Category Energiya-Buran

The ultimate dream of the Energiya designers was to develop a rocket that would be fully reusable (Energiya-2 or GK-175). The plan was to achieve full reusability in various steps, first of all by having the strap-on rockets parachute back to Earth for recovery. In the next step the core stage was to be turned into a reusable winged stage with three RD-0120 engines and a payload compartment in the upper section. Despite the lower amount of propellant, the overall dimensions of the core stage remained the same, freeing up some 610m3 of volume in the payload section, which compared favorably with the 350m3 offered by Buran’s cargo bay.

The massive nose fairing would not separate during ascent, but open in space, somewhat like the forward cargo door of a Lockheed C-5 transport aircraft, allowing it to be reused on subsequent flights. After deployment of the payload (30 to 40 tons), the fairing would slide down over the LOX tank so that the core stage would shrink in size from 60 to 44 m to prevent stability problems during re-entry. In order to cut costs, the core stage would inherit as many systems as possible from Buran (wings, vertical stabilizer, landing gear, avionics, and hydraulic systems). It was not con­sidered expedient though to cover the stage with Buran’s heat-resistant tiles and efforts focused instead on using innovative non-ablative and active cooling thermal protection systems.

The next phase was to replace the standard strap-ons by the same type of flyback boosters being envisaged for Energiya-M. The parachute recovery system imposed a

heavy weight penalty on the rocket, the impact zones limited the number of launch azimuths, and recovery from those distant impact zones would have been a laborious and costly undertaking. The flyback strap-ons would be equipped with long foldout wings, a V-shaped tail, and a small jet engine enabling them to fly back to the launch site after separation from the core stage. Although the idea was tempting, the landing of four strap-ons in quick succession on the single Baykonur runway would probably have caused tremendous logistical problems.

In the final step the four flyback strap-ons would be replaced by a first stage equal in size to the second stage and with similar landing systems, but without thermal protection, and equipped with four RD-170 engines. This vehicle would have a payload capacity of between 30 and 50 tons. By using two such first stages it would be possible to increase payload capacity to 200 tons, about the same as the Vulkan with its eight strap-ons. There was even an idea to use four of the large first stages and lengthen the second stage to reach a phenomenal payload capacity of 500 tons. Despite the impressive prospects, any of those three variants would probably have required major modifications to the existing Energiya launch facilities.

Not surprisingly, all these bold proposals faced an uphill battle as the budgets for the space program became ever tighter towards the end of the 1980s. The only spin-off from the Energiya-2 studies is Baykal, a reusable flyback booster now being proposed as a first stage for the Angara rocket family. This incorporates many ideas that had originally been conceived for Energiya-2’s flyback strap-ons [67].

NPO Molniya advertised MAKS-OS by pointing out the following advantages:

– a high degree of reusability (with Mriya partially replacing the traditional rocket first stage + the return of the RD-701 engine aboard the spaceplane);

– its ability to fly from any first-category airfield outfitted with proper ground support and propellant loading facilities;

– an impressive 2,000 km cross-range capability, allowing the vehicle to land on runways located far from the orbital plane;

– an almost unlimited range of launch azimuths + short launch preparation times, combining to make it ideal for quick-response missions such as rescue of space station crews;

– an environmentally clean system thanks to the use of non-toxic propellants and the absence of rocket stage impact zones.

MAKS-OS was primarily seen as a launch system for both government and commercial small and medium-size satellites, the hope being that it would reduce launch costs by as much as ten times compared with expendable launch vehicles. NPO Molniya estimated that the system would break even after just three years if an annual launch rate of 20-25 missions was achieved. Although rarely mentioned, MAKS-OS also inherited the military advantages of the canceled Spiral system and was considered for reconnaissance, inspection, and attack missions [7]. Vladimir

Skorodelov, a deputy chief designer at NPO Molniya, later acknowledged that MAKS was conceived with both civilian and military goals in mind and that the civilian applications came to the foreground only as the Cold War drew to a close [8]. One may even wonder if MAKS wasn’t at least partially inspired by the US Air Force’s Space Sortie system, a quick-response military spaceplane studied in the early 1980s that would be launched with an external fuel tank from the back of a modified Boeing 747.

