Category Energiya-Buran

For regions exposed to temperatures of up to 370° Buran had multiple-layer, square­shaped panels of flexible insulation, similar to the Felt Reusable Surface Insulation (FRSI) employed by the Shuttle. Known as ATM-19PKP, the material was similar to that used for the felt pads under the tiles and was applied to the upper payload bay doors, portions of the upper wing surfaces, and portions of the mid fuselage.

Carbon-carbon

The areas where Buran incurred the highest heating during re-entry (up to 1,650°C) were the nosecap and the leading edges of the wings. As on the Orbiter, these parts were covered with a reinforced carbon-carbon (RCC) material. Until 1978 efforts focused on an RCC material known as KUPVM-BS, but despite its high thermal resistance and strength, it turned out to be too difficult to use. Eventually, the choice fell on a material called GRAVIMOL, an acronym reflecting the names of the three organizations that developed it (NII Grafit, VIAM, and NPO Molniya). There were some small differences in the composition of the RCC material used in the nosecap and the wing leading edges (GRAVIMOL-B in the wing leading edges). The material’s density was 1.85 g/cm3. The RCC had a coating of molybdenum disilicide to prevent oxidation. As on the Shuttle Orbiter, each wing leading edge was covered with 22 RCC panels.

Thermal barriers

Flexible thermal seals protected the vehicle in between certain types of thermal protection material and also in areas containing movable parts. Brush-type seals covered small gaps between sections of the payload bay doors and also in the vertical stabilizer, body flap, and elevons. Seals made of quartz fibers protected areas between the thermal protection system and various doors and hatches. Seals composed of silicon carbide fibers were used in areas exposed to extremely high temperatures such as the gaps between the RCC panels on the wing leading edges and areas where the RCC material bordered on the tiles. An ablative material capable of withstanding temperatures up to +1,800°C covered the gaps between the elevons [14].

A rescue option unique to the Soviet space program was the ability to send a Soyuz spacecraft to an incapacitated orbiter. That plan could have been set in motion in any scenario where Buran would have been unable to return to Earth, such as a propulsion system failure, major damage to the thermal protection system, etc. In any given situation, it would have taken the Soyuz at least several days to reach Buran, making it necessary for the crew to conserve power and consumables until the rescue craft arrived. After crew evacuation, Buran could then either have been sent on a destructive re-entry over unpopulated regions or—if deemed feasible—safely brought back to Earth unmanned.

Of course, a Soyuz rescue could only have been conducted in certain well-defined circumstances. First, it assumed that Buran was equipped with an APAS docking adapter. Second, the ship needed to have at least some level of control (navigation systems and steering thrusters) enabling it to be positioned for the active Soyuz vehicle to dock with it. Third, the crew should have numbered no more than two cosmonauts, since the three-man Soyuz had to be launched with a “rescue com­mander” to assist the stranded pilots in boarding the Soyuz. In the late 1980s/early 1990s the Russians had a cadre of “rescue commanders’’ for emergency flights to Mir who could quite easily have been cross-trained for Buran rescue missions. The Soyuz rescue scenario seems to have been worked out specifically for the early two-man test flights.

Even if Buran carried more than two crew members, a Soyuz rescue was not entirely out of the question, at least if the ship was on a space station mission. Fuel reserves permitting, the Soyuz could have evacuated all crew members by making repeated flights between the stricken Buran and the space station. This was only in the very unlikely event that the vehicle had a problem preventing it from landing and could not reach the station or return to it. Otherwise, the Buran crew could simply

have stayed aboard the space station until rescue arrived. Taking into account the fact that the bulk of Buran missions would have been to Mir and Mir-2, this is a luxury that most Buran cosmonauts would have had long before NASA even began thinking about the “safe haven” concept in the wake of the 2003 Columbia accident.

The Russians took the Soyuz rescue option very seriously. They were even planning to simulate it during the second mission, in which the ship would have launched and landed unmanned but would have been temporarily boarded by a Soyuz crew while in orbit (see Chapter 5). If Buran had ever flown its two-man test flights, a Soyuz vehicle would very probably have been on stand-by at the Baykonur cosmodrome to come to the rescue. The early Buran pilots would have needed some limited Soyuz training, even if the Soyuz would be piloted by a rescue commander. This is probably one of the reasons Buran pilots Igor Volk and Anatoliy Levchenko made Soyuz flights in 1984 and 1987, although the primary goal of these flights was to test their ability to fly aircraft after a week in zero gravity (see Chapter 5).

