Category Energiya-Buran

NPO Energiya, the former “Korolyov design bureau”, was the organization in charge of the Energiya-Buran project as a whole, performing a role comparable with that of a “prime contractor” in the West. NPO Energiya was responsible for making all key technical decisions and coordinating work between the numerous organiza­tions. Situated in the Moscow suburb of Kaliningrad (renamed Korolyov in 1996), it was initially part of the N11-88 rocket research institute founded in 1946, but split off from that organization along with Factory 88 to form the independent OKB-1 (Experimental Design Bureau 1) in 1956. It was renamed Central Design Bureau of Experimental Machine Building (TsKBEM) in 1965, NPO Energiya (NPO standing for “Scientific Production Association”) in 1976, and RKK Energiya (RKK standing for “Rocket and Space Corporation”) in 1994. Factory 88 was renamed Factory of Experimental Building (ZEM) in 1967.

Placed in charge of NPO Energiya in May 1974 was Valentin P. Glushko, who thereby relinquished his duties as chief designer of KB Energomash, the rocket engine design bureau that had merged with TsKBEM to form NPO Energiya. Being a member of the Academy of Sciences (since 1953) and a member of the Central Committee of the Communist Party (since 1976), Glushko had considerable political clout and enjoyed almost unconditional support from Dmitriy Ustinov. Initially, Glushko was both “general designer” and “director” of NPO Energiya, but in June 1977 Vakhtang D. Vachnadze was assigned to the newly created post of “general director” to handle the organization’s day-to-day administrative affairs. Glushko died in January 1989 and was replaced in August 1989 by Yuriy P. Semyonov, who was initially only general designer, but also took over the post of general director from Vachnadze in March 1991.

By late 1977 work on the Energiya-Buran project at NPO Energiya was con­centrated in Department 16. Igor N. Sadovskiy was the chief designer of Energiya – Buran as a whole, with Yakob P. Kolyako, the former head of the heavy-lift launch vehicle section, serving as deputy for the rocket, and Pavel V. Tsybin as deputy for the orbiter. There were changes in the wake of a December 1981 party and government decree calling for organizational improvements in the Energiya-Buran program. Responsibility for the orbiter was transferred to design Department 17 of Yuriy P. Semyonov (Soyuz-Salyut), while Department 16 remained in charge only of the rocket. In January 1982 Sadovskiy, who had been on bad terms with Glushko, was replaced as chief designer of Energiya-Buran by Boris I. Gubanov, a veteran of KB Yuzhnoye in Dnepropetrovsk, who had played a key role in the development of missiles such as the R-14, R-36, and R-36M. From that moment on Gubanov was chief designer of the Energiya-Buran system as a whole and also chief designer of the rocket, while Semyonov was chief designer of the orbiter. Sadovskiy became Gubanov’s first deputy, while Vladimir A. Timchenko served as Semyonov’s deputy for the orbiter.

On the production side, NPO Energiya’s ZEM manufacturing facility was in

charge of building many key systems needed for orbital flight—in particular, the orbital maneuvering engines and primary thrusters of the ODU propulsion system as well as the power supply system. These parts were then shipped either to the Tushino Machine Building Factory or to Baykonur for installation in the vehicle. ZEM also housed a full-scale “electrical analog” of Buran (the so-called “Integrated Stand” or OK-KS).

ZEM also manufactured several parts of Energiya’s strap-on boosters. In the mid-1970s an agreement had been reached that KB Yuzhnoye in Dnepropetrovsk would only build the so-called “modular part’’ of the strap-ons—in other words, the part that was common to the strap-ons and the Zenit first stage. Most of what was unique to the strap-ons would have to be built at ZEM—in particular, the nose and tail sections of the boosters, the parachute containers, drain valves, and actuators. According to original plans, final assembly of the strap-on boosters was to take place at ZEM, but later it was decided to move this work to the Baykonur cosmodrome. The parts manufactured at ZEM were delivered to Baykonur by rail and integrated with the modular part in situ at the cosmodrome. Finally, ZEM also manufactured the pneumatic and hydraulic systems for the Energiya core stage. Directors of ZEM during the Buran years were Viktor M. Klyucharyov (1966-1978) and Aleksey A. Borisenko (1978-1999) [2].

Very early on in the program a decision was made to build a runway at the Baykonur cosmodrome not only to receive Buran at the end of its missions, but also to deliver Buran and elements of the Energiya rocket to the cosmodrome by the VM-T Atlant and eventually Mriya. NPO Molniya was assigned as prime contractor for the construction of the runway by a party/government decree on 21 November 1977.

Baykonur has had an aerodrome (“Krayniy”) since the early days of its existence, but this is situated close to the city of Leninsk, many dozens of kilometers to the south of the launch facilities, and was therefore not suited for this role. Requirements for the location of the new runway were that it had to be outside the “blast zone” of the Energiya pads and be capable of receiving Buran from either side, both during nominal missions and in launch emergencies. The new facility (called PK OK or 11P72) was eventually built some 6.5 km to the northwest of the UKSS complex and 11 km to the northwest of the Raskat complex.

The central part of the landing complex was a 4.5 km long and 84 m wide runway called Yubileynyy (“Jubilee”), capable not only of receiving Buran, but also planes with a take-off mass of up to 650 tons. The surface layer was made of reinforced concrete with a thickness varying between 26 and 32 cm above an 18 to 22 cm sand/ cement ground layer. This concrete, which was about 1.5 to 2 times stronger than the type used on ordinary runways, was produced in six factories located at a consider­able distance from the runway. This created serious transportation problems since the concrete could remain in liquid state for only one and a half hours before being poured onto the runway. The surface had to be extremely flat, with deviations of no more than 3 mm over a 3 m stretch (compared with 10 mm on ordinary runways). To achieve this, the complete 378,000 m surface of the runway had to be ground like parquet floor with special milling machines.

At either end of the runway was a 500 m long and 90 m wide stretch of asphalt to give Buran more leeway during emergency landings. Running parallel to the main runway at a distance of some 50 m was a 4.5 km long and 100 m wide dirt runway apparently intended for emergency landings by planes, with no role in the Buran program.

Adjacent to the runway were several facilities:

– A platform to drain liquid oxygen, gaseous oxygen, and liquid hydrogen from Buran’s fuel cells and the ODU propulsion system.

– A platform to off-load Buran and elements of the Energiya rocket from their carrier aircraft. This has two mate-demate devices called PKU-50 and PUA-100 capable of handling payloads of 50 and 100 tons, respectively.

– A “waiting platform” for vehicles needed to service Buran after landing.

– A parking platform for airplanes.

– An airplane-servicing area.

Also located in the vicinity of the runway was the ground segment of the Vympel navigational aid system (Vympel-N). This included six transponders for the RDS system (only three of which were required for landing), one beacon for the RSBN system, four beacons for the RMS microwave landing system, and a set of radars.

The nerve center of the landing complex was a six-story high command and control building (OKPD) that acted as a control center for the landing phase, work­ing in conjunction with the TsUP Mission Control Center near Moscow. The build­ing had one big control room for Buran and another for ordinary air traffic control tasks [16].

By mid-1989, several months after Buran’s maiden flight on 15 November 1988, plans were finalized for a second mission that would far exceed the first one in complexity. The mission would use the second flight vehicle (2K, sometimes called “Buran-2”) and was therefore dubbed 2K1. The plan was for the orbiter to be launched unmanned and fly to the Mir space station, where it would dock with the axial APAS-89 docking port of the Kristall module. Before that, Kristall would be relo­cated from its lateral port on the Mir multiple docking adapter to the station’s front axial port. After docking, the Mir resident crew would board the orbiter to determine the state of its on-board systems, with one of the possible objectives being to use the vehicle’s remote manipulator arm to move a payload from the payload bay to Kristall’s lateral APAS docking port. One NPO Energiya official said that the pay­load was a small one-ton module housing a Fosvich X-ray telescope similar to the one on Mir’s Kvant module. See [69]. Also installed in the payload bay would have been a pressurized module (37KB) about the size of the Kvant module with instrumentation to record various flight parameters.

Subsequently, the orbiter would undock and continue its flight autonomously. Around the same time, a manned Soyuz equipped with an APAS-89 docking port would be launched to dock with the orbiter. The crew would transfer to the orbiter and perform one day of testing. After the Soyuz undocked, it would fly on to Mir to link up with Kristall, while the unmanned 2K orbiter returned back to Baykonur after a one-week mission [70].

In the late 1980s NPO Energiya was ordered to build three Soyuz spacecraft (serial numbers 101, 102, 103) with APAS-89 docking ports. These vehicles were intended in the first place for possible rescue missions to stranded Buran crews during the test flight program, but it was decided to use the first one in the framework of the

2K1 mission [71]. The flight was partially seen as a dress rehearsal for such a potential rescue mission.

LII demanded that at least one of its Buran pilots be included in the Soyuz crew to give him the necessary experience for the first manned Buran mission [72]. With no or few Soyuz seats available in the mainstream Mir program, this was the ultimate opportunity for a Soyuz familiarization flight, the more so because it involved Buran itself. However, in 1990 a training group was formed for the Soyuz mission consisting of three GKNII pilots and three TsPK military engineers:

Stepanov and Fefelov were assigned in April 1990 and the others in October/ November 1990. It is not entirely clear if training advanced to the point that actual crews were formed, although Kadenyuk has claimed he was in the second back-up crew with Fefelov [73]. The most active training was performed by the three pilots, who faced the unprecedented task of docking Soyuz with Buran. All three spent many hours in TsPK’s Soyuz simulators, practicing dockings both with Buran and Mir. The three engineers reportedly never underwent any dedicated mission training [74].

The 2K1 mission was originally scheduled for 1991, but kept slipping as future prospects for the Buran program grew ever dimmer. Officially, the three pilots and Illarionov remained assigned until March 1992, and Fefelov and Stepanov until October 1992 [75]. Kadenyuk has said the mission was officially canceled in August 1992 [76].

Soyuz craft nr. 101 was eventually launched as Soyuz TM-16 on 24 January 1993, carrying another resident crew (Gennadiy Manakov and Aleksandr Poleshchuk) to the Mir space station. Equipped with an APAS-89 docking port, it was the only Soyuz vehicle ever to dock with the Kristall module. Soyuz “rescue” vehicles nr. 102 and 103, which had been only partly assembled, were modified as ordinary Soyuz TM spacecraft with standard “probe” docking mechanisms and were given new serial numbers [77].

Testing of Buran’s ODU propulsion system was the prime responsibility of the so-called Primorskiy Branch of NPO Energiya in the Leningrad region on the shores of the Gulf of Finland. This was set up in 1958 as a branch of Glushko’s OKB-456, mainly to test engines with exotic propellants such as the RD-301 fluorine/ ammonia engine destined for a Proton upper stage. When OKB-456’s successor KB Energomash merged with TsKBEM in 1974 to form NPO Energiya, the Primorskiy Branch became part of the new conglomerate and remained subordinate to it even after Energomash regained its independence in 1990. Its first assignment as part of NPO Energiya was to test the RD-120 engine for the second stage of Zenit. The old RD-301 test stand was refurbished for a series of horizontal test firings of 11D58M engines for the Proton rocket’s Blok-D upper stage in 1978­1982, which were probably seen as precursors to similar tests with the Orbital Maneuvering Engines (DOM or 17D12) for Buran. Between May 1985 and Septem­ber 1988 six 17D12 engines underwent 114 horizontal test firings lasting a total of 22,311 seconds.

Meanwhile, in 1981 construction had begun of a new vertical test stand called V-1 to test complete ODU engine units called EU-597, containing not just the 17D12 engines, but also thrusters and verniers. The first such ODU unit (nr. 10S) began testing in June 1986 but was destroyed in a fire in February 1987, seriously damaging the test stand. V-1 was refurbished for a series of tests with a new unit (nr. 12S) between September 1987 and April 1988 that underwent the complete ODU firing program planned for the first Buran mission. Those tests uncovered a problem that would delay the Buran flight for several months (see Chapter 7). More tests were conducted with unit nr. 31L between June and December 1988 and unit nr. 11S between January 1991 and March 1993. After cancellation of the Energiya-Buran program the unit was mothballed and eventually removed from the test stand. The 17D15 thrusters and 17D16 verniers apparently also underwent individual tests at Nllkhimmash near Zagorsk. Test firings of the ODU integrated in Buran were conducted at Baykonur’s test-firing platform [14].

The Auxiliary Power Units (VSU) underwent a test program at the IS-104 and IS-105 test stands of Nllkhimmash, which included simulated hydrazine leaks to test the fire suppression system. The VSU hydrazine tank was put to the test in simulated weightless conditions aboard an Ilyushin-76 aircraft and also at various ^-levels at Star City’s TsF-18 centrifuge. The VSU test program culminated in the units being installed on Buran and activated at the Buran test-firing stand at Baykonur.

The next step in the launch preparation process was for Buran to be mated with its launch vehicle (Energiya rocket 1L) for an experimental roll-out to pad 37. The 1L rocket had always been well ahead of Buran in its launch preparations. Assembly of the core stage in the Energiya assembly building had begun back in October 1986, shortly after work with the core stage for vehicle 6SL had been completed. In early 1988 (14 January-2 February) the 1L rocket had already spent about three weeks on pad 37 for a variety of tests, including firing tests of the hydrogen igniters and retraction tests of the various platforms connecting the launch towers with the rocket.

The Energiya 1L-Buran stack arrived on the pad in the third week of May (the roll-out date has been given both as 19 May and 23 May). Once again a multi­tude of tests were performed, although none of them involved actual fueling of the rocket or the orbiter. One goal of the pad tests was to see if various sources of electromagnetic radiation at Baykonur did not interfere with the operation of on­board systems. The main problems uncovered during the pad tests were with the interaction between the orbiter and rocket computers and with the ground software needed to analyse telemetry at the cosmodrome and in Mission Control.

Actually, the pad tests in May-June were only part of a broader series of exercises at the cosmodrome intended to simulate pre-launch and post-landing operations, including numerous off-nominal situations. Involved in the exercises were not only the launch and recovery teams, but also the LII pilots, who simulated automatic landings on board Tu-154LL aircraft, with the MiG-25-SOTN performing the role of escort aircraft as it would during Buran’s final descent. The exercises also offered the opportunity to test virtually the entire communication network for the mission, including tracking stations, Mission Control in Kaliningrad, and orbiting communications satellites. Ground crews rehearsed post-landing operations and were trained how to deal with a return-to-launch-site abort during ascent. For this purpose, the OK-MT Buran mock-up was transported to the Yubileynyy runway.

The Energiya-Buran stack returned to the assembly building after about 3-4 weeks of tests (the roll-back date has been given both as 10 June and 19 June). Apparently, the original plan was for the orbiter and rocket to undergo some additional tests and then return to the pad for launch in the summer of 1988. Internal planning documents show that in early 1988 the launch was scheduled for July [37]. However, program managers felt that several problems that had surfaced during testing over the preceding weeks needed to be dealt with and decided to remove Buran from the rocket and return it to its MIK OK processing facility.

The most serious problem had cropped up in April during test firings of an ODU propulsion module at the Primorskiy Branch of NPO Energiya near Leningrad. A valve used in the liquid-oxygen gasification system of the primary thrusters failed

to close when commanded to do so, a problem that could jeopardize the operation of the thrusters in flight. Because of this and other issues with the ODU, it was deemed necessary to remove Buran’s ODU module and partially disassemble it to carry out modification work. This also required changes to the flight software, which had already been adapted numerous times in the preceding months, a penalty the

Russians had to pay for flying Buran unmanned. In the end, Buran went into orbit with the 21st version of the flight software.

After repairs to the ODU and integrated electrical tests with the final version of the flight software, Buran was moved back to bay 4 of the Energiya assembly building on 29 August for reintegration with Energiya 1L. With that work complete, the stack was rolled over to the nearby MZK building on 13 September for a series of hazardous and other operations. These included various loading operations (kerosene for the Buran propulsion system, hydrazine fuel and nitrogen gas for the Auxiliary Power Units, ammonia for the thermal control system, air for the cabin repressurization system), installation of batteries aboard Buran, solid-fuel separation motors on the strap-on boosters, and pyrotechnics for the Buran/core stage separation system.

Finally, the large doors of the MZK were opened in the early hours of 10 October and four diesel locomotives began pulling the impressive 3,500-ton combination of Energiya, Buran, and transporter to launch pad 37. In an old tradition, coins imprinted with the roll-out date were placed on the rails before the assembly passed

by and collected afterwards as souvenirs. It took the assembly some 3.5 hours to inch its way to the launch pad. Then another three hours were required to place the stack into vertical position and another hour to connect the Blok-Ya launch adapter to the launch table. All was now ready for final launch preparations to begin [38].

By mid-1991 the 2K1 mission had slipped to 1992 from its original launch date in the first quarter of 1991. Beyond that Buran was now scheduled to take part in the assembly and operation of the Mir-2 complex, where the emphasis would be on the industrial production of ultra-pure medicines and semiconductor materials and also on remote sensing. The plans were presented in detail by Yuriy Semyonov at the congress of the International Astronautical Federation in Montreal in October 1991.

First, the 2K orbiter would go up again in 1993 on an unmanned solo flight (2K2) to test some of the biotechnological installations to be flown under the Mir-2 pro­gram. Then in 1994 the 1K vehicle would fly the first manned mission (1K2) as part of a plan sometimes light-heartedly referred to as “Mir-1.5”, in which Mir would gradually be replaced in orbit by Mir-2. After the launch of the Mir-2 core module by a Proton rocket, Buran would rendezvous with the module, grab it with its two remote manipulator arms, and dock it to a bridge in the cargo bay. Buran would then

link up with a small docking module on Mir’s multiple docking adapter and again use its manipulator arms to transfer the Mir-2 core module to a lateral docking on Mir previously occupied by the Spektr module. The two modules would remain docked for about two years. After the transfer of the Priroda Earth resources module to the Mir-2 core, Mir and its remaining add-on modules would then have been undocked and discarded, setting the stage for the four-year assembly of the Mir-2 complex (1996-2000).

Before that, in 1995, vehicle 2K would be launched on another autonomous flight (2K3) to test a biotechnological module called 37KBT, based on the original 37KB instrumentation modules. With the emphasis having shifted from fundamental scientific research to biotechnological production, the original plans for the 37KBI scientific add-on modules had been scrapped in late 1989. Buran would now regularly fly two biotechnological modules (37KBT nr. 1 and nr. 2), carrying one up and bringing the other down.

Between 1996 and 2000 there would be two missions annually, one using vehicle 2K to swap out the 37KBT biotechnological modules (2K4, 2K5, 2K6, 2K7, and 2K8) and another using the 1K orbiter for assembly and logistics missions (1K3,1K4, 1K5, 1K6, 1K7). Planned for addition to Mir-2 was a 37KBE “power module’’ equipped with extra solar panels. Further Buran missions would have been required to add a large 85 m truss structure to Mir-2 and outfit it with solar arrays, large radiators, and an array of scientific instruments [30].

The “Mir 1.5’’ plan was dropped in 1992, when it was decided that Mir-2 would

only be launched after Mir had outlived its usefulness. This would also allow the new station to be placed into a higher inclination orbit (65° vs. 51.6° for Mir) for better remote-sensing coverage. At this point the big Buran-launched 37KB-type modules were abandoned in favor of smaller modules based on the Zenit-launched Progress – M2 cargo ship. The new Mir-2 concept was approved by the Council of Chief Designers in November 1992. Although it left open the option of launching the add-on modules and the station’s truss structure with Buran, Zenit was clearly the preferred option. By the time Mir-2 was merged with Freedom to become the Inter­national Space Station in late 1993, work on Buran had been suspended.

Bolstered by the success of the maiden Energiya launch in 1987, NPO Energiya worked out a series of ambitious plans for future use of the rocket. Taking into account the changing international climate, those missions focused not so much on national, but global needs. Some of these projects bordered on the realm of science fiction and were way beyond even the generous budgets of the Soviet days, which is why the Russians were clearly counting on international partners to join them. The following missions were studied in 1987-1993:

– A constellation of 30 to 40 satellites to restore the depleted ozone layer by aiming laser beams at the stratosphere, causing excited oxygen molecules to break up under the influence of solar radiation and to recombine into ozone molecules. Weighing 60 to 80 tons each, the satellites would have flown in Sun-synchronous orbits at an altitude of 1,600 km, using electric propulsion systems to maneuver from their initial insertion orbits. Using this satellite constellation, it would have taken an estimated 30 years to solve the ozone depletion problem.

– Containers with radioactive waste to be placed into heliocentric graveyard orbits between Earth and Mars at a distance of approximately 1.2 astro­nomical units from the Sun. Weighing 50 tons each, the hardened containers could house 6 to 9 tons of radioactive waste. It was estimated that 10 to 15 Energiya missions would be required annually to dispose of the 100 tons of high-level radioactive waste produced around the world each year. Each container was to be boosted to an 800 km parking orbit by a conventional upper stage before being sent on an escape trajectory by a nuclear electric propulsion system.

– A constellation of solar reflector satellites to illuminate the polar regions, provide energy from space, and improve crop yields by stimulating photo­synthesis. With each of the satellites weighing 5-6 tons, a single Energiya was capable of placing a cluster of 10 to 12 such satellites into a low parking orbit with the help of an upper stage. A reusable, solar electric interorbital space tug would have boosted the satellites to a 1,700 km polar orbit inclined 103° to the equator. Each satellite had a 10 year lifetime and would be usable 8 hours daily, illuminating a 17 km diameter circular area on the Earth’s surface.

– An Earth-to-Moon shuttle service to collect helium-3 on the lunar surface for use in nuclear fusion reactors.

– 20-ton environmental monitoring satellites in geostationary orbit. Using the same UKP platform as the Globis satellites, they would monitor the Earth with optical, infrared, and microwave remote-sensing instruments, study Sun-Earth relations with ultraviolet spectrometers and particle detectors, and relay data from low-orbiting satellites in radio and optical wavelengths.

– 30-ton UKP-based satellites in 600 km polar orbits to monitor observance of international disarmament treaties and perform remote-sensing tasks such as studies of natural resources and environmental monitoring. The 12-ton payload would have included a videospectrometer, optical electronic cameras, and phased-array antennas.

– Satellites to clear the geostationary belt of space debris. Equipped with an engine unit and grappling devices, they would each spend about half a year in 0° to 14° inclination orbits at geostationary altitude, moving defunct satellites and debris to graveyard orbits.

– A 27-ton space-based radio telescope to provide Very Long Baseline Inter­ferometry (VLBI) in concert with ground-based radio telescopes. Called IVS (International VLBI Satellite), this was a joint Soviet-European project put forward in response to a 1989 Call for Mission Proposals for the second medium-size mission under ESA’s Horizon-2000 program. The IVS was to consist of NPO Energiya’s UKP bus and a European-built 20 m diameter radio telescope. With an inclination of 65° and a perigee of 6,000 km, the apogee would be varied from an initial height of 20,000 km to 40,000 km and 150,000 km over the satellite’s five-year operational lifetime. IVS was picked along with five other projects for further assessment in 1991, but was not approved for further development. Had it been selected, it could have flown in 2001 [62].

Even though the Skif-DM launch had demonstrated that Energiya was capable of being used as a heavy cargo carrier, Buran-T failed to gain impetus, mainly due to a lack of interest from the military, who were supposed to be the system’s main customers. A government decree in August 1985 had ordered the Ministry of Defense to work out “technical requirements” for Buran-T and Vulkan in a three-month period and NPO Energiya to prepare a draft government decree on these systems in the first quarter of 1986, outlining their objectives and setting a timeline for their development. The draft was sent for review to the VPK by July 1986 and called for starting Buran-T flights in 1988, with the introduction of the Smerch cryogenic upper stage expected in 1995. It was not until December 1987, one and a half years later, that the VPK responded by rejecting the draft, claiming it had not been agreed upon with the military. For the military a rocket could only be declared operational if there was a concrete payload for it, which was hardly the case for Buran-T. Eventually, the military even withdrew their “technical requirements” for Buran-T [63].

Even as the newly created NPO Molniya got down to Buran development in 1976, the Mikoyan bureau contingent in the organization seemingly had a hard time parting with the air-launched Spiral concept. In fact, one NPO Molniya veteran recalls that

Lozino-Lozinskiy was never overly enthusiastic about Buran, which had been forced upon him from above, and that his real passion remained with air-launched systems [3]. Realizing that one of the major drawbacks of Spiral had been the need to develop a futuristic hypersonic aircraft, the Mikoyan designers began drawing up plans for spaceplanes launched from existing subsonic transport planes. The aim was to expand their missions beyond military reconnaissance and offensive operations to satellite deployment/retrieval and space station support. Unlike Buran, such space – planes would be suited to launch payloads usually orbited by expendable launch vehicles and had many other advantages such as quicker turnaround, more launch flexibility, and a wider range of attainable orbits. The new air-launched concept benefited heavily from experience gained in the Spiral, BOR, and Buran programs.

The RLA’s LOX/kerosene engines were to be the first such engines developed under Glushko in almost 15 years. In the mid to late 1950s Glushko had supervised the development of the RD-107/RD-108 engines for the R-7 missile and derived launch vehicles (sea-level thrust around 80 tons) and the RD-111 for the R-9ICBM (sea-level thrust 144 tons). All of these were four-chamber LOX/kerosene engines using an open combustion cycle, in which the gases used to drive the turbopumps are vented overboard. This system is also known in Russian terminology as “liquid-liquid”, because both the fuel and the oxidizer are injected into the combustion chamber in a liquid state. However, the development of the RD-111 was plagued by serious prob­lems, including high-frequency oscillations in the combustion chamber, intermittent combustion, and the need to protect the chambers and nozzle walls from overheating.

In the early 1960s Glushko turned his attention to closed-cycle engines, in which the gases used for driving the turbines are routed to the combustion chamber to take part in the combustion process. This, together with the increased chamber pressure, produced much higher specific impulses than had been obtained earlier. One of the propellants entered the combustion chamber in a liquid form and the other in a gaseous form (which is why this system is also called the “gas-liquid” system by the Russians). Given the painful experience with the RD-111, Glushko was wary of using LOX/kerosene for these even more powerful engines. Instead, he decided to con-

centrate on storable propellants based on unsymmetrical dimethyl hydrazine (UDMH), which he had already mastered while developing open-cycle engines for the R-12, R-14, and R-16 missiles. In fact, Glushko’s preference for storable over cryogenic propellants can be traced back all the way to his years as a rocket pioneer at the GDL and RNII rocket research institutes in the 1930s.

All this had dire implications for the N-1 program. Glushko’s reluctance to build closed-cycle LOX/kerosene engines and Korolyov’s refusal to use the highly toxic

storable propellants for the rocket effectively ended the cooperation between the two chief designers. It forced Korolyov to rely on LOX/kerosene engines of the much less experienced OKB-276 Kuznetsov design bureau in Kuybyshev (which actually were of the closed-cycle type).

For the remainder of the 1960s, Glushko was mainly engaged in building closed – cycle engines with storable propellants for a variety of missiles and launch vehicles of the Chelomey and Yangel bureaus. Except for the R-7 derived rockets, all Soviet space launch vehicles that were operational around the turn of the decade (Kosmos, Tsiklon, Proton) were powered by such engines. They had been derived from nuclear missiles, which traditionally use storable propellants to enable them to be launched at short notice.

Energomash didn’t end its boycott on LOX/kerosene engines until the late 1960s, by which time enough experience had been gained with the closed-cycle combustion principle for engineers to feel confident enough to apply it in powerful LOX/kerosene engines. An opportunity to build such an engine arose in 1969, when the Chelomey bureau drew up plans for a mammoth rocket called UR-700M, intended to send Soviet cosmonauts to Mars. One version of the rocket that Chelomey looked into would have 600-ton thrust LOX/kerosene engines in the first and second stages. In 1970 Glushko’s engineers worked out plans for such an engine called RD-116 or 11D120, which presumably was a modified LOX/kerosene version of the single­chamber RD-270, a hypergolic engine earlier planned for Chelomey’s (unflown) UR-700 Moon rocket [31]. Although the UR-700M remained no more than a fantasy, Energomash was reportedly also ordered to investigate the possibility of using the same engine on the first stage of the N-1, which had suffered two launch failures in 1969. A small cluster of RD-116 engines would be enough to replace the N – 1’s thirty NK-15 first-stage engines [32].

In the end the idea was dropped because it would also have implied a radical redesign of the N-1 rocket. However, it does seem to have whetted Glushko’s appetite to continue studies of such engines, the more so because a new policy was emerging in the early 1970s to abandon storable propellants in favor of cryogenic and hydro­carbon propellants in new, dedicated space launch vehicles.

As a result, work on high-thrust LOX/kerosene engines at Energomash resumed in earnest in 1973. The studies focused not only on standard kerosene, but also an advanced synthetic hydrocarbon fuel known as tsiklin or sintin. Based on furfural and propylene, it had a higher specific impulse than ordinary kerosene, but was also much more expensive.

In the course of 1973 proposals were presented for single-chamber, two-chamber, and four-chamber versions of a 500+ ton thrust LOX/kerosene engine. There was serious debate between the proponents of the single and four-chamber versions, which both had their advantages and drawbacks. A key meeting at Energomash in the second half of 1973 opted for the four-chamber version. After all, Energomash had had experience with multi-chamber engines since the 1950s. Furthermore, there had been numerous problems with the development of the 640-ton single-chamber RD-270 for the UR-700. Finally, by using four smaller combustion chambers it would be easier to test them by modifying test models of existing combustion chambers for storable propellants. The meeting also approved a so-called “modular design” for the engines, making it possible to use them in a standardized fleet of rockets [33].

Still, all these were no more than internal decisions within Energomash that didn’t stand much chance of being implemented until the bureau merged with TsKBEM to form NPO Energiya in May 1974 and Glushko got the opportunity to advance his RLA idea. But even at this stage there was no consensus what the LOX/kerosene engines should look like. Some of the disagreements centered around such things as the pressure in the combustion chamber and the type of combustion cycle. Some claimed the pressure in the combustion chamber shouldn’t exceed 200 atmospheres, making the engine more reliable. However, a lower pressure translates into bigger combustion chambers and less payload, and the compromise reached was to have a pressure of 250 atmospheres. Others felt the engine should use a fuel – rich combustion cycle, lowering the risk of turbopump burn-throughs. That was countered by the argument that an oxidizer-rich preburner engine is more efficient and easier to reuse because it leaves behind less soot residue [34].

There was also more fundamental debate over the thrust of the engine. Some felt that the task of building a four-chamber engine with a single, powerful turbopump assembly was too challenging and instead preferred single-chamber engines in the 150-ton thrust range with smaller, individual turbopumps. In other words, rather than having a handful of very powerful engines, it would be better to install a large number of low-thrust engines [35]. One concern with the high-thrust engines was that they would expose the rocket to serious vibrations in case of a sudden emergency shutdown, making it necessary to strengthen the rocket’s structure and lower its payload capacity [36].

Bearing in mind these two schools of thought, two design departments at Energomash got down to studying engines in two thrust classes. Department 729 focused on engines ranging in thrust from 112.5 to 263.5 tons: the RD-128, RD-129, and RD-124 for the first stage of the RLA family and the RD-125, RD-126, and RD-127 for the second and third stages. Department 728 initially concentrated on an engine with a phenomenal thrust of 1,003 tons (the RD-150), but then scaled back its ambitions to a 600-ton thrust engine called RD-123 [37]. This is the engine that finally got selected in 1975 for use in the first stage of the Soviet space shuttle stack and the progenitor of the eventually developed RD-170. A determining factor in this choice must have been the negative experience of flying many low-thrust engines on the first stage of the N-1. Moreover, Glushko must have feared that if the choice did fall on the low-thrust engines, there would have been attempts to de-mothball the Kuznetsov bureau’s N-1 engines rather than introduce his new LOX/kerosene engines. However, the debate would flare up again in the early 1980s when the RD-170 was plagued by serious development problems (see Chapter 6).

ENERGIYA CORE STAGE

The core stage was designed jointly by NPO Energiya in Kaliningrad and its Volga Branch in Kuybyshev, with manufacturing taking place at the Progress factory in Kuybyshev. With a length of 58.7 m and a maximum diameter of 7.75 m, it was the backbone of the Energiya-Buran stack, providing structural support for attachment with the strap-on boosters and orbiter. It was very similar in design to the Space Shuttle’s External Tank (ET), with the exception of a tail section housing the engine compartment. The core stage was made up of an upper liquid oxygen (LOX) tank, an unpressurized intertank, a lower liquid hydrogen (LH2) tank, and a tail section containing the four RD-0120 engines. The wet mass was 776 tons.

Both the LOX and LH2 tanks were made of a 1201 aluminum alloy. The 552m3 LOX tank could hold about 600 tons of oxidizer. It consisted of a forward ogive section (itself made up of three sections), a cylindrical section (itself made up of two sections), and a spherical aft dome. All sections were welded together. The tank had anti-slosh baffles to dampen any motions of the LOX that might throw the rocket off course. The LOX feed line exited the LOX tank at a 7° angle to the longitudinal axis of the tank to facilitate oxidizer supply during the final moments of the launch. It ran to the tail section right through the LH2 tank. This is a major difference with the Space Shuttle External Tank’s LOX feed line, which emerges from the ET’s intertank area to convey the oxidizer to the aft right-hand ET-Orbiter disconnect umbilical. The LOX feed system had gas accumulators to dampen longitudinal oscillations (“pogo’’). These were located in the lower part of the LOX feed line in the bottom of the LH2 tank and also in the engines’ turbopump inlet ducts.

The intertank was the structural connection joining the liquid hydrogen and oxygen tanks. Flanges were affixed at the bottom and top of the intertank so the two tanks could be attached to it. Also installed in the intertank was the instrumenta­tion for the core stage’s flight control system. Prior to launch the intertank was

Energiya core stage 93

purged with nitrogen gas to prevent the build-up of moisture and explosive mixtures of hydrogen and oxygen gas.

The 1,523m3 LH2 tank, which could hold about 100 tons of liquid hydrogen, consisted of spherical aft and forward domes and a large cylindrical section. The tank walls were machined in a waffle-grid pattern, something not employed in the ET hydrogen tank until the introduction of the Super Lightweight External Tank in 1998. Although LH2 is so light that sloshing does not induce significant forces, Energiya’s LH2 tank, unlike that of the ET, did have anti-slosh baffles.

Just like the Shuttle’s ET, the Energiya core stage was covered with a combina­tion of polyurethane spray-on foam insulation (Ripor-2N, PPU-17) and ablative material (PPU-306) for thermal insulation and thermal protection. This reduced boil-off losses during the countdown, maintained the propellants at the proper temperatures for normal engine operation, limited ice formation on the outer surface, and protected the core stage against the flames from the strap-on boosters’ separation motors. The original plan not to apply thermal insulation to the upper part of the LOX tank was abandoned due to fears that ice might break off from that part of the core stage and damage the orbiter’s fragile heat shield. Various non-destructive methods were used to test these materials after they were applied: electric methods to check their thickness, radioisotope techniques for density, and acoustic methods to detect debonding.

Shedding of tank insulation became a big issue in the US after the February 2003 Columbia accident, caused by a piece of foam insulation breaking off the tank and inflicting lethal damage to one of the Reinforced Carbon-Carbon panels on the Shuttle’s left wing. Russian sources do not mention whether Energiya’s foam insula­tion was less or more prone to shedding simply because this was not a matter of major concern in the pre-Columbia days. Moreover, any foam loss that might have occurred on the two Energiya launches in 1987 and 1988 would have been virtually impossible to photographically document because the first launch took place in darkness and the second in poor weather conditions. Tile damage suffered by Buran on its sole mission has usually been attributed to ice falling off the core stage and the launch pad and not to foam impacts.

Electric power for the core stage was provided by four simultaneously operating turbogenerators driven by air, nitrogen, hydrogen, and helium gas. In order to simplify the design and reduce mass, common plumbing was used for all four gases. Each generator weighed 330 kg and provided 24kWt of power.

The cryogenic propellants were loaded at lower temperatures than on the Space Shuttle (—255°C vs. —253°C for the liquid hydrogen and —195°C vs. —182°C for the liquid oxygen). This made the propellant denser and also significantly reduced boil – off losses. Techniques for subcooling liquid oxygen were pioneered by the Russians with the R-9 missile in the early 1960s, but Energiya marked the first use of subcooled liquid hydrogen. The liquid hydrogen was subcooled by passing it through two double-walled cooling devices in which tubular heat exchangers were immersed in a bath of liquid hydrogen boiling at reduced pressure.

Loading of the core stage began several hours before launch with a slow-fill mode to condition the tanks. When the core stage was 2 percent full, the fueling process was

sped up to 19,000 liters per minute for the liquid oxygen and 45,000 liters per minute for the liquid hydrogen. This fast-fill mode continued until 98 percent of the pro­pellant was loaded. Topping off continued until T — 3m02s for the LOX tank and T — 1m52s for the LH2 tank.

After the start of fueling, electrically powered pumps in the main engines began to circulate the liquid hydrogen in the fuel tank through the four engines and back to the tank to chill down the liquid hydrogen lines, ensuring that the path was free of any gaseous hydrogen bubbles and was at the proper temperature for engine start. The LH2 was recirculated to the tank rather than returned to ground facilities because it loses a lot of pressure during the circulation process. The engines’ LOX lines were also thermally preconditioned, but the LOX used for this purpose was dumped overboard.

In the final minutes of the countdown the tanks were pressurized to maintain their structural integrity during launch, minimize the build-up of volatiles in the tanks, and to prevent cavitation of the main engine low-pressure boost pumps. Pre-launch pressurization was performed with ground-supplied helium and began at T — 2m23s for the LOX tank and T — 1m20s for the LH2 tank. After lift-off the LOX tank was pressurized with hot gaseous oxygen produced by heat exchangers in the main engines and the LH2 tank with gaseous hydrogen tapped from the turbines of the main engine LH2 boost pumps. At lift-off the LOX tank was pressurized to 2.6 atmospheres and the LH2 tank to 3.1 atmospheres. During launch the pressure in the LOX and LH2 tanks was maintained between 1.41-1.55 atmospheres and 2.25-2.39 atmospheres, respectively.

Each tank had a dual-function vent and relief valve at its forward end. It could be opened by ground-supplied helium before launch for venting or by excessive tank pressure for relief during launch. Excess hydrogen gas left the core stage via the intertank area, while excess oxygen gas was directly vented overboard [1].