Iac-13, C4.1, 2×17679

IAC-13, C4.1, 2×17679

DEVELOPMENT STATUS OF THE CRYOGENIC OXYGEN/HYDROGEN YF-77 ENGINE FOR LONG-MARCH 5

Weibin Wang Beijing Aerospace Propulsion Institute, Beijing, China, [email protected]

Dayong Zheng*, Guiyu Qiaot

The YF-77 engine, developed by the Academy of Aerospace Launch Propulsion Technology (AALPT), China, is a high performance and reliability booster designed for Chinese next-flagship expendable launcher, called Long- March 5 (CZ-5). The YF-77 engine is the first high-thrust cryogenic engine developed in China, which takes a big technological step with respect to previous Chinese cryogenic Oxygen/Hydrogen engine. The engine utilizes gas generator cycle with cryogenic LOX/LH2 propellants. Two YF-77 engines fly on the first stage of the Long-March 5 (CZ-5), and each engine provids 700-kN in vacuum at an oxidizer-to-fuel mixture ration (O/F) of 5.5. This discussion covers engine system and component characteristics as well as the development status of YF-77 engine. The reliability and safety of YF-77 is well demonstrated in engine testing during development before its maiden journey.

I. INTRODUCTION The YF-77 engine is the first booster rocket engine Long-March 5 (CZ-5) is the next generation of the in China with cryogenic Oxygen/Hydrogen. The YF-77 Long-March launcher family, under study of the China engine utilizes a gas generator cycle and each engine Academy of Launch Vehicle Technology (CALT). has a thrust rating of 700-kN in vacuum at an oxidizer- Long-March 5 (CZ-5) is the first launcher utilizing to-fuel mixture ration (O/F) of 5.5. The cryogenic 5-m- cryogenic and nontoxic propellants (liquid oxygen, diameter main-stage is shown on Fig. 2. LOX/liquