Russian Liquid Propellant Engines

5/27/2018 Russian Liquid Propellant Engines

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NFM Special Issue No. 10/2000 Page 2

RD-170 FAMILY

The RD-170 liquid oxy-gen/kerosene rocket engine was de-signed in 1974-1986 at NPO Ener-gomashby a team led by Chief De-signer V.P.Radovsky for the side

engine units (first stage) of theURKTS Energiya-Buran Heavy

Multipurpose Space Rocket Trans-portation System. Until today, theRD-170 remains the worlds mostpowerful and quite promising enginedesign.

RD-170 is a four-chamber, re-coverable liquid-propellant rocketengine with the fuel turbopump-feedsystem. The design provides forhigh-pressure afterburning of the ex-haust turbine gas in the main com-

bustion chambers. The ignition offuel components is ensured by self-igniting starting fuel.

The engines fuel system in-cludes the Main Turbopump Unit(MTU) incorporating a single-stage active ax-ial-flow turbine co-shafted with rotors of theoxidizer and the fuel pumps. The turbopumpassembly also includes two screw boostpumps: the oxidizer pump driven by the gasturbine, and the fuel pump driven by the hy-

draulic turbine. The pneumatic control systemincludes: pressure helium bottles, solenoid-operated electro-pneumatic valves and systempiping.

According to Mr. Feliks Chelkis, Chief

and Lead Designer of the RD-170 rocket en-gine, there is hardly another national or West-ern liquid-propellant engine showing same per-formance in terms of full-thrust gainingsmoothness, which makes the loads applied tothe engine and the launch vehicle considerably

lower.Engine designers say development of a

recoverable high-thrust power unit is a chal-

lenging task. To ensure controllable throttlingto full thrust and cut-off smoothness, it is vitalthat consumption of propellants is adequatelyregulated by means of powerful electro-hydraulic devices.

The described engine went through full-

scale development, bench trials and flight tests.Ground test-firing of the engine was conductedon a fully built-up unit.

As development works continued on theRD-170 engine, the designers encountered fireaccidents in the engines hot gas-flow duct re-sulting in that MTU assembly bearings weredestroyed, in addition, there were high-frequency vibrations in the combustion cham-ber. The fire accidents were found to havebeen caused by the presence of foreign materi-

als (like chips or fillings) in the gas ducts. Cer-tain bottlenecks in the turbopump assembly op-

Review Of R ussias Advanced Liquid -PropellantRock et Engines

RD-170


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Table 1.RD-170 vs. F-1: COMPARED PERFORMANCE

Performance Characteristic RD-170 F-1*

Engine configuration Closed-loop Open

Oxidiser Liquid oxygen Liquid oxygen

Fuel Kerosene Kerosene

Oxidiser/fuel ratio 2.6 2.2

Combustion chamber pressure, atm 250 80

Ground thrust, tf 740.2 690.1

Vacuum thrust, tf 806.7 792.7

Specific impulse on ground, s 308 265.4

Specific impulse in vacuum, s 337 304.8

Fuel consumption, t/s 2.4 2.6

Nozzle area ratio 33.5 16.1

Engine mass, t 5.2 7.9

Engine mass-to-thrust ratio, kg/tf 6.45 9.97

Height, m 4.0 5.6

Diameter, m 3.8 3.8

Turbopump assembly horsepower 220,000 60,000

Rated duration, s 140 150*) F-1 engine was developed by Rocketdyne in 1959-1967 for Saturn-5s first stage.

eration had to be overcome, as well. Yet, thehigh-frequency vibration able to cause the in-

jector assembly to destroy was something thedesigners had to address again and again tillthe engine was just about to be commissioned.

The combustion chamber developmentphase included as many as 300 bench firingtrials of summarized duration of 23,000 s; atotal of six different cooling system optionswere tested along with more than 20 injectordesign options for a variety of nozzle parame-ters. A replacement of the combustion chamberhead does not require the engine to be totallydisassembled. The complete RD-170 enginecan be re-started at least 20 times before totalengine overhaul is required.

So far, RD-170 has had a record of over120,000 s in bench firings, showing high reli-ability. Program managers, however, seem tobe not quite satisfied with the already attainedreliability level, and have been planning devel-opment of a more reliable and more economi-cal RD-170 design option for reduced fuel con-sumption.

The RD-170 engine is designed to ensure

failure-free operation, i.e., provide normalfunctioning under any operating conditionswithout failures that might result in any disas-trous consequences for the LVs or the powerplants relevant systems, like hydraulic drivepiping destruction or seal failure.

The design of the RD-171 engine (amodification of RD-170 installed in Zenit-2LVs first stage) incorporates four pivotedcombustion chambers, unlike the RD-170 ver-sion for the URKTSSystem, where two cham-bers are pivoted. Pivoting is ensured by auxil-iary power units (APU) operating on hydrazinedecomposition products. Each booster unit ofthe URKTS Systems first stage includes oneAPU; similar APUs were installed on the Bu-ranorbital spacecraft to provide for actuationof the aerodynamic control devices (like elev-ons, flaps, rudder) and landing gear. Table 1below compares performance of RD-170 andthe United States F-1 liquid-propellant engine.