Of course, it should be understood that, while all these abort scenarios were theoretically possible, it is far from certain that all of the situations described above would have been survivable. Much would have depended on the exact circumstances. Also, at least several of them were only feasible with a limited number of crew members on board (two to four). On the whole, though, it can be said that Buran crews would have stood a better chance of surviving in-flight emergencies than any Space Shuttle crew to date [32].

Internal documents obtained by the authors show that there was fierce debate between LII/MAP and NPO Energiya/MOM in the 1980s over crewing for the first manned missions. It was all very reminiscent of similar disagreements between the Korolyov design bureau and the Air Force over crewing for Voskhod and Soyuz missions in the 1960s. The documents show that the Council of Chief Designers decided on 26 January 1983 to assign only LII test pilots to the first two manned Buran missions, but that NPO Energiya disagreed with the plan in September 1983, putting forward its own flight engineers to occupy the second seat. By the autumn of 1985 NPO Energiya had mustered enough support to secure a joint decision from MOM, MAP, and the Ministry of Defense on the formation of four preliminary crews for the first two manned flights:

Somewhat later the pairings were changed as follows:

Volk Levchenko Stankyavichus Shchukin

Ivanchenkov Strekalov Balandin Krikalyov

Another source claims the crews initially were Volk-Ivanchenkov, Levchenko – Strekalov, Stankyavichus-Balandin, and Shchukin-Lebedev, with Lebedev being replaced by Krikalyov in 1986 [45].

On 6 December 1985 the Military Industrial Commission (VPK) went along with the plan, ordering formation of final crews by December 1986 based on the training results obtained by then. The first phase of training for the NPO Energiya engineers would see theoretical, simulator, and aircraft training. LII demanded that the flight engineers fly a total of 398 hours on five different aircraft, but in the end five engineers (Ivanchenkov, Strekalov, Balandin, Krikalyov, and Lebedev) accumulated just 11 hours of flying time during 26 flights on four aircraft in November 1986. In June 1987 Volk and Ivanchenkov flew 10 different landing profiles on the PDST simulator at NPO Molniya, which according to Volk’s official protocol showed that the engineers would not be able to safely land Buran in case of an emergency.

Based on the preliminary results of the training program, both MAP and the Air Force recommended in 1987 only to fly experienced LII test pilots on the first Buran missions. With their limited aircraft training, the engineers were not even considered capable of flying in the co-pilot seat of the Tu-154LL training aircraft or the BTS-002. Although the prime landing mode even for manned missions was automatic, MAP and the Air Force argued that the crew would have to take manual control if they were diverted to an emergency landing site not equipped with the necessary navigation equipment to support hands-off landings. Moreover, it was felt that the second crew member needed flying skills equal to those of the commander in order to deal with various off-nominal scenarios. Among those were malfunctions in the commander’s flight displays and control panels and a situation where the commander was partially disabled by space motion sickness. A joint LII/TsPK research program called “Dilemma” had shown that the engineers would not be able to render the necessary assistance to the commander in case of these and other emergencies.

Predictably, NPO Energiya and MOM, citing the December 1985 VPK decision, ignored the conclusions of MAP and the Air Force and insisted on a continued training program for the engineers, including simulated flights on the PDST simu­lator and real flights on the Tu-154LL and BTS-002. One of the arguments in favor of including a flight engineer on the first manned flight (then scheduled to be mission 1K2) was that it would be a conservative 3-day flight, with most systems operating in the automatic mode. On the other hand, LII used the same argument to claim that the limited engineering tasks planned for the flight might just as well be performed by a test pilot. At any rate, the Council of Chief Designers ordered on 23 March 1988 to draw up a new training schedule for the NPO Energiya engineers, but it looks as if NPO Energiya pursued its plans with less vigor as the months went on. The launch date for the first manned mission kept slipping and the exact flight plan remained vague, complicating the formation of a training program. Moreover, by the end of the 1980s virtually all of the NPO Energiya flight engineers involved in Buran had

either been reassigned to the Mir program or left, with only Ivanchenkov remaining until 1992 [46].

In later interviews the LII pilots did not hide their opposition to Energiya’s push to include engineers in the first crews. Volk said that at one point he went to Minister of General Machine Building Oleg Baklanov, asking him what the use of flying engineers was. According to Volk, Baklanov quoted Glushko as saying that “they would keep an eye on the devices.’’ Losing his temper in a subsequent argument with Glushko over the crew assignments, Volk told the chief designer: “Then let Strekalov and Ivanchenkov fly! And if there is a crash or whatever, then of course the news will be all over the world’’ [47].

The US Defense Department estimated in the early 1980s that the HLLV would fly first in 1986-1987, followed by the Soviet orbiter in 1987-1988. Ironically, this prediction was more realistic than what was being planned by the Russians, who had a history of setting optimistic timelines for their space projects. When the Energiya-Buran program was approved in February 1976, the goal had been to fly the maiden mission in 1983, but this date started slipping soon. A government decree in December 1981 moved the target date to 1985 and another one on 2 August 1985 set the mission for the fourth quarter of 1986 [26]. Even that must have been a completely unrealistic goal given the progress made by that time in rocket and orbiter testing. A major factor in the delays probably were the serious problems with test firings of the RD-170/171 engines in the early 1980s, although other technical as well as budgetary issues must also have come into play.

As mentioned in the previous chapter, original plans apparently called for launching the first two missions of Energiya with mock-up orbiters that would remain attached to the rocket and re-enter together with it. Later those plans were dropped in favor of launching a real orbiter on the first Energiya mission. Then, as Buran ran into delays, it was decided to turn a test model of Energiya (6S) into a flightworthy version (6SL) and launch that with the Polyus/Skif-DM payload, moving the Buran mission to the second flight of Energiya.

Since the core stage was suborbital, another element that needed to be developed for Buran-T besides the GTK were the upper stages to place payloads into orbit. One of these was a modification of the Proton rocket’s Blok-DM upper stage. Having a diameter of 3.7 m and a length of 5.56 m, it was to carry between 11 and 15 tons of LOX/kerosene. Its engine was to have a thrust of up to 8.5 tons and have the capability of being ignited up to seven times. It could also act as a retro- and correction stage for long-duration deep-space missions, in which case it would need a special propellant-cooling system.

The other upper stage, known as 14S40 or Smerch (“Tornado”), was to use liquid oxygen and hydrogen. It was only one in a family of cryogenic upper stages that the KB Salyut design bureau (part of NPO Energiya in the 1980s) had been tasked to develop by a government decree in December 1984. The others were Shtorm (“Gale”) for the Proton rocket, Vikhr (“Whirlwind”) for Groza (an Energiya with two strap – ons), and the 11K37 (a “heavy Zenit’’) and Vezuviy (“Vesuvius”) for Vulkan (an Energiya with eight strap-ons). Manufacturing was to take place at the Krasnoyarsk Machine Building Factory.

By late 1985 KB Salyut came up with a plan for using the cryogenic 11D56M engine, an improved version of the 11D56 engine developed back in the 1960s by KB Khimmash for the N-1 rocket. With its thrust of 7.1 tons and specific impulse of 461 s, it was well suited for KB Salyut’s own Proton, but did not meet the requirements that NPO Energiya had laid down for Smerch. In July 1988 Minister of General Machine

Building Vitaliy Doguzhiyev directed NPO Energiya and its Volga Branch to propose its own upper stages for Buran-T and Vulkan. NPO Energiya set its sights on the RO-95, an open-cycle LOX/LH2 engine under development at KBKhA in Voronezh.

With a thrust of 10 tons and a specific impulse of 475 s, the RO-95 outperformed the 11D56M by a considerable margin and was also optimized for use in Vulkan’s Vezuviy upper stage. Unlike the upper stage that KB Salyut had proposed, NPO Energiya’s Smerch had the LOX tank on top, which was more favorable in terms of center-of-gravity requirements and also made it easier to ignite the engine in zero gravity. In this configuration Smerch was 5.5 m wide and 16 m long with a propellant mass of up to 70 tons. The engine could be re-ignited up to ten times. Technical requirements for the RO-95 were sent to KBKhA in December 1988 and test firings of the engine were expected to begin in 1991-1992. Yet in February 1989 Doguzhiyev seems to have turned around his earlier decision by limiting work on cryogenic upper – stage engines to KB Khimmash’s 11D56M, arguing that there were no payloads in the pipeline for Buran-T and Vulkan that justified the development of an entirely new engine.