The preliminary design for MAKS was finished in 1988 and the system was first publicly presented by Gleb Lozino-Lozinskiy at the 40th Congress of the International Astronautical Federation in Malaga, Spain in October 1989. Realizing that MAKS stood little chance as an exclusively government-funded project, NPO Molniya sought international partners to join the project. There was considerable European interest in the early 1990s. British Aerospace saw MAKS as a possible intermediate step towards its own Interim HOTOL, an An-225 launched version of the original British HOTOL single-stage-to-orbit spaceplane. ESA displayed interest in MAKS as an alternative to its own Hermes spaceplane. In 1993-1994 ESA sponsored a joint study by British Aerospace, NPO Molniya, TsAGI, and the Antonov design bureau on a MAKS look-alike Rocket Ascent Demonstrator Mis­sion (RADEM) to prove the technology for a possible European/Russian/Ukrainian reusable air-launched system. However, the results of the study were never imple­mented as ESA lost interest in a European-funded space transportation system. The study did result in NPO Molniya’s later MAKS-D proposal [9].

MAKS received little support within the Russian Space Agency, with Yuriy Koptev having spoken out against it even before becoming the head of the agency

in 1992 [10]. In 1998 Koptev claimed the system would cost $6-7 billion to develop, which was the same amount projected by British Aerospace in the early 1990s and twice the amount estimated by NPO Molniya itself [11]. The main objection raised against MAKS was the high launch rate required to make it cost-effective. The number of domestic satellite launches in the 1990s was quickly dwindling and estimates showed MAKS would be able to launch only 30 percent of the Russian payloads planned until 2010. Many also doubted that the system would ever capture a major share of the international launch market. After all, MAKS was a funda­mentally new launch system and not well suited to launch geostationary satellites, which comprise the bulk of the international commercial payloads.

Questions were also raised about the announced reusability (100 missions for each spaceplane and up to 15 missions for the RD-701 engine). NPO Molniya was also said to underestimate the cost of equipping airfields all over the world with the necessary support infrastructure, such as satellite-processing buildings and propellant storage facilities. Another major concern was that NPO Molniya poorly addressed safety issues related to MAKS’ use of cryogenic propellants and its all-azimuth launch capability. The latter in many cases required the vehicle to fly over populated, not to mention foreign territory [12].

Despite all the objections, NPO Molniya continued low-level research on MAKS using shoestring government funds (at least partially thanks to continued support from the military) and other financial means, some provided by the Moscow city government. Full-scale mock-ups were built of the OS spaceplane and the external fuel tank. A crude experimental version of the RD-701 began testing at NPO Energomash in 1994. Mriya re-entered service in 2001 after having been grounded for seven years.

MAKS was given a new chance in late 2005, when it competed with proposals by RKK Energiya and the Khrunichev Center in a tender to develop a successor for the Soyuz spacecraft. The spaceplane proper now had a slightly differently shaped fuselage and no longer had foldable wings. However, the Russian Space Agency canceled the tender in July 2006, preferring to develop a capsule-type vehicle in collaboration with ESA. One of the main drawbacks cited for MAKS was the considerable Ukrainian involvement—namely, the Antonov bureau’s Mriya aircraft. One of the requirements in the tender had been to limit foreign contributions. MAKS is now destined to go down in history as yet another unrealized Russian spaceplane project.

Spaceplanes were also studied in the early 1960s at the OKB-156 bureau of the Soviet Union’s most famous aircraft designer Andrey N. Tupolev. These studies had their roots in research conducted in 1957-1960 on an unmanned Long-Distance Glider (DP) intended to deliver thermonuclear warheads to enemy territory. According to original plans the DP was to be launched to an altitude between 50-100 km by a missile, either the R-5 or R-12, or a booster built at the Tupolev bureau itself. After separation from the rocket, it would gradually glide to its target, located up to

4,0 km from the launch pad. An on-board altimeter would then detonate the thermonuclear bomb at the required altitude.