Starting from late 80s, agencies responsi-ble for the national space program have beendisplaying interest in exporting RD-170 to gainmuch-desired convertible currency. After per-


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Table 2.RD-170M, RD-180, RD-191 ENGINES PERFORMANCE

Performance Characteristic RD-170M RD-180 RD-191

Engine configuration Closed-loop, w/afterburning

Fuel components

Ground thrust, tf 780.5 375.1 188

Vacuum thrust, tf 807.4 408.2 204.6

Combustion chambers 4 2 1

Height, m 4.0 4.0 3.9

Liquid oxygen/kerosene

mission had been granted to sell the engineabroad, representatives of NPO Energomashturned their eyes westwards immediately. In aninterview Mr. F.Chelkis said the engine did notrequire additional modification to suit anyspecified rocket and could be offered for sale

at any time out of the available stock. How-ever, according to NPO leaders, RD-170 willnot to be supplied for military applications.

Top managers withEnergomashdid theirbest to make sure the issue of technology trans-fer was addressed adequately as RD-170 wasoffered for sale, and assumed that the design ofthe liquid propellant engine being sold couldbe protected as a commercial secret.

Mr. F.Chelkis said in end-1991 that theRD-170 engine augmented by Energomashs

manufacturing and testing capability coulddramatically reduce the expenditure by the

US on their national space program. However,apart from the obstacles resulting from tradingrestrictions, any application of a liquid-propellant rocket engine of that power hasbeen and remains limited by non-existence in

the United States of a launch vehicle in whichit could be incorporated.

Engineers with Pratt & Whitney areknown to have been considering a two-stage,non-recoverable medium-class LV with RD-160 as the first stage engine. Provided the sec-ond stage of the described LV includes thelox/kerosene RD-120, the system could beused to inject max. 12.7 t payload into a loworbit, i.e., to show performance similar to Ze-nit-2. In the meantime, with the advanced

RD-171

RD-180


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Performance Characteristic RD-120 RD-120K

Engine configuration Closed-loop, w/afterburning

Oxidiser Liquid oxygen

Fuel Kerosene


Ground thrust, tf - 80


Specific impulse on ground, s - 304.4

Specific impulse in vacuum, s 350 330.5

Specific fuel consumption, kg/s 242 263

Combustion chamber pressure, atm 166 179.8

Nozzle area ratio 114.5 4.96

Engine mass, kg 1,285 1,080


Height, m 3.87 2.435

Diameter, m 1.95 1.080


Turbopump assembly r.p.m. 19,000 2,220

Guaranteed duration, s 2,220 2,220

Table 3.RD-170 vs. RD-120K: COMPARED PERFORMANCE

lox/hydrogen RL-200 liquid-propellant rocketengine by Pratt & Whitney as the second stagemotor, the payload capacity could be raised toas much as 22.45 tons causing the LV to sur-pass the performance of Proton. Finally, withthe second stage provided with two traditionallox/hydrogen RL-10A-4 engines (an advancedmodification of Atlas-Centaurs second stageengine) the PL capacity could reach 9.75 tonsfor low orbit missions, placing the LV in be-tween SoyuzandZenit-2 LVs.

Another potentially feasible applicationfor RD-170 is the engines incorporation intoliquid-propellant boosters designed to replaceSpace Shuttle Systems solid-propellantbooster units. According to engine designers,that would enhance the launch vehicles flexi-bility and reliability owing to the added thrustcontrol and liquid-propellant engines in-flightcut-off capabilities not available to solid-propellant rocket engines. In addition, liquid-propellant boosters show greater environment-friendliness and provide for Space Shuttlesgreater payload capacity in low orbit injec-RD-191


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RD-120

tions.The RD-170 liquid-propellant rocket en-

gine combined with an external expendablefuel tank and three SSME lox/kerosene enginesof the Space Shuttle System could provide ba-sis for the development of a heavy launch vehi-cle designed to bring payloads into low orbits.A rocket of that configuration with two liquid-propellant boosters could put a payload of70.76 tons into orbit, while six boosters wouldraise the payload capacity to as high as 176tons.

Various RD-170 design options havebeen under development since late eighties foradvanced national and foreign launch vehicles.Specifically, the thrust of the basic liquid-propellant rocket engine was supposed to beincreased by raising the chamber pressure from250 to 263.5 atm, which required insignificantmodification of certain components so as toreduce the overall engine mass. The describeddesign option was proposed for Zenit-2 LVmodification intended to carry both unmannedspacecraft, like the cargo Progressor modulesof orbital stations, and launch manned trans-portation spacecraft like Soyuz.

It was at that time that designers withNPO Energomash proposed a new enginebased on RD-170. The advanced RD-180

could be used on both the new or advanced,national or foreign LVs and as a liquid-propellant booster for the Space Shuttle Sys-tem. The design of the RD-180 engine includestwo of the four combustion chambers availableon the predecessor, and several standard or up-graded units and assemblies of the basic designengine, like the turbopump assembly or otherfuel feed assemblies. The Russian-US RD-

AMROSS Joint Venture made up by RussiasNPO Energomashand the United States Pratt& Whitney of United Technologies is knownto be manufacturing the engines for the Ameri-can advanced launch vehicles of LockheedMartins Atlas-3A and Atlas-5 family (see

NFM Special Issue 8/99 for details). Currentstock of orders counts as many as 18 engines(Editors note: Lockheed Martin intends to buy

from Energomash a total of 101 engines overthe next 10 years), and there is a possibility toinitiate manufacturing of RD-180 in the UnitedStates for installation on launch vehicles de-signed to orbit governmental payloads.

Furthermore, work is underway at NPOEnergomash to develop the single-chamberedRD-191 engine based on the RD-170 combus-tion chamber design, new automatics and newturbopump assembly. Brief performance char-acteristics of the modified RD-170 engine arepresented in Table 2 along with RD-180s andRD-190 performances.

RD-120/RD-120K

The development of the RD-120 liquidoxygen/kerosene engine that was presentedoriginally at the Aviadvigatel-91 air engineshow was led between 1976-1986 by ChiefDesigner V.P.Radovskyunder the Statement OfWork issued by KBYuzhnoyeDesign Bureau.The engine was intended for the second stageof the medium payload capacityZenit-2launchvehicle earmarked as future replacement of theSoyuzfamily launch vehicles.

The RD-120 is a high-altitude enginewith turbopump-assisted fuel feed system. Thedesign of the engine provides for exhaust tur-bine gas afterburning. Fuel components are ig-nited by means of self-igniting starting fuel.

The engine includes the combustionchamber with dual-component (gas-fluid) in-

jectors, the high-altitude nozzle, the turbopumpassembly incorporating the oxidizer gas gen-erator in which the entire liquid oxygen (i.e.,


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the oxidizer) is beingconverted and heated,and the automaticstart-up and opera-tion control systems.Thermal insulation ofthe combustionchambers wall is en-sured by the fuel-enriched wall bound-ary layer generatedby the chamber injec-tor head, and twofilm cooling belts(slots).The engine wentthrough full-scale de-velopment, bench tri-als and flight tests.Ground test-firing ofthe fully built-up en-

gine with a nozzle was conducted on the ex-haust diffuser test bench providing for simu-lated high-altitude operation of the engine.

Similar to what happened as RD-170 wasdeveloped, the designers of RD-120 encoun-tered fire accidents in the engines hot gas-flow passage, failure of the turbopump assem-bly bearings and high-frequency vibrations inthe combustion chamber. Some experts wouldexplain the accidents by the less than adequatedesign of the combustion chamber injector.

The RD-120K engine, a modification ofRD-120 distinguished for a shorter nozzle, dif-fers from the RD- 120 design in modified lay-out of the engines top. Indeed, the RD-120version installed in the central opening of the

Zenit-2 second stages toroidal tank lookssomehow compressed on its sides and ex-tended vertically. Unlike that, the RD-120Kmounted on LVs bottom face is shorter anddiffers in the layout of the top components, in-cluding the combustion chamber, turbopumpassembly, gas generator, piping and automat-ics.

The RD-120K developed as part of homeinitiative byNPO Energomashis designed foruse in the first stages of future light- and me-dium-class launch vehicles. To support theproposed applications, the engine is providedwith both the incorporated thrust control capa-bility (3.5 percent throttling up or 8.5 percent

throttling down) and the mixture ratio variation

capability ( 10 percent). The engine can be

fixed or hinge-mounted, as required to ensure

thrust-vector control capability within 6 de-

grees.

The described engine is proposed for in-stallation on lower stages of the advancedSoyuz-2 (Russ Project) and Yedinstvo (Unity)launch vehicles developed, accordingly, by theProgress Design Bureau and the Makeyev. Acomparison of RD-120 and RD-120K perform-ance is provided in Table 3.

RD-161/RD-161P

The experimental RD-161 and RD-161PV engines, which are among the latest de-

velopments by NPO Energomash, were origi-nally presented at the K Zvezdam 93 (To theStars-93) show.

The liquid oxygen/kerosene RD-161 en-gine offered by Energomash is a home initia-tive development designed for liquid-propellant engine applications in upper stagesof launch vehicles and/or orbital transfer vehi-cles (Block DM), including both the existingand future vehicles. The design of RD-161P inwhich the propellants are high-concentration

(93-97 percent) hydrogen peroxide and kero-sene is based on an earlier prototype initiallydesigned for the orbital maneuver system oftheMAKSProjects spacecraft.

Both engines are single-chambered unitswith turbopump fuel-feed system and exhaustturbine gas afterburning. The ignition of thefuel components in the RD-161 engine is en-sured by plasma generated as high-voltage,high-frequency current is passed through gase-ous oxygen (a development by RKK Energia

initially implemented in Burans steering mi-croengines).The engines include practically identical

combustion chambers with dual-component(gas/fluid) injectors, high-altitude nozzles andturbopump assemblies. However, there is dif-ference in the gas generation flow duct of thedescribed liquid-propellant engines.

The gas generator duct of the RD-161 en-gine is a dual-component, single-injector unit(with rotary-type fuel injection and peripheraloxidizer atomization). The system is designedto operate under rich oxidizer conditions as itgenerates hot gas at approx. 550oC essentially

RD-120K


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Table 4.RD-161 vs. RD-161P: COMPARED PERFORMANCE.

Performance Characteristic RD-161* RD-161P

Engine configuration

Oxidiser Liquid oxygen Hydrogen peroxide

Fuel Kerosene Kerosene


Vacuum thrust, tf 2.03/2.0 1.5/2.5**

Specific impulse in vacuum, s 365/360 319

Specific fuel consumption, kg/s 5.5.6 7.84


Nozzle area ratio 19.25/18.75 16.28

Thrust duration, s 900 900

Total firings per mission Max. 17 50+

Hinge angle, degrees 6 -

Engine mass, kg 141/119 105

Engine mass-to-thrust ratio, kg/tf 69.46/59.5 51.22

Height, m 2.205/1.700 1.450

Diameter, m 1.020/0.780 0.540

Turbopump assembly horsepower 444 500

Turbopump assembly r.p.m. 80,000 -*) numerator/denominator values show performance of engine with/without nozzle extension;**) numerator/denominator values show single/dual-component mode performance

composed of oxygen and a certain percentageof water steam and carbon dioxide gas.

The gas generator duct of the RD-161Pengine is a single-component catalytic system:hydrogen peroxide passed through the catalystpack is decomposed to produce hot steam-and-gas at approx. 850oC, containing basicallysteam of water and oxygen.

Upon working on the turbopump assem-bly blades, the steam-and-gas is fed to thecombustion chamber for afterburning bymeans of the fuel. Owing to the single-component design of the gas generator, the fuelsupply and engine start-up system are substan-tially simplified. There is a feature worth beingmentioned specifically if RD-161P is to be de-scribed: in case no fuel is fed to the combus-tion chamber, the engine will operate in the so-called single-component mode and develop re-

markably high thrust. The greatest difficulty inthe development of the fuel supply system wasencountered as the materials for the gas gen-

erators catalyst package were selected. Table4 presents compared performance of the RD-161 and RD-161P engines.

NPO Energomash have been out ofsmall-size engine development business for al-most forty years. Yet, designing a small liquid-propellant rocket engine with high specific im-pulse fuelled by environment-friendly fuel

components is really on the agenda.In reviewing engines operating on the hy-drogen peroxide/kerosene pair of fuel compo-nents, one cannot but mention the family ofBritains Gamma liquid-propellant rocket en-gines developed in the early and mid-sixtiesand installed on the Black Knight high-altituderocket and the Black Arrow launch vehicle.Perhaps, no one else but the British can boastexpertise in designing, development and pro-duction of peroxide engines.

The scientific and production potentialavailable to the Energomash Association al-lows new technology to be mastered within the


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Table 5.NK-33 vs. NK-43: COMPARED PERFORMANCE

Performance Characteristic NK-33 NK-43

Engine configuration Close-loop. w/afterburning

Oxidiser Liquid oxygen

Fuel Kerosene

Oxidiser/fuel mixture ratio

Ground thrust, tf 154 -

Vacuum thrust, tf 171 179

Specific impulse on ground, s 298 -

Specific impulse in vacuum, s 331 346

Specific fuel consumption, kg/s 517.3 517.3


Nozzle exit section pressure, atm 0.55 0.13

Engine mass, kg 1,340 1,400




Thrust duration, s 600 600

2.8

shortest possible time if the required liquid-propellant rocket engine is to be created. Let usnot forget the available expertise with experi-mental engines operating on hydrogen perox-ide/kerosene and hydrogen perox-ide/pentaborane fuel: those were developedand tested byEnergomashin the mid-sixties aspart of the Soviet Lunar program.

The RD-161 includes an original carbonplastic nozzle extension, some 500 mm inlength. The radiation-cooled extension pro-vides for the engines Isp to be increased ap-proximately 5-fold.

The RD-161 engine should be describedas an advanced unit: indeed, even though themore energy-capacious Synthine synthetic hy-drocarbon fuel is not being used, the engineshows the highest specific impulse among allavailable Lox/kerosene engines. Experts be-lieve turning to Synthine fuel is possible in thefuture: despite certain potential problems withcooling, the new fuel could add at least 10units to the engines specific impulse parame-ter.

In the proposed configuration, the RD-161P engine is quite small-sized, essentiallybecause the nozzle is relatively short. Despitethe fuel-saving scheme including generator gas

afterburning, the 161P version provides com-paratively low specific impulse (in fact, it iseven lesser than that of similar thrust level en-gines by KBKhM fuelled by the nitrogentetroxide/asymmetric dimethyl hydrazine pairof fuel components). Representatives ofEner-gomash claim their RD-161P peroxide-fuelledliquid-propellant rocket engine is intended toprove feasibility of a relatively simple small-size recoverable motor to be fuelled by non-toxic, environment-friendly propellants, thatcould provide for essentially unlimited dura-tion of space operation. No task to attain anyultimate specific impulse was set, for as longas the value could be increased by 10-15 unitsin the future, if required.

NK-33/NK -43Many Russias and international compa-

nies have been displaying major interest latelyin the NK-33 and NK-43 rocket engines by

Dvigately NK of Samara. The engines werepresented to public for the first time atAviadvi-gatel-90show.

The two products are Russias first everrecoverable liquid-propellant rocket enginesoriginally developed more than 25 years ago.Since then, the engines have passed full-scale


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bench trials to confirm the initially incorpo-rated performance. Yet, neither was ever in-stalled on any craft, even despite batch produc-tion was in place.

Prototype NK-33 lox/kerosene engineknown to broad public as NK-15 was devel-oped between 1962-1970 at the KuznetsovDe-sign Bureau in Kuibyshev for the first stage ofthe super-heavy Lunar N-1 launch vehicle.

N-1s second stage included NK-15V en-gines - the high-altitude modification of thefirst stage liquid-propellant rocket motor, dif-fering from the original engine mainly bygreater nozzle area expansion ratio.

Flight tests of N-1 LV showed the firststage engines required additional developmenteffort, let alone inefficiency of the adoptedquality control practice applied to the liquid-propellant rocket engines being supplied.Those factors stimulated KuznetsovBureau al-ready in mid-1970 to initiate development ofbasically new recoverable liquid-propellant en-gines for longer service life.

The NK-33 engine was developed be-tween 1970-1974 based on a novel work state-

ment accounting for the initial flight test re-sults of N-1 rocket. The idea was to dramati-cally improve reliability, ensure failure-freeperformance and safety of the engine withoutchanging the layout, though somewhat boost-ing the thrust and raising the specific impulse.Interdepartmental tests of the first stage en-gines (the NK-33) and the second stage en-gines (NK-43) were finally completed in Sep-tember 1972.

Engine recovery capability allowed eachcommercial engine to be proof-tested first andretested later as part of the overall rocket sys-tem.

Similar to its prototype, the NK-33 engineis a single-chambered unit with a turbopumpfuel feed system designed for environment-friendly non-self-inflammable fuel (the fuel iskerosene, the oxidizer is liquid oxygen). Theunit is a close-loop engine with the exhaust tur-bine gas afterburned at high pressure in themain combustion chamber. The propellant forthe turbopump assemblys turbine is the burn-ing products of the main propellants burned inoxidizer-rich environment. Practically the en-tire oxidizer is gasified in the gas generatorwith just a little fuel added.

The NK-33 engine differs from the proto-type NK-15 unit by a simplified hydro-pneumatic scheme, more advanced automaticsand improved turbopump and combustionchamber assemblies. For instance, the numbersof igniter automatic components incorporatedin the design dropped from twelve to only asfew as seven. Plug-in connections and inter-changeability of parts add to the enginesmaintainability.

Certain NK-33 production units deliveredthe thrust of almost 205-207 tf on test bench asthe gas generator burning was intensified(which manifested itself in temperature rise);that meant those engines were to be classifiedas units of a different thrust standard. Thethrust control range (varied between 50 and105 percent) for NK-33 was dictated essen-tially by the available service life of the en-gine. With the required service life reducedonly insignificantly, the thrust control rangecould be increased to as high as 135 percent.

Despite the design of NK-33 includes astarting pyroturbine, that engine has lesserweight compared toEnergomashsRD-253 in-

NK-33 mock-up


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Table 6.NK-39 vs. NK-31: COMPARED PERFORMANCE

Performance Characteristic NK-39 NK-31

Engine configuration Close-loop. w/afterburning

Oxidizer Liquid oxygen

Fuel


Engine mounting Fixed Hinged


Specific impulse in vacuum, s 352 353

Specific fuel consumption, kg/s 116.1 116.1


Engine mass, kg 700 722


Nozzle exit section diameter, m 1.3 1.4

Rotation speed, r.p.m. 32,000 32,000

Thrust duration, s 1,200 1,200

Kerosene

stalled in the first stage of the Proton launchvehicle. Indeed, NK-33 engine may well be de-scribed as one of the worlds best in terms ofmass-to-thrust ratio.

In its turn, the NK-43 engine is theworlds most powerful high-altitudelox/kerosene rocket engine. The engines top(combustion chamber, turbopump assembly,automatics and the initial section of the nozzle)is similar to that of the NK-33. However, thenozzle area expansion ratio is much greater.Engine reliability was verified by increasingtotal firings to as many as ten.

According to some foreign experts, theremarkable performance of the engines manu-factured in Samara allows one to assume large-scale co-operation of foreign partners with theRussian manufacturer is quite possible, start-ing, for instance, from joint Russian-US designefforts for new liquid-propellant rocket enginesto direct installation of engines byN.D.Kuznetsov on the existing and future for-eign rockets.

As an example, by replacing two boosterengines currently installed on Atlas LV withtwo NK-33s, payloads injected to transfer or-bits for eventual geostationary orbiting couldbe increased from 3,630 kg to 4,173 kg.

By agreement withDvigately NK, Aerojet

General doing marketing research for the en-gines in Western markets performed five benchtrials of NK-33 at their test facility in Sacra-mento.

So far,Dvigately NKhave been known tohave orders in stock for delivery of NK-33 andNK-43 engines for future K-1 launch vehiclesby the United States Kistler Aerospace, aswell as for YamalLV (see NFM 12/98 for de-tails) by Energia and the Air Launch devel-oped by the corporation of the same name (see

NFM 15/2000).

NK-31/NK-39

The attractiveness of the NK-31 and NK-39 engines by Dvigately NK is explained byessentially the same reasons as is the case withNK-33 and NK-43.

Development of the prototype NK-31 en-gine, known to public as NK-9 and designedoriginally for the first stage of the GR-1 globalrocket by Korolyov OKB-1, began in 1959.The GR-1s second stage included a high-altitude version of the same liquid-propellantengine designated as NK-9V. That latter onewas in fact the immediate predecessor of NK-39 and NK-31.

After development works on GR-1 werediscontinued, the NK-9V version with ex-


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Table 7.KVD-1 vs. RL-10A-3: COMPARED PERFORMANCE

Performance Characteristic KVD-1 RL-10A3-3*

Engine configuration Close-loop Close-loop

Turbopump assembly gas generationmethod

Dual-component gas generator Hydrogen conversion in the cool-ing jacket

Control elements drive Electropneumatic Electromechanical

Ignition Pyrotechnical Electric spark


Gas generator pressure, atm 82.3 -

Cruising chamber thrust, kgf 7,100** 6,830

Specific impulse, s 462 444

Fuel consumption, kg/s 15,37 15.39

Components ratio 6.0 5.0

Dry engine mass, kg 282 132

Height, m 2.14 1.78

Diameter. m 1.58 1.0

Turbine rotor speed, r.p.m. 42,000 30,240*) developed in 1958-68 by Pratt & Whitney for upper stages of Atlas-Centaur and Saturn-1

**) the thrust of each of the two steering chambers: 200 kgf

tended service life was used in the upper (the3rdand the 4th) stages of the super heavy LunarN-1 carrier rocket. In order to suit for the newlaunch vehicle, the engine power was degradedthrough a reduction of the turbopump opera-tion intensity.

The NK-39 engine for N-1s third stage,and the NK-31 for the same rockets fourthstage were developed between 1970-1974based on a work statement issued anew to al-low for the initial flight test results of N-1rocket and improve reliability, ensure failure-free performance and safety of the enginewithout changing the layout. Interdepartmentaltests of the engines were completed by 1972.

Engine recovery capability allows eachcommercial unit to be proof-tested first and re-tested later as part of the overall rocket system.

The NK-31 engine is a single-chamberedunit with a turbopump fuel feed system de-signed for environment-friendly non-self-inflammable fuel (the fuel is kerosene, the oxi-dizer is liquid oxygen). The unit is a close-loopengine with the exhaust turbine gas afterburnedat high pressure in the main combustion cham-ber. The propellant for the turbopump assem-blys turbine is the burning products of the

principal fuel components burned with highexcess of oxidizer.

The NK-31 and NK-39 engines differfrom the prototype NK-9V unit by a simplifiedhydro-pneumatic scheme, more advanced auto-matics and improved turbopump and combus-tion chamber assemblies. Plug-in connectionsand interchangeability of parts add to the en-gines maintainability.

Pitifully, similar to NK-33 and NK-43,neither the prototype engine, nor NK-39 orNK-31 have never been flown realistically inany LV, even despite numerous proposals weremade both before and after the described liq-uid-propellant engines were used in the N-1Project. Certain interest in the engines has beenrecently shown by national and foreign compa-nies (including a business from France plan-ning installation on the VEHRA demo air-craft).

KVD-1

So far, the one and only oxygen/hydrogenliquid-propellant rocket engine in Russiaknown to have passed through full-scaleground testing routine has been the KVD-1 en-gine. Even despite more than 35 years have


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gone since the time its development was firstinitiated, the engine has never been tested in

flight. In the meantime, that is precisely theone Indian Space Research Organization se-lected for the cryogenic upper stage of theirGSLV rocket.

KVD-1s prototype known as 11D56wasdeveloped between 1965-1972 by KhimmashDesign Bureau under a work statement issuedby S.P.Korolyovs OKB-1 Bureau. It was ini-tially designed for the fourth stage (theboost/deboost system) of a future version ofheavy Lunar N-1 launch vehicle. Bench trials

of the engine commenced in 1966.KVD-1 engine is a single-chambered unitwith a turbopump system designed to feed pro-pellants; unlike RL-10A-3 - a similar US liq-uid-propellant engine of the same thrust stan-dard - KVD-1s design includes afterburning: afeature characteristic of any powerful Russianliquid-propellant rocket engine design. Yet,while a gas generator producing oxidizing tur-bine gas (the so-called sour gas, i.e., containingexcess oxygen) is used to drive turbopump as-

semblies of national non-hydrogen close-loopengines, the case is different with KVD-1s gasgenerator, where the propellant is burned with

excess fuel to produce reducing(the so-called sweet) gas.See Table 7 for compared per-formance of KVD-1 and RL-10A-3 engines.KVD-1 includes a single-shaftturbopump assembly. To makethe oxidizer and the fuel pumpscomply with their cavitationcharacteristics at equal speed ofthe turbopump shaft, boosterpumps are provided in the design.With built-in vane boost pumpsincorporated in the propellantcomponents feeder lines, the de-scribed liquid-propellant enginecan operate with outside, externaldrive boost oxidizer and fuelpumps delivering the requiredpropellant components pressureto ensure cavitation-free opera-tion of the main turbopump as-sembly pumps.Ignition of the propellant compo-nents during engine starting is at-

tained by initially igniting the propellants inthe gas generator (by means of pyrotechnicaldevices); thereafter, the gas flowing off the gasgenerator is fed to the turbine and later - to thecombustion chamber for afterburning, wherelacking oxidizer is added.

Despite no specific purpose of the KVD-1engine was disclosed within the framework ofcurrent Russian space program, an apparentassumption is that the unit can be used in cryo-genic upper stages designed to put payloadsinto high-altitude elliptical, geostationary or-bits or escape trajectories.

The first ever cryogenic upper stage withKVD-1 will be 12KRB - the third stage of In-dias GSLV designed and built by Russian spe-cialists of the KB Salyut Bureau andKhrunichev. The one to follow will be theKVRB alias the fourth stage of the future Pro-ton-M LV which is to replace todays Protonwith its lox/keroseneDMblock. Utilization ofthe oxygen/hydrogen propellant instead of theoxygen/kerosene is anticipated to add some 30-55 percent to payload capacity of those LVs assatellites are put into geostationary orbits.

In addition, the oxygen/hydrogen upperstage can make a good substitute for the DM

KVD-1 (left) and its precursor 11D56


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Table 8. S5.92 ENGINE PERFORMANCE

Performance Characteristic High-Thrust Low-Thrust

Engine configuration Open

Oxidizer Nitrogen tetroxide

Fuel Asymmetric dimethyl hydrazine

Oxidizer/fuel ratio 1.95-2.05 2.0-2.1

Cruise chamber thrust, kgf 2,000 1,400

Exhaust nozzle thrust, kgf 40 19

Specific impulse, s 327 316

Fuel consumption, kg/s 6.12 4.43


Gas generator pressure, atm 118 61


Total thrust duration, s 2,000

Overall dimensions, m 0.677 x 0.838 x 1.028

Dry engine mass, kg 37.5

block of todaysZenit-3LV (at Sea Launch) ora future modification of same, thus allowingthe LVs capabilities to be dramatically en-hanced during high-altitude elliptical and/orgeosynchronous satellite launches.

S5.92The prototype S5.92 engine was devel-

oped as part of the Soviet N-1-L3 Project andwas to be installed on the "Zh" block or thepower unit designed for departure of themanned Lunar spacecraft integral to the N-1 -L-3 system from the Lunar orbit back to Earth.KRD-1 engines were also developed in parallelwith the described engine to provide lift-offcapability from the Lunar surface and returnmission to Earth for the return rockets of theautomatic space stations Luna-16, -20, -24; inaddition, development works continued on theKTDU-417 engines designed to provide mis-sion trajectory correction and Moon landing ofthose stations. In fact, the mentioned engineshad much in common, even despite their pro-pellants differed.

The development of todays S5.92 beganin 1978. The engine is designed for the multi-functional rocket unit integratable into a vari-ety of spacecraft, including Fobos-1, -2 and

Mars-96. The power plant of which S5.92 is apart was to boost up an automatic interplane-

tary station after separation from the final stageof the LV to as fast as the escape velocity andto ensure numerous flight trajectory correc-tions, deboosting, Martian orbit entry and orbitmaneuvering.

Though the missions of the mentionedstations failed, nothing could be said to repri-mand the power plant, and a proposal followedto use the S5.92 engines in both the existingand the future power plants of spacecraft andlaunch vehicles.

The single-chambered S5.92 engine is anopen layout unit with the propellants fed by aturpopump. The turbine is operated by themain propellants, the exhaust is effectedthrough fixed steering nozzles. A feature of theengine is the unusually hinged (rather thangimballed) mount of the chamber to ensure itsplane-parallel motion inside the power plant.The approach provides for the engine thrustvector to be displaced relative to the center ofmasses, which in the described power plant isfound very close to the engine head (under thesituation, gimballing would not provide thearm required to generate the moment of thrust).S5.92 engine is operational in two modes: thehigh-thrust mode and the low-thrust mode.High-thrust mode is applied in major speed al-teration maneuvering, while low-thrust mode is


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RD-0110

applied in maneuvers that require precisionspeed pulse variation.

Table 8 presents S5.92 performance char-acteristics

Among the features of the describedS5.92 engine (that are similar to actually anyliquid-propellant rocket engine by Khimmash)are the extremely small dimensions, low spe-cific mass with quite high thrust and specificimpulse values. By the said parameters, the en-gine outruns the similar class RD-161M byEn-ergomash, even despite the more perfect close-loop configuration with afterburning is used inthe latter. Portability of liquid-propellant en-gines by Khimmash is basically achieved ow-ing to optimized combination of the turbopumpassembly parameters, the combustion chamberpressure and the nozzle area expansion ratio. Inaddition, an open configuration allows a liq-uid-propellant engine wherein the vertical di-mensions would not exceed the dimensions ofthe combustion chamber.

Most similar Western analogies to S5.92are the liquid-propellant engines developed inthe late eighties for the US Air Force ASPSupper stage. Those are: XLR-132 by Rocket-dyne and Transtar-1 by Aerojet Tech Systems,both operating on the AT - Aerozin-50 propel-lant (a mixture of 50 percent hydrazine and 50percent asymmetric hydrazine). However, nei-ther of the two engines ever went beyond thetest bench prototype.

The S5.92 engine has been installed on

the Fregat multipurpose rocket unit by La-vochkin(see NFM 4/2000 for details) - the onedesigned for use in the existing (Proton, Soyuz)

and future (Russ) LVs as the top stage for pay-load injections into a variety of orbits, includ-ing solar-synchronous, high-altitude elliptical(geo-transfer), high-altitude circular(geostationary) orbits and/or mission trajecto-ries to the Sun, the Moon, planets, comets orasteroids. The Fregatupper stage can performas a cruise power plant for spacecraft, orbitaltransfer vehicles, orbital or orbital-landingmodules.

RD-0124The second half of the eighties was

marked at the KBKhA Design Bureau forChemical Automation of Voronezh by com-menced development of the advanced versionof a Lox/kerosene engine to replace the stan-dard RD-0110 mounted in the third stage of theSoyuz family LVs. The work went on withinthe frame of the national space program andtransformed in the early nineties in the Russ(Soyuz-2) Project.

The basic version of the RD-0124 enginedesigned to replace RD-0110 is a four-chamberunit. The pivot chambers are used to provide

Model of RD-0124M


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Table 9.RD-0110 vs. RD-0124: COMPARED PERFORMANCE

Performance Characteristic RD-0110 RD-0124

Engine configuration Open Close-loop

Propellant components

Vacuum thrust, tf 30.0 30.0Specific impulse in vacuum, s `320 353

Total combustion chamber 4 4

Liquid oxygen/kerosene

launch vehicle control at the third stage leg ofthe flight. Staying within the same dimensionsas they switched over to close-loop configura-tion with exhaust turbine gas afterburning inthe main chambers and increased chamberpressure and nozzle area expansion ratios, en-gine designers from Voronezh succeeded inraising the specific impulse by as many as 33units, which is equivalent to adding 950 kg toSoyuzpayload capacity.

The engine has been through a compre-hensive cycle of test trials, and preparatorywork is going on now to launch batch produc-tion at the Voronezh Mechanical Plant.

An attractive option (for a foreign cus-tomer, in particular) is the single-chamber ver-sion of the engine designated RD-0124M com-plete with assemblies (turbopump, gas genera-tor, piping, automatics) borrowed from thefour-chamber version. The chamber is com-pletely new, while the nozzle borrowed from asubmarine-launched ICBM can act as either ahigh-altitude or a ground unit.

Some words also must be said about the

most popular problem of using hydrogen forrocket fuel. Hydrogen is called the rocket fuelof the future. In this field the vast experiencehas been accumulated by the KBKhA designBureau. The Russian RD-0120 Engine is not at

all inferior to the American SSME, which

powers the Shuttles.The engine is designed with the staged

combustion cycle (the turbine gas is suppliedto the main injector). RD-0120 is an advancedliquid rocket engine which combines high per-formance, long service life, and low operationcost.

The engine allows reusability applicationsand considerable thrust throttling. The enginehas a system of diagnostics and safety system.

The engine operates with the ecologically

tidy propellants - liquid oxygen and liquid hy-drogen.

The engine provides helium heating forthe oxidizer tank pressurization, producesgaseous hydrogen for the fuel tank pressuriza-tion, and on-board power supply systems.

During the phase of development the en-gine was fire-tested 800 times with total testingduration of 170000 sec. Eight engines weretested on theEnergia launcher.

RD-0120 Performance features:

Vacuum thrust 200 tf

Vacuum specific impulse 455 kgf- s/kg

Propellants LOX/LH2

Com bustion Cham ber pressu re 228 kgf/cm 2

Operating dur ation 500 c

Total engine weight 345 0 kg

Overall dim ensions

Length 4550 mm

max. nozzle exit diameter 2420 mm

RD-0124


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NPO Energomash

Rocket Engine Manufacturers In Russia

Energia-Buran

RD-170

RD-171

Zenit 1st

stageRD-180

Atlas 3A 1st

stage

RD-191

Angara 1ststage

RD-120

RD-120K

Zenit 2nd

stage

Soyuz-2 1st& 2nd st

Yedinstvo

RD-161

RD-161K

Dvigateli NK

NK-43

NK-33

NK-31

NK-39

Yamal

Air Launch

Kistler

KB Kmimmash

KVD-1

GSLV

KBKhA

RD-0124

RD-0110S5.92

Soyuz 2nd stage

Soyuz-2 3rdstage

Soyuz/Fregat

upper stage

RD-0120

Energia 2nd stage

Soyuz-2

3rdstage

Soyuz-2

3rd

stage

Russian Liquid Propellant Engines

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