Initially, three upper-stage configurations were studied for Buran-T: only the Blok-DM derived stage for low-orbiting payloads (up to 1,000 km), only the Smerch for payloads destined for geostationary orbit, lunar libration points, and lunar orbit, and the two stages combined for lunar-landing missions, flights to Mars and Jupiter. Payload capacity would have been about 88 tons to low Earth orbit, 18-19 tons to geostationary orbit, 21.5-23 tons into lunar orbit, 9-10 tons to the lunar surface, and 10-13 tons into Martian orbit [57].

In 1993 the Russian Space Agency initiated a research and development program called Oryol (“Eagle”) to devise a strategy for the development of reusable space transportation systems in the 21st century. While the program was mainly aimed at technology development, several design bureaus were also invited to work out poss­ible schemes for a Russian Aerospace Plane (RAKS), although it is hardly likely the intention was to actually build one. The focus was both on SSTO and two-stage-to – orbit (TSTO) concepts.

Schemes for vertically launched, partially reusable TSTO systems were devised by RKK Energiya, KB Salyut (which became part of the Khrunichev Center in 1993), and TsNIImash. All these revolved around the use of winged flyback boosters and expendable second stages, capable of placing about 25 tons into low 51° inclination orbits. All the concepts relied on the use of LOX/LH2 engines, with the RD-0120 figuring prominently in three of the four schemes. The payloads could either be traditional satellites placed under a payload fairing or spaceplanes. Primarily intended for space station support, these Reusable Orbital Ships (MOK) would have an expendable instrument and cargo compartment.

Attention was also given to air-launched systems. It would seem that NPO Molniya got some funding under Oryol to continue work on its air-launched MAKS versions. Meanwhile, the Mikoyan bureau studied a fully reusable TSTO system called MiGAKS, consisting of a turbojet/ramjet powered hypersonic carrier aircraft and a spaceplane with rocket engines. The aircraft would propel the spaceplane to Mach 6 before releasing it and would then return either to its home base or to a runway downrange. The Mikoyan bureau studied hypersonic planes burning a combination of kerosene and hydrogen (total take-off mass 420 tons) or hydrogen alone (take-off mass 350 tons). Payload capacity to a low 51° inclination orbit was 12.3 tons for the first version and 10 tons for the second version.

In the SSTO area, the Mikoyan bureau came up with an unmanned spaceplane called MiG-2000. Weighing 300 tons at take-off, the 54 m long vehicle would be accelerated to Mach 0.8 by a liquid-fueled rocket sled, with ramjets propelling it to Mach 5 before rocket engines burning LOX and subcooled liquid hydrogen took

over to boost it to orbit. Payload capacity was 9 tons to a low 51° inclination orbit and cross-range capability was up to 3,000 km.

RKK Energiya proposed a 1,400-ton SSTO spaceplane called MKR (Reusable Space Rocket Plane). This would be launched on its own vertically, powered by seven tripropellant LOX/LH2/kerosene engines with a sea-level thrust of 250 tons each. Externally resembling a Buran orbiter, most of the mid and aft fuselage was occupied by propellant tanks, leaving room only for a 8.0 x 4.5 m payload bay. Payload capacity was anywhere from 10 to 18 tons to low 51° orbits, depending on whether the vehicle was manned (maximum crew of three) or unmanned. Missions would last no longer than seven days. Cross-range capability was 2,000 km [22].

There was other SSTO research in the 1990s apparently not funded under Oryol. Khrunichev’s KB Salyut worked on a vertical take-off/horizontal landing system reminiscent of America’s VentureStar, and the Makeyev bureau designed a vertical take-off/vertical landing system called Korona similar to the American DC-X and its Delta Clipper prototype [23]. Finally, NPO Molniya did paper studies of sled – launched SSTOs (VKS-R) as well as vertical take-off/horizontal landing systems (VKS-O) [24].

Perhaps the most exotic SSTO concept was Ajax, originally conceived in the late 1980s by Vladimir L. Frayshtadt at the holding concern Leninets in Leningrad, but not made public until the 1990s. The basic principle is that Ajax turns the kinetic energy produced by the incoming airflow into chemical energy and power. Hydrocarbon fuel circulating under the skin is decomposed into several constituents by aerodynamic heating (“endothermic fuel conversion”) and routed to a so-called magnetohydrodynamics (MHD) propulsion system, consisting of an MHD genera­tor, a scramjet, and an MHD accelerator. The MHD generator extracts energy and thereby slows down the airflow before it enters the combustion chamber, circum­venting the problems associated with mixing fuel and air at high Mach numbers.