Scale models of the DP were launched to speeds of up to Mach 2 with small solid rocket motors from Tu-16LL aircraft. OKB-156 also developed an experimental prototype of the DP called 130 or Tu-130. Weighing 2.5 tons, the tailless glider was 8.8 m long and 2.2 m high with a wingspan of 2.8 m. However, on 5 February 1960, just as the first Tu-130 was being readied for launch on a modified R-12 missile, the Soviet government issued a decree to cancel the DP project, now considered useless in the wake of the early ICBM successes. By this time OKB-156 had been aiming to launch the DP with a three-stage rocket built in-house, enabling the glider to cover distances of 9,000 to 12,000 km and carry a thermonuclear warhead weighing 3 to 5 tons.

The experience gained during the research on the DP came in handy for Tupolev’s spaceplane project, presumably started around 1960 under the names Aircraft 136, Tu-136, or Zvezda (“Star”) (Tu-136 is also the name of a recently developed regional cargo/passenger plane). The ultimate goal was to build a 10- to

20-ton spaceplane to be orbited by a newly developed launch vehicle. Several aero­dynamic shapes were studied, one closely resembling that of the 130 glider and another that of Dyna-Soar. In the end the designers opted for a canard configuration. If the experimental version was successful, it would serve as the basis for a whole series of rocket planes to be used for reconnaissance, bombing, and anti-satellite missions. Tupolev envisaged a grueling two-phase test program to verify the design at hypersonic speeds in the lower and upper atmosphere and to try out re-entry and landing techniques.

The first phase would see scale models of Zvezda being launched from Tu-16 aircraft and with the help of R-5 and R-14 missiles. The air-launched version would have a built-in solid rocket motor to reach an altitude of up to 40 km and a speed of

9.0 km/h. Models launched by the R-5 and R-14 would climb to 45 km and 90 km, respectively, and develop speeds of 14,000 km/h and 23,000-28,000 km/h.

The second phase involved the use of three manned test vehicles. One was a scaled-down version of Zvezda known as 136-1, air-launched from a Tu-95K. Having reached a peak altitude of 10 km and a top speed of 1,000 km/h, it would land at a speed of about 300 km/h, just like the real Zvezda. The next step was to use the Tu-95K as a launch platform for a hypersonic vehicle designated “139”. The Soviet equivalent of the X-15, it was to fly as high as 200 km and develop a speed of

8.0 km/h. The third vehicle was dubbed 136-2, an improved version of the 136-1 with an additional rocket engine to reach speeds of up to 12,000 km/h and a max­imum altitude of about 100 km.

After this the stage would be set for the first launches of the actual Zvezda vehicle, which would fly between altitudes of 50 and 100 km and therefore be limited to single-orbit missions. There were also plans for an unmanned version called 137, Tu-137, or Sputnik, capable of performing multi-orbit missions. The only launch vehicle capable of launching Zvezda was Chelomey’s UR-500/Proton, but this was only in the very early stages of development when Zvezda was conceived. Therefore OKB-156 worked out plans for its own two – or three-stage rocket to launch the spaceplane. Also considered was a scheme in which the spaceplane would be launched with a missile from the back of a strategic supersonic plane (the Tu-135 or Tu-139).

Work on Zvezda was discontinued in 1963 for reasons that have not been dis­closed [21].

The basic configuration of the RLA rockets was a common core stage complemented by a different number of standard first-stage strap-on boosters, depending on the mass of the payload. Very little has been revealed about the RLA launch vehicles studied in 1974-1975 and various sources have also given different designators for rockets with similar capabilities. Apparently, the design evolved significantly even during that short period, dictated by the progress made in the concurrent research on kerosene/hydrocarbon and hydrogen engines. It would appear that original plans for large clusters of low-thrust engines eventually gave way to small clusters of high-thrust engines as the confidence in the latter grew.

