Category Energiya-Buran

On 26 October the State Commission in charge of test flights of the Energiya-Buran system met at Baykonur to set a launch date for the mission. Such commissions were set up routinely in the Soviet Union to oversee launch preparations for specific projects. The composition of the Buran State Commission reflected the importance attached to the flight. Established in December 1985, it was headed by none less than the Minister of General Machine Building himself, initially Oleg Baklanov, replaced in April 1988 by Vitaliy Doguzhiyev. In all, the commission numbered 44 people, including 9 ministers, 10 deputy ministers, the President of the Academy of Sciences, leading Ministry of Defense officials, and several general and chief designers (Glushko, Gubanov, Semyonov, Lozino-Lozinskiy, Konopatov, Radovskiy, Barmin, Andryushenko, and Lapygin). The “technical leader” of the State Commission, somewhat comparable with a “launch director” in the US, was NPO Energiya head Valentin Glushko. However, the 80-year old Glushko, who had suffered a stroke only months earlier, was recovering in a Moscow hospital and had to be replaced by his deputy Boris Gubanov. Final preparations at the pad were the responsibility of the military teams of the so-called 6th Test Directorate under the leadership of Major – General Vladimir E. Gudilin.

Despite last-minute concerns over problems with another ODU test firing near Leningrad on 19 October, the commission declared Energiya and Buran ready to go. With meteorologists predicting excellent weather conditions, the commission set the launch for 29 October at 6:23.46 Moscow time (8:23.46 local time at Baykonur, 3: 23.46 gmt), a decision announced by TASS the same day. Lift-off was timed such that the launch could be observed by the orbiting crew of the Mir space station (Vladimir Titov, Musa Manarov, and Valeriy Polyakov). Mir would be a minute or two short of a Baykonur flyover at lift-off time, allowing the crew to watch virtually the entire launch. Mir’s orbit had been adjusted on 16 October to permit close coordination during the brief Buran flight, including launch and retrofire. However, the observations from Mir were not a strict requirement and the launch was not likely to be scrubbed if that objective could not be met. The main constraint for the launch window was to ensure a landing at Baykonur well before local sunset, ideally around noon. Speaking to reporters later that same day, Doguzhiyev did not hide the tension felt around the cosmodrome:

“No one is indifferent or passive at the cosmodrome… Behind outward calm

there is much nervous pressure. Even we [State] Commission members find it

difficult to answer questions’’ [41].

Among the final tasks to be accomplished at the pad in the last three days prior to launch was the retraction of the 145 m high rotating service structure, a relic of the N-1 days, which was now moved back to its parking position, fully exposing Buran to the elements. Strict safety measures were in place to protect personnel against any potential accidents on the pad. The region around the launch complex was divided into four safety zones. Zone 1 (2 km radius around the pad) was completely evacuated 12 hours before launch. By that time any personnel involved in final countdown operations were required to go to hermetically sealed and heavily armored bunkers, from where all final launch preparations (including fueling) were controlled. The bunkers were said to be capable of surviving impacts of rocket debris. Zone 2 (5 km radius) was cleared of personnel at T — 8 hours as final preparations got underway for loading of liquid hydrogen. Zones 3 and 4 (8.5 km and 15 km radius) were evacuated at T — 4 hours and T — 3 hours to ensure safety of people in case of an explosion during engine ignition and during the early stages of ascent. The rules were much stricter than at the Kennedy Space Center, where people are allowed to watch Shuttle launches in open air from a distance of just about 5 km.

Two days before the planned launch, concern arose over some equipment in Buran’s automatic landing system. VNIIRA, the design bureau in charge of the system, requested installing back-up equipment aboard the orbiter and first test that aboard a Tu-134B aircraft. Although this required activation of all the navigation and landing support systems at the Yubileynyy runway, the landing tests were authorized given the potentially catastrophic consequences of a failure in the auto­matic landing system. The back-up equipment was successfully tested during several approaches to the Yubileynyy runway on 28 October.