Subsequently, the extracted energy is re-injected into the system by the MHD accel­erator (located behind the combustion chamber) which speeds up the airflow. Another novelty on Ajax is the creation of plasma at the leading and trailing edges of its body to ensure a smoother air flow across the fuselage [25].

The Oryol program was finished in 2001. The general conclusion was that the best way to go forward in the near future was to develop partially reusable TSTO systems with flyback boosters and conventional rocket engines. Including space – planes as a means of satellite deployment in TSTO systems would only be effective if they could lower launch costs by 5-7 times compared with expendable launch vehicles and if they could be made five times more reliable, both of which are unattainable goals at the present time. Therefore, preference was given to TSTO systems with conventional satellite deployment techniques. The partially reusable Angara rockets using the Baykal flyback stage were seen as a first step in that direction. SSTOs were considered worth developing only if their dry mass could be made 30 percent lower than that of systems like the Space Shuttle or Energiya – Buran, which is unrealistic for the time being. The most promising SSTO designs were considered to be vertical take-off/horizontal landing systems [26].

Under the Federal Space Program for 2001-2005 Oryol was followed by another research program called Grif (“Vulture”), focusing among other things on studies of new, heat-resistant materials, construction materials, and air-breathing engines [27]. The latest Federal Space Program (2006-2015) only envisages the development of a partially reusable TSTO system with a flyback booster, an indication that SSTO has been shelved for many years to come. A tender to develop the TSTO is to be held in 2009 and the system is supposed to be fielded in 2016, although this is subject to further review. Payload capacity should be 25-35 tons to low orbit and launch costs should be reduced 1.5 times by avoiding the expenditures associated with clearing first-stage impact zones.

A possible contender is the RN-35, a TSTO system designed by the Keldysh Research Center in 2001-2003. Having a payload capacity of 35 tons, it would have a winged flyback booster burning liquid oxygen and methane. This may eventually be followed around 2030 by the RN-70, a similar system with a 70-ton payload capacity. There may be cooperation with the French CNES space agency under a program known as Ural [28]. At any rate, given the conclusions of the Oryol studies, it is unlikely that spaceplanes will be part of the TSTO program.

Called the Reusable Vertical Landing Transport Ship (MTKVP), the lifting body was a 34 m long vehicle consisting of three main sections: a front section with the crew cabin, a mid-section containing a huge payload bay, and an aft section with orbital maneuvering engines. After using its limited aerodynamic characteristics during the hypersonic stage of re-entry, the vehicle would deploy a series of parachutes at an

altitude of 12 km and a speed of 250 m/s. Vertical landing speed would be dampened with small soft-landing engines and horizontal speed with a ski landing gear.

One of the big advantages of this design was that the ship did not need expensive runways, although some of the plans did envisage landing on a prepared dirt surface. The absence of wings, which are dead weight for most of the flight anyway, also saved a lot of mass. The MTKVP would weigh 88 tons and have a payload capacity of 30 tons to a low 50.7° inclination orbit and a return capacity of 20 tons. Moreover, the MTKVP could rely on proven technologies such as other aerodynamically shaped objects (in particular, the Soyuz descent capsule and nuclear warheads) and parachute and soft-landing systems that had been used for some years by airborne troops to safely land heavy cargos. The idea also was to retrieve the RLA strap-on boosters in a similar fashion, leading to additional cost savings.

However, a major disadvantage of the MTKVP was its low cross-range cap­ability. This was particularly important for the Russian orbiter, because unlike its American counterpart, it could land only on Soviet territory. At a later stage in the design process an attempt was made to improve the vehicle’s cross-range capability from about 800 km to 1,800 km by giving the fuselage a slightly triangular shape. Another problem was that a vehicle of this type would be exposed to extremely high temperatures (about 1,900°C), placing high demands on the heat shield and requiring long turnaround times for repair work. Many doubted if it would be reusable at all. Moreover, since the vehicle was supposed to land in the steppes, it would have been a cumbersome process to recover it and transport it back to the launch site.

The launch vehicle for the MTKVP was known as RLA-130V. It consisted of a 37.4m high core stage powered by two 250-ton thrust LOX/LH2 RD-0120 engines and six 25.7 m high strap-on rockets with 600-ton LOX/kerosene RD-123 engines [54].