Glushko is known to have presented plans for three RLA rockets (RLA-120,135, and 150) during a meeting on 13 August 1974, which was attended by most of the chief designers and also by Dmitriy Ustinov. This was Glushko’s third RLA proposal in the barely three months he had been in office at NPO Energiya. The August 1974 RLA plans revolved around the exclusive use of kerosene and sintin, the 1,003-ton thrust RD-150 engine, and massive 6m diameter rocket modules. The RLA-120, expected to be ready in 1979, would have a payload capacity of about 30 tons and among other things launch modules of a permanent space station. The RLA-135 had a 100-ton payload capacity and could be used to orbit a reusable space shuttle or

elements of a lunar base and was expected to make its debut in 1980. Finally, there was the massive RLA-150, capable of placing up to 250 tons into low orbit and seen by Glushko as the rocket that would eventually send Soviet cosmonauts to Mars. Its first flight was anticipated in 1982 [48].

Other sources have identified three rockets known as RLA-110 or Groza (“Thunderstorm”), RLA-120 or Grom (“Thunder”), and RLA-130 or Vulkan (“Vol­cano”). The RLA-110, equipped with two boosters, would have a payload capacity “higher than the Proton rocket”. The RLA-120, using four boosters, would have about the same payload capacity as the N-1. Finally, the RLA-130, toting eight boosters, would play a key role in establishing a Soviet lunar base [49].

While Glushko’s RLA plan may have looked very appealing on paper, upon closer analysis it did raise the necessary questions among fellow designers. Some warned that because of the unification not all the launch vehicles in the series would be the most efficient in their particular payload class. The design of the common core stage had to be tailored to the heaviest 250-ton class booster, which was the one expected to fly least. The implication was that the core stage was oversize for the 30-ton class RLA, exactly the one that would probably be launched most frequently [50]. This is probably the very reason preference was eventually given to Yuzhnoye’s 11K37 “heavy Zenit’’ to fill this niche in the payload spectrum. Some felt that a better way of developing a standardized rocket fleet was to first fly a light booster, then use its first stage as the second stage for a heavier booster, subsequently turn that second stage into a third stage for an even heavier rocket, etc. [51]. This was the approach that Korolyov and Chelomey had suggested for their respective N-I/N-II/N-III and UR-200/UR-500/UR-700 families in the 1960s. One big disadvantage, however, was that each vehicle in the fleet would require its own launch pad.

In the end, the only rocket that emerged from the RLA plans was the 100-ton class booster that later became known as Energiya. Proposals were later tabled for Energiya derivatives such as Energiya-M (30-ton class), Groza (60-ton class), and Vulkan (200-ton class), but these never made it off the ground (see Chapter 8). And so the dream of a standardized rocket fleet in the heavy to super-heavy class never materialized either, although the main reason here was the absence of payloads to justify its existence.

The RD-170 (also known as 11D521), designed and manufactured by KB Energo – mash in Khimki near Moscow, was a LOX/kerosene engine employing the staged combustion cycle. Providing 740 tons of thrust and a specific impulse of 308.5 s at ground level, it remains not only the most powerful LOX/kerosene engine built to date, but also the highest-thrust liquid-fuel engine flown on any launch vehicle in the world.

Although Energomash had gained significant experience with staged-combustion cycle engines burning hypergolic propellants, the RD-170 marked the bureau’s first foray into closed-cycle LOX/kerosene engines. The only other closed-cycle LOX/ kerosene engines built in the Soviet Union until then had been much less powerful single-chamber engines such as the ones used on the Blok-L and Blok-D upper stages (built by the OKB-1 Korolyov bureau) and the NK engines for the first three stages of the N-1 rocket (developed by the Kuznetsov bureau in Kuybyshev). The United

States has never built a staged-combustion cycle LOX/kerosene engine. The only powerful LOX/kerosene engine ever flown by the United States was the F-1, five of which powered the first stage of the Saturn V. This was an open-cycle engine inferior in most aspects to the RD-170.

The RD-170 consisted of four combustion chambers, one turbopump assembly, and two gas generators. The turbopump assembly incorporated a single-stage active axial-flow turbine, an oxidizer pump, and a two-stage fuel pump. Connected to the assembly were low-pressure oxidizer and fuel pumps to increase the pressure of the propellant and thereby prevent cavitation of the turbopump assembly. The turbo­pump was driven by two oxidizer-rich gas generators. Originally, it was planned to have a single gas generator consuming 1.5 tons of propellant per second, but this would have been too big. In the RD-170 the entire oxidizer supply and just a small fraction of the kerosene (6% of the overall propellant mass) passed through the gas generators. The turbopump produced about 257,000 horsepower, which the Russians like to compare with the combined horsepower of three of their heavy nuclear icebreakers.

The RD-170 could be throttled down to 50 percent of rated thrust and could be gimbaled about 8° with the help of hydraulic actuators. The engine could be gimbaled in two axes, whereas the Zenit’s RD-171 had only single-axis gimbal capability. Therefore, each RD-170 required a total of eight hydraulic actuators, two for each combustion chamber. Unlike the RD-171 nozzles, those of the RD-170 entered the air stream impinging on the rocket when they were swiveled, requiring the use of more powerful actuators to counter the aerodynamic pressures.

With a nominal flight burn time of 140-150 seconds, the engine was designed to be used at least ten times, a capability confirmed during bench tests. Although the RD-170 was used only for the two Energiya missions in 1987 and 1988, its nearly identical twin (the RD-171) continues to fly today on the two-stage Zenit launch vehicle and its three-stage Sea Launch version. A derived version with just two combustion chambers (the RD-180) now powers America’s Atlas-5 rockets and a single-chamber version (the RD-191) is expected to become the power plant of Russia’s Angara family of launch vehicles (see Chapter 8). [4]

Although Buran lacked main engines for ascent, it did have engines and thrusters for on-orbit maneuvers and attitude control functions. Buran’s propulsion system was known as the Combined Engine Installation (ODU or 17D11) and consisted of an integrated set of orbital maneuvering engines, primary thrusters, vernier thrusters, and associated plumbing.

While the overall number and general location of these engines were similar to those of the Space Shuttle Orbiter’s Orbital Maneuvering System (OMS) and Reac­tion Control System (RCS), there were some fundamental differences between the two vehicles, notably the types of propellant used. Orbital maneuvering and attitude control engines on manned spacececraft have traditionally used hypergolic propel­lants or hydrogen peroxide, which can be stored for long periods of time and do not require complex ignition and turbopump systems. The Space Shuttle Orbiter uses a hypergolic mix of nitrogen tetroxide and dimethyl hydrazine for both its OMS and RCS engines. Although Soviet designers also planned to use hypergolic propellants in their original orbiter concepts (OS-120 and OK-92), they eventually opted for a combination of liquid oxygen and a synthetic hydrocarbon fuel known as sintin. This marked the first time that such propellants were used in any type of orbital maneuvering and attitude control system. Next to the absence of main engines, this was probably the most significant difference between Buran and the Space Shuttle Orbiter.

Cryogenic propellants offered a number of advantages. They gave the orbital maneuvering engines a better performance than those of the Shuttle (although

thruster performance was virtually identical) and were safer to handle by ground personnel because of their non-toxicity. Moreover, the LOX could be cross-fed to the storage tanks of Buran’s electricity-generating fuel cells, providing extra redundancy to the power system and, indirectly, to the life support system, which drew oxygen and water from the fuel cell system. The drawbacks were that the plumbing was more complex, making the ODU 1,100 kg heavier than the Shuttle’s RCS/OMS system. Also, the mix did not ignite spontaneously on contact, such as was the case with hypergolic propellants, but required an electric ignition source. In addition to that, extra measures needed to be taken to prevent the cryogenic oxidizer from boiling off during long missions.

It is interesting to note that in the 1990s US Shuttle engineers considered a cryogenic OMS/RCS as a long-term Shuttle upgrade. This would have used a combination of LOX and ethanol and would have enabled the forward RCS, aft RCS, and OMS engines to draw propellant from common tanks, just as on Buran. It is not clear if this upgrade was in any way inspired by the design of Buran’s ODU.

Like the Space Shuttle Orbiter, Buran was a versatile vehicle that could have been used for a wide range of orbital operations. The following possible tasks were later identified by the Russians:

(a) Deployment of satellites or other cargos: the maximum payload was 30 tons into a 50.7° inclination 200 km orbit and 16 tons into a 97° orbit. The payload bay could house a payload with a maximum length of 15 m and a maximum diameter of 4.15 m. Because of the less stringent center-of-gravity requirements resulting from the absence of main engines, Buran’s maximum payload capacity was actually higher than that of the Space Shuttle.

(b) Servicing satellites in orbit

(c) Returning satellites back to Earth. The maximum mass that could be returned from a 50.7° inclination 200 km orbit was 20 tons.

(d) Space station missions: resupply, assembly, crew exchange, crew rescue.

(e) Missions to assemble large structures in space.

(f) Autonomous scientific missions.

Three basic types of operational mission durations were envisaged for the vehicle. The first would be short-duration missions (up to 3 days) to place heavy payloads into orbit, deliver emergency supplies to space stations, or rescue space station crews. Such missions would be characterized by multiple operations and maneuvers in a relatively short time span, heavily taxing both the crew and the ground and also requiring many of them to be conducted automatically.

Medium-duration missions (up to 8 days) were expected to be the most frequent ones and would have several objectives or one particularly time-consuming and demanding goal. Typical medium-duration flights would include routine missions to space stations, multiple satellite deployment missions, satellite-servicing missions, assembly flights, etc. Although comparable in the number of operations with the short-duration flights, the longer time in orbit would make it possible to more evenly spread the workload for the crew.

Finally, long-duration missions (9 to 30 days) would primarily be devoted to scientific, materials-processing, and biotechnological experiments, which take a relatively long time to produce the necessary results. For this purpose, the Russians were planning to develop a Spacelab-type module to be placed in the cargo bay. The longest missions would have required the installation of an extra cryo kit for the fuel cells. The number of maneuvers performed during this type of mission would have been very low. In terms of the daily workload for the crew and the ground, such missions would have been comparable with a routine workday on a space station.

Range safety restrictions at the Baykonur cosmodrome, mainly dictated by the impact zones of the strap-on boosters, limited the possible orbital inclinations of the spacecraft to 50.7-83°, 97°, 101-104°, and 110°. The vehicle could have operated at altitudes between 200 and 1,000 km, although the higher of these would have necessitated the installation of extra propellant tanks in the cargo bay [29].

The modular part of the Energiya strap-on boosters was designed by KB Yuzhnoye in the Ukrainian city of Dnepropetrovsk. This originated in 1954 as OKB-586 and under the leadership of Mikhail K. Yangel pursued the development of ballistic missiles using storable propellants as well as derived launch vehicles such as the Kosmos and Tsiklon series of boosters. Renamed KB Yuzhnoye in 1966, it was also in charge of developing a wide array of military and scientific satellites. Yangel died in 1971 and was succeeded by Vladimir F. Utkin, who headed the organization until 1990.

Yuzhnoye’s production facility was the Yuzhnyy Machine Building Factory (YuMZ or “Yuzhmash”), originally founded in 1944 as the Dnepropetrovsk Automobile Factory. In 1951 it was renamed Factory nr. 586 and ordered to switch to the serial production of OKB-1 missiles (the R-1 and R-2). Several years later it began producing missiles and eventually launch vehicles and satellites for KB Yuzhnoye. During the Buran years Yuzhmash was headed by Aleksandr M. Makarov (1961-1986) and Leonid D. Kuchma (1986-1992), the later President of the Ukraine. Eventually, serial production of the modular part of the Energiya strap – ons was also to be transferred to PO Polyot in Omsk, but this apparently never happened.

The Cosmonaut Training Center was the first to select a dedicated group for the Buran program. On 23 August 1976, just six months after the official approval of the Energiya-Buran program, nine pilots were chosen, the first selection by TsPK in six years. They were:

• Leonid Georgyevich Ivanov;

• Leonid Konstantinovich Kadenyuk;

• Nikolay Tikhonovich Moskalenko;

• Sergey Filippovich Protchenko;

• Yevgeniy Vladimirovich Saley;

• Anatoliy Yakovlevich Solovyov;

• Vladimir Georgyevich Titov;

• Vladimir Vladimirovich Vasyutin;

• Aleksandr Aleksandrovich Volkov;

All of them were young, relatively inexperienced Air Force pilots in their mid to late twenties. The rationale behind their selection at this early stage may have been that they would need several years to advance their flying skills while the more experienced LII and GKNII test pilots conducted the early Buran test flight program.

Not surprisingly, it was decided that the cosmonauts would first have to undergo test pilot training before beginning the standard cosmonaut training course. They began studying and training at TsPLI in Akhtubinsk in September 1976, becoming Test Pilots 3rd Class (the lowest test pilot rank) in June 1977. In addition, in August they conducted parachute training. From October 1977 until September 1978 they then underwent the standard basic cosmonaut training course (“General Space Training” or OKP) at TsPK.

After graduation most members of the group (Ivanov, Kadenyuk, Moskalenko, Protchenko, Saley, Solovyov, and Volkov) returned to Akhtubinsk to resume test pilot training with the goal of becoming Test Pilots 2nd Class. It was during this follow-up course that Sergey Protchenko was medically disqualified and dismissed from the cosmonaut team in April 1979 [3]. More than a year later, on 24 October 1980, the group suffered another loss when Leonid Ivanov was killed in the crash of a MiG-27 in Akhtubinsk.

On 22 June 1981, Kadenyuk, Moskalenko, Saley, Volkov, and Solovyov were awarded the title Test Pilot 2nd Class. After that, the first four went on to conduct Buran-related training, but Solovyov was transferred to the Salyut space station program, together with Titov and Vasyutin.

Leonid Kadenyuk was the next to be dismissed. He left the cosmonaut team in March 1983 after he had divorced his wife. In the Soviet Union of the 1970s and 1980s, getting a divorce usually resulted in the end of a cosmonaut career for those who were still awaiting their first mission.

The 1976 selection had been limited to the Air Force and it was therefore decided that another screening would take place in the Soviet Navy and Air Defense Forces

[4]. As a result, on 23 May 1978 one additional candidate from each of these two branches of the military was added to the detachment:

• Nikolay Sergeyevich Grekov (Air Defense Forces);

• Aleksandr Stepanovich Viktorenko (Navy).

In October 1978 the two began training in Akhtubinsk, graduating as Test Pilot 3rd Class on 2 July 1979. They then returned to Star City, where they underwent OKP, finishing that in February 1982. Viktorenko almost died in a bizarre accident during a medical check-up in 1979. He was wearing a band with electrodes around his body to have an ECG made when a 220-volt current was accidentally sent through it. Apparently his heart stopped and he was brought back to life using CPR. According to Viktorenko the incident cost him quite some time in training as the doctors wanted to be 100% certain that he had not suffered any ill effects from the incident [5].

As the training went on, it was becoming increasingly apparent that Buran’s first flights would be significantly delayed. Since TsPK was becoming confronted with a shortage of commanders for Soyuz and Salyut, it was decided in late 1983 to transfer all members of the 1976 and 1978 selections to the space station program. In the following years, they would become the core of the cosmonaut detachment, with several of them flying record-breaking missions (for details on further careers see the cosmonaut biographies in Appendix B).

Like the RD-0120, Energomash’s RD-170 engines for the strap-on boosters under­went a step-by-step test program, moving from autonomous tests of individual components to full-scale test firings at increasing rates of thrust. Bench tests were not only carried out with the engines alone (at Energomash’s own test facility), but also with the engines mounted on the modular part of the strap-ons (at

Nllkhimmash). Test firings of complete strap-ons at Baykonur’s UKSS pad were ultimately canceled, at least partially because the RD-171, the near twin of the RD-170, had already undergone flight testing on the first stage of the Zenit rocket. The road to success for the RD-170 proved much more arduous than for the RD-0120, with several setbacks plunging the program into deep crisis in the early 1980s.