting tests at 100 percent thrust, was ready in 1984.

The first test firing at NIIMash’s test stand nr. 201 took place on 19 January 1980, also at low thrust. By early 1981 the engine was being tested at 70 percent thrust, and it wasn’t until May 1984 that the RD-0120 worked at 100 percent thrust for 600 seconds (with the nominal operating time during launch being 467 seconds). Considerable delays in reaching nominal thrust were caused by problems with the impeller of the liquid hydrogen pump, a rotating disk with a set of vanes that produces centrifugal force within the pump casing. The problem was eventually solved by using a different type of titanium to manufacture the component.

Although the RD-0120 test program took longer than expected, it does not appear to have been a major factor in the delays of Energiya’s first launch. By the time of that maiden launch in May 1987 the Russians had accomplished 523 test firings of 103 different RD-0120 engines with a total duration of 73,891 seconds. NIIMash’s test stand nr. 301 saw its first test firing on 30 July 1987. Records set during that same period were a 100-second test at 123 percent rated thrust in September 1987 and a maximum-duration burn of 1,202 seconds in January 1988. By early October 1988, about a month before the maiden flight of Buran, a total of 126 engines had undergone 635 test firings lasting a total of 120,454 seconds. By comparison, NASA accumulated 110,253 seconds of burn time on the Space Shuttle Main Engines in 726 test firings before STS-1.

The RD-0120 test program continued at Nllkhimmash even after the cancella­tion of the Energiya-Buran program under a deal between KBKhA and the US Aerojet company to study the feasibility of using the engine on future American launch vehicles. Several test firings were also performed in the mid-1990s in the framework of the joint European-Russian RECORD program (Russia-Europe Cooperation on Rocket Engine Demonstration), led by the French SEP company. In those tests, the engine was equipped with European instrumentation to allow European engineers to create a detailed software model of the engine and gain experience with staged-combustion cycle cryogenic engines for possible use in future reusable systems. The final test firing of the RD-0120 was conducted at Nllkhimmash in 1997. The maximum accumulated test time for a single engine was about 5,000 seconds and the maximum number of ignitions for a single engine was thirty [2].

A typical BOR-4 mission would begin with a launch from the Kapustin Yar cosmo­drome near Volgograd using a modified two-stage Kosmos-3M booster known as K65M-RB5. Baykonur no longer supported that rocket at the time and although Plesetsk did it was situated too far north to place the spaceplane models into the proper inclination. The rockets used for the BOR-4 missions had originally been earmarked for other missions, but had already exceeded their guaranteed “shelf life’’ and would have been used for test launches anyway.

The vehicle would be launched with its two wings completely folded so that it fitted under the rocket’s fairing. After release from the launch vehicle, the wings were unfolded to a position that would keep the vehicle stable during re-entry at an angle of attack of between 52° and 57° between altitudes of 70 and 60 km. Orientation in orbit was carried out with the help of eight microthrusters burning hypergolic propellants. After a single revolution of the Earth, BOR-4 initiated its descent back to Earth, firing what is believed to have been a jettisonable solid-fuel motor mounted on top of the vehicle. At an altitude of 30 km the on-board control system sent BOR-4 on a steep spiralling trajectory to decrease speed and at 7.5 km the spacecraft deployed a parachute that reduced the vertical landing speed to 7-8 m/s.

Since the vehicle was not equipped with landing gear, it needed to land on water to ensure that its heat shield remained intact for post-flight analysis. The only major bodies of water on Soviet territory that would be in the BOR’s flight path were the Black Sea and Lake Balkhash. However, the Russians had never returned a winged vehicle or lifting body from orbit and were not confident they could aim the space­craft for precision splashdowns in the Soviet Union. Therefore they opted to land the first vehicles in the Indian Ocean, where they would still come down in water even if they fell short of or overshot the planned landing area. That did, however, signifi­cantly increase the cost of the recovery operations, which, moreover, would be hard to conceal from the prying eyes of Western reconnaissance aircraft.

After splashdown a conically shaped float was inflated on top of the spacecraft to improve its buoyancy. The float also had flashing lights and antennae to make it easier for the recovery forces to locate the BOR. Before being hoisted on board a recovery ship, a crew was sent out to the vehicle to disarm an on-board self-destruct system.

While BOR-4 was designed in Zhukovskiy under the leadership of LII chief Viktor V. Utkin, the vehicle was manufactured and covered with heat-resistant materials at the Tushino Machine Building Factory. The man in charge of the BOR-4 program at NPO Molniya was Stepan A. Mikoyan, a deputy of chief designer Gleb Lozino-Lozinskiy. The BOR-4 test flights were coordinated by a State Commis­sion headed by former cosmonaut Gherman Titov, then serving as a deputy head of GUKOS. Titov, incidentally, had also been part of the Spiral cosmonaut training group at Star City in the 1960s.

The orbital flights were preceded by an experimental suborbital launch on 5 December 1980 in the direction of Lake Balkhash. This mission was intended to test the rocket, the aerodynamic characteristics of the vehicle, and the performance of the aerodynamic surfaces and rocket thrusters. Designated BOR-4S (serial nr. 401), the vehicle only had the original ablative heat shield. The final part of the flight was monitored by two Ilyushin 18RT aircraft flying in the vicinity of Lake Balkhash. These were modified Ilyushin 18D aircraft specially adapted to perform tracking in areas that were not covered by Soviet ground-based or sea-based tracking means.

In the spring of 1982 seven Soviet ships set sail for the Indian Ocean to support the first orbital mission of a BOR-4 vehicle. These included two vessels to ensure communications between the fleet and the home front, namely the Navy’s Chumikan and the Academy of Sciences’ Kosmonavt Georgiy Dobrovolskiy. In no time Royal Australian Air Force P-3C Orion reconnaissance aircraft deployed from RAAF Base Williams in Point Cook were circling overhead to monitor the vessels’ activities.

Finally, on 3 June 1982 the first BOR-4 covered with Buran’s heat shield materials (serial nr. 404) was successfully placed into orbit. After a single orbit the vehicle fired its deorbit engine and re-entered the Earth’s atmosphere, performing a cross-range maneuver that took it about 600 km to the south of its orbital path. The craft gently splashed down some 560 km south of the Cocos Islands, which was about 200 km from its intended landing point. The recovery operation was seriously hampered by stormy seas. Battling high waves, recovery forces on board the vessel Yamal needed several attempts to hoist the BOR-4 on deck. During one of those

attempts, the vehicle accidentally bumped into the Yamal, causing significant damage to the spacecraft’s nose section. The entire operation was photographed by an Australian Orion aircraft, which according to Russian eyewitnesses flew so low that the slipstream nearly knocked them off their feet.

The TASS news agency issued a routine statement saying a satellite called Kosmos-1374 had been launched for “the further study of outer space’’, providing no hint of its real mission. It only differed from the standard Kosmos launch announcement by adding that “the scientific research envisaged by the program had been carried out.’’ Within a week US media reports were suggesting the mission had been a test of a small shuttle vehicle, although some argued it had been a test of a prototype spaceborne nuclear weapon targeted on US and British naval forces in the Indian Ocean.

State Commission leader Gherman Titov, who had already pushed for a Black Sea landing on the first mission, now turned to the Military Industrial Commission with a request to have the next BOR-4 land in the Black Sea, expressing his fear the ship could be captured by the Americans. However, he was overruled by his superiors, possibly because Kosmos-1374 had landed well off target and a splash­down in the much smaller Black Sea could not yet be guaranteed. On 15 March 1983, Kosmos-1445 (serial nr. 403) was launched on a repeat mission, coming down 556 km south of the Cocos Islands. On station in the Indian Ocean apart from Navy vessels of the Black Sea fleet were the tracking ships Kosmonavt Vladislav Volkov and Kosmonavt Pavel Belyayev. Two Il-18RT tracking aircraft were in the skies over

Afghanistan to monitor the final part of the re-entry. Coming in the middle of the Soviet-Afghan war, their missions were not without risk and they were protected by a whole squadron of Soviet fighter jets.

Kosmos-1445 was again retrieved by the Yamal. Better prepared than during the Kosmos-1374 mission, the Australian Air Force once again sent out P-3C Orion aircraft to monitor the recovery operation and obtained even better pictures than before, some of which were released to the public by the Australian Ministry of Defense in April 1983. Also keeping a close eye on events were several Australian Navy vessels, which reportedly came so close that the Soviet crew members could use their binoculars to catch a glimpse of the movies shown on giant screens on the upper decks in the evening.

Confident enough now they could bring back the BORs with sufficient precision, the Russians decided to land the next two BORs in the Black Sea just west of Simferopol. The first was launched on 27 December 1983 as Kosmos-1517 (serial nr. 405) and the second was orbited on 19 December 1984 as Kosmos-1614 (serial nr. 406). The TASS launch announcements differed from the earlier ones in acknowl­edging that the satellites “had performed a controlled entry into the atmosphere and

landed in the pre-designated area of the Black Sea.” While Kosmos-1517 was successfully retrieved by the Yamal, it was later revealed that Kosmos-1614 was lost, having either burned up in the atmosphere or sunk in the Black Sea. The recovery vessels, aircraft, and helicopters searched the 70 x 30 km landing ellipse for about a week, but to no avail. Talking about the cause of the mishap many years later, State Commission leader Gherman Titov said that “while fixing one problem, engineers had created another.”

Despite the failure to recover the final vehicle, it was felt that enough data had been gathered during the four orbital flights that a fifth mission was reportedly canceled. The Russians later said the missions had allowed them to test the effects of aerodynamic, temperature, and acoustic loads as well as vibrations on the heat shield between altitudes of 100 and 30 km and speeds of between Mach 25 and Mach 3. Particularly helpful had been the temperature data obtained in critical areas such as the nosecap and the underbelly of the vehicle. The BOR-4 missions had helped to determine the ideal size of gaps between the tiles, measure the “catalytic activity” of the heat shield in real plasma conditions, and also to study the risks associated with losing one or more tiles. The flights had also made it possible to “outline measures to reduce the mass of Buran’s heatshield”, although there is no evidence those measures were actually implemented [17].