Buran had 38 primary thrusters (“Control Engines’’ or UD) (exactly the same number as on the Space Shuttle Orbiter) and eight verniers (“Orientation Engines’’ or DO) (two more than on the Orbiter). Together the primary thrusters and verniers formed the Reaction Control System (RSU). The primary thrusters provided both attitude control and three-axis translation, and the verniers only attitude control. They were used for these functions during the launch, separation from the core stage, on-orbit, and re-entry phases of the flight (up to an altitude of 10 km). If needed, some of the UD thrusters could also act as a back-up for the DOM engines.

The UD thrusters (17D15), built in-house at NPO Energiya, had a thrust of 390 kg and a specific impulse between 275 and 295 s. Unlike the DOM engines, they used gaseous rather than liquid oxygen as an oxidizer. This was obtained with a small turbopump assembly mounted on the ODU LOX tank. First, liquid oxygen from the LOX tank passed through the pump, where its pressure was increased to 78.4 MPa. Then it entered a gas generator where it was ignited with a minute amount of sintin fuel (ratio of 100: 1) to form a mix of gaseous oxygen, carbon dioxide, water vapor, and droplets with a temperature of 60°C. After any residual liquids had been dumped

overboard, the gaseous oxygen was used to drive the turbine and was then stored in separate tanks at pressures ranging from 2.45 to 4.9 MPa. From there it was delivered to the combustion chamber to react with liquid sintin through electrical ignition. Each UD thruster could be fired for a duration of anywhere between 0.06 and 1,200 seconds and be ignited up to 2,000 times during a single mission. The thrusters were designed to sustain 26,000 starts and 3 hours of cumulative firing.

The DO verniers (17D16 or RDMT-200K) provided 20 kg of thrust and had a specific impulse of 265 s. They were developed by the Scientific Research Institute of

Machine Building (Nil Mashinostroyeniya) in Nizhnyaya Saida, which had been a branch of the Scientific Research institute of Thermal Processes (Nil TP) until 1981 and specialized in small thrusters for spacecraft. The RDMT-200K was probably a cryogenic version of the RDMT-200, a thruster with similar capabilities built for the Almaz space station but burning storable propellants. The verniers were similar in design and operation to the UD thrusters, but used liquid oxygen and a different cooling system. They were intended for short-duration burns with an impulse time between 0.06 and 0.12 seconds and could be ignited up to 5,000 times during a single mission. Thrusters based on the RDMT-200K were supposed to fly on the upper stage of the Yedinstvo/ULV-22 rocket, a launch vehicle studied by the Makeyev bureau in the late 1990s to fly from Australian territory.

Aside from the ODU engines and thrusters, Buran had four small solid-fuel motors (thrust 2.85 tons each) to instantly separate the vehicle from the Energiya core stage in case of a multiple engine or other catastrophic launch vehicle failure. Presumably developed by NPO iskra, they were situated in the nose section of the vehicle and should have given Buran enough speed to stay clear of the out-of-control rocket after separation. They were not supposed to be used in a standard separation from the core stage after main engine cutoff. The solid-fuel motors were apparently not installed on the first flight vehicle that made the one and only Buran mission in 1988.

Since the main production facilities for Energiya-Buran were located at great dis­tances from Baykonur, a practical way had to be found of transporting the various elements to the cosmodrome. This was not so much of a problem for the 3.9 m wide strap-on boosters. The first stage of the 11K77/Zenit rocket, which served as the basis for the strap-ons, had already been tailored for rail transport to the launch site and the strap-ons could therefore use the same infrastructure. However, this was not the case for Buran itself and the core stage, which were too big to be transported by conventional means. There was also the problem of returning Buran to Baykonur in case it was forced to make an emergency landing on its back-up landing strips in the Crimea and the Soviet Far East.

One way of avoiding this problem would have been to concentrate the bulk of the assembly work at the cosmodrome itself. In other words, elements of Buran and the core stage would have been transported to the cosmodrome in many small pieces using conventional means of transport and then assembled together at the launch site. This had been done with the massive first stage of the N-1 rocket, the major parts of which were welded together at the cosmodrome. However, for Energiya-Buran this was not considered a viable solution. It would have required the construction of costly new facilities at the launch site and thousands of skilled engineers and workers would have had to be sent away from their home base on lengthy assignments to the cosmodrome.

Transportation by road and/or water was considered, but all proposals were deemed too costly because of the need to perform major construction work and make changes to existing infrastructure. One option studied for the core stage was to transport it by barge over the Volga river from Kuybyshev to the Volgograd/ Astrakhan region and from there to Baykonur over a specially constructed railway.

The only solution left was to transport the elements by air, either by helicopter or airplane. Serious consideration was given to using the Mi-26 helicopter of the Mil design bureau, which had become operational just as the Energiya-Buran program got underway in the second half of the 1970s. The Mi-26 is still the heaviest and most powerful helicopter in the world, capable of lifting about 20 tons. In this scenario, the

orbiter airframe or elements of the core stage would have been mounted on an external platform and then lifted by a combination of two to four Mi-26 helicopters (depending on the mass of the payload). Test flights with the mid fuselage of a defunct Il-18 aircraft were staged from the Flight Research Institute in Zhukovskiy, but showed that this transportation technique was cumbersome and even dangerous. During one test flight the pilots were forced to drop the payload after it had begun dangerously swaying from one side to the other due to air turbulence. Another problem was the helicopter’s limited range, which would have made it necessary to make several refueling stops on the way to Baykonur.

As foreseen by the government decree of 17 February 1976, the Ministry of the Aviation Industry began looking at a number of aircraft to solve the transportation problem. Two airplanes considered were the Tupolev Tu-95 and Ilyushin Il-76, but it soon turned out they would not be up to the task at hand. The most advanced Soviet cargo plane available at the time was the An-22 Antey of the Antonov design bureau in Kiev. In service since 1965, it was capable of lifting 60-80 tons. Engineers studied the possibility of mounting Energiya-Buran hardware on the back of the aircraft or inside by increasing the diameter of its aft fuselage to 8.3 m (the latter version was known as An-22Sh), but both configurations presented insurmountable aerodynamic and stability problems. A more capable Antonov cargo plane, the An-125 Ruslan, was under development in the late 1970s. However, its single vertical fin made it impossible to install long payloads and its relatively small undercarriage could not handle high-crosswind landings with a big payload installed on the back of the fuselage.

By the middle of 1980 preparations had been completed for the long-awaited inaugural test firing of a complete RD-170. Mounted on Energomash’s test-firing stand nr. 2, the engine was ignited on 25 August 1980, but shut down just 4.4 seconds later. It was only the first in a long string of setbacks for the RD-170/171. The next 15 test firings were also less than satisfactory, leading to a decision to perform the 17th test firing at a lower thrust of 600 tons. This resulted in a first successful, full-duration 150-second test firing of the RD-170 on 9 June 1981.

Subsequent test firings at the same thrust rate also produced satisfactory results, giving Energomash engineers enough confidence to move on to ground tests of the nearly identical RD-171 integrated with a Zenit first stage. These tests were carried out at the IS-102 test stand of NIIkhimmash, originally used in the 1950s for testing the first stage of the R-7 missile and later the scene of test firings of the Proton first stage and the second, third, and fourth stages of the N-1. The engine earmarked for the test (serial nr. 18) had already undergone a successful test firing at Energomash’s facilities in September 1981. Later analysis did show that a turbopump rotorblade had been damaged by particles that had somehow entered the turbopump assembly, but this was considered benign enough to press on with the test firing of the Zenit first stage on 26 June 1982. To the amazement of onlookers, the test ended in disaster near the end of its scheduled 6-second duration, when the turbopump assembly burnt through and caused a massive explosion that completely destroyed the stage and the entire test stand.

The disaster raised serious questions about the fundamental design of the RD-170/171, the more so because the test had been performed at only 600 tons of thrust rather than the nominal 740 tons. It led to the creation of an interdepartmental commission to look into the status of the RD-170 development program and consider

possible alternatives for powering the Zenit first stage and Energiya’s strap-on boosters. Headed by Valentin Likhushin, the head of Nil TP, the commission included such luminaries of the Soviet rocket industry as Arkhip Lyulka, Nikolay Kuznetsov, and also Valentin Glushko himself.

One idea, proposed by I. A. Klepikov at Energomash, was to equip each combus­tion chamber with its own, smaller turbopump assembly, transforming the RD-170 into four engines with 185 tons thrust each (hence their designation MD-185, with the “M” standing for “modular’’, because the idea was to use the engine on a variety of rockets). Actually, an order to study such an engine had already come from the Minister of General Machine Building Sergey Afanasyev as early as 11 October 1980. Wary of witnessing a repeat of the N-1 fiasco, Afanasyev had ordered to set up a complete department within Energomash to design such an engine in order to safe­guard against any major development problems with the RD-170/171. It was felt that the 2UKS experimental engine, successfully tested in 1977-1978, could serve as a prototype for the MD-185.

Another option was to use the NK-33 engines developed by the Kuznetsov design bureau (under the Ministry of the Aviation Industry) for a modified version of the N-1 rocket. Although the N-1 had been canceled before the NK-33 engines ever had a chance to fly, forty of these reusable engines had undergone an extensive series of test firings up to 1977, proving their reliability. By making small modifications to the turbopumps, Kuznetsov’s engineers had managed to uprate the NK-33’s thrust from about 170 tons to just over 200 tons, meaning that four would be sufficient to replace the RD-170. Energiya’s chief designer Boris Gubanov flew to Kuznetsov’s plant in Kuybyshev, where he was shown more than 90 such engines lying in storage.

The most radical alternative studied was to replace the Blok-A strap-ons with solid-fuel boosters. That task was assigned to NPO Iskra in Perm (chief designer Lev N. Lavrov), an organization specialized in solid-fuel motors that had already built several small solid-fuel systems for Energiya-Buran. NPO Iskra devised a plan for a 44.92 m high booster consisting of seven segments. Weighing 520 tons (460 tons of which was propellant), the booster would produce an average thrust of 1,050 tons (specific impulse 263 s) and operate for 138 seconds before separating from the core stage.

In the end, none of the three proposals was accepted. Although the MD-185 was probably the least radical alternative, research showed that it would not solve the turbopump burn-through problems as the temperature of the generator gas would be virtually the same as in the RD-170/171. A major problem with both the MD-185 and NK-33 was that they increased the total number of engines on Energiya from eight to twenty, leaving more room for failure.

One can also safely assume that Glushko had second thoughts about using the NK-33 engines. After all his efforts to erase the N-1 from history, it is hard to imagine he would have accepted using engines that had originally been built for this rocket. What’s more, in 1977 Glushko had secured a decision from the Council of Ministers to ban all work on powerful liquid-fuel rocket engines not only at Kuznetsov’s design bureau, but at any organization under the Ministry of the Aviation Industry. Understandably, Kuznetsov was not about to come to Glushko’s rescue just like that.

One of the conditions he laid down for participating in the Energiya program was that his team be officially rehabilitated after the abrupt and humiliating cancellation of its efforts several years earlier.

It was even easier to find arguments against NPO Iskra’s solid rocket motors. Aside from the safety and ecological concerns inherent in solid-fuel rockets, the Soviet Union had no experience in building solid rocket boosters of this size. More­over, they would not have been reusable and it would have been difficult to operate them in the temperature extremes of Baykonur. It would have taken an estimated 8 years to get them ready for flight.

In fact, any of the three alternative proposals would probably have delayed the first flight of Energiya by many years and would only have added to the already soaring costs of the program. In September 1982 the interdepartmental commission decided to continue test firings of improved versions of the RD-170/171 and at the same time continue research work on the MD-185. The official investigation into the June 1982 accident had concluded that it was probably the direct result of the engine being tested in a vertical position (as opposed to the near-horizontal position for the

Energomash tests). However, Energomash engineers disagreed and believed it had been caused either by aluminum particles entering the turbopump assembly from the propellant tanks or by high vibrations of the turbopump assembly.

Among the measures taken to prevent a repeat of the accident were the instal­lation of filters to prevent particles from entering the turbopump assembly and the strengthening of certain components of the turbopump. Those efforts paid off with the first successful full-duration 142-second test firing of the RD-170 at nominal thrust (740 tons) on 31 May 1983, which by many was considered a make-or-break test for the engine. In the following months, the engine performed better and better, clearing the path for another test of the RD-171 as part of a Zenit first stage. Bearing in mind the disastrous outcome of the first such test, a commission was set up to decide if it could proceed. In October 1984 the commission gave a negative recom­mendation (even KB Yuzhnoye chief Vladimir Utkin), but that was overruled by the new Minister of General Machine Building Oleg Baklanov, who had replaced Afanasyev in the spring of 1983 and proved to be a more avid supporter of the RD-170 than his predecessor. In the end, the Zenit first stage operated flawlessly in a test firing at the refurbished IS-102 test stand of NIIkhimmash on 1 December 1984, repeating that performance at the end of the same month.