On the eve of the launch, Soviet officials backed down from their earlier promises to provide live television coverage of the launch and were now planning to show the recorded launch 35 minutes after the event. The landing would not be carried live either. As expected, the planned launch time of 3: 23 gmt went by without any comment from the Soviet media. Anxiety grew as nothing was heard in the following 45 minutes or so. Finally, shortly after 4:00 gmt, TASS broke the silence by issuing a brief two-line statement:

“As has been reported earlier, the launch of… Energiya with the orbital ship Buran had been planned for 6.23 Moscow time on 29 October. During pre­launch preparations a four-hour delay of the launch has been announced.’’

This now theoretically put the launch time at 7: 23 gmt, raising serious doubts among observers that the launch would take place that day at all. According to the original flight plan, landing would have taken place at 6: 49 gmt, but the launch delay would now move this to around 11: 00 gmt (16: 00 local time at Baykonur), close to sunset. At around 7: 30 gmt, shortly after the rescheduled launch time, TASS reported what had been obvious all along:

“During final launch preparations for the rocket carrier Energiya with the orbital ship Buran there was a deviation in one of the launch support systems. As a result of this an automatic command was issued to stop further work. At the present time work is underway to eliminate the problems. A new launch date and time will be announced later.’’

It wasn’t until later in the day and in press interviews the following days that officials began providing details about the exact cause of the scrub. It turned out the countdown had been halted at T — 51 seconds because a platform had failed to properly retract from the rocket. When it came to describing the exact nature of that platform, Soviet space officials, not used to communicating problems to the media, did a poor job. Talking to a Pravda reporter, Vladimir Gudilin, the head of launch pad operations, said:

“51 seconds before the launch one of the servicing platforms did not move away from the rocket. To be more precise, it visually moved away, but the signal confirming this did not reach the computer checking the launch readiness of all systems. Until the last seconds this platform holds an aiming platform, controlling the gyroscopes. The computer [did not receive the retraction signal] and instantaneously stopped the launch program” [42].

Despite the fact that the Soviet orbiters had long been mothballed by the time ISS construction began, one piece of Buran technology does play a vital role in station

operations. This is the Androgynous Peripheral Docking System (APDS—Russian acronym APAS), a Russian-built docking mechanism that allows Space Shuttles to dock with the US-built Pressurized Mating Adapters (PMAs) on the ISS “node” modules. Built at RKK Energiya under the leadership of Vladimir Syromyatnikov, the first APAS (APAS-75) was developed back in the 1970s for the Apollo-Soyuz Test Project. A modified version (APAS-89) appeared in the 1980s to enable Soviet orbiters to dock with the axial APAS docking port of Mir’s Kristall module. In the end, Buran never flew to Mir and the Kristall APAS docking port was used only once by Soyuz TM-16 in 1993.

In July 1992 NASA initiated the development of the Orbiter Docking System (ODS) to support Shuttle flights to Mir. Mounted in the forward end of the payload bay, the ODS consists of an external airlock, a supporting truss structure, and an APAS docking port. While the first two elements were built by Rockwell, the APAS was manufactured by RKK Energiya. Although Energiya’s internal designator for the Shuttle APAS is APAS-95, it is essentially the same as Buran’s APAS-89. While the ODS was slightly modified for Shuttle missions to ISS, APAS remained unchanged. There was even a suggestion to launch Buran to Mir to test the docking system prior to the beginning of the Shuttle-Mir flights [31].

The APAS consists of a three-petal androgynous capture ring mounted on six interconnected, ball screw shock absorbers that arrest the relative motion of the two vehicles and prevent them from colliding. The APAS-89 differs from APAS-75 in several key respects. It is much more compact (although the inner egress tunnel diameter is more or less the same), has twelve structural latches rather than eight, the guide ring and its extend/retract mechanism are packaged inside rather than outside the egress tunnel, and the three guide petals are pointed inboard rather than outboard [32].

Russian plans to sell their entire Buran Docking Module to NASA fell through, but its design did serve as the basis for the construction of the Russian Pirs airlock module, docked to ISS in September 2001. This retains the central part of the adapter’s spherical section (2.55 m in diameter). Mounted to its aft end is a small section of the Buran airlock’s cylindrical tunnel (without the extendable part) and attached to the front end is the forward part of a Soyuz-TM/Progress-M orbital module. The design was more complex than that of the Docking Module of Mir (316GK), which did not have to be used as an airlock and was merely an extension to the Kristall module to facilitate Shuttle dockings [33].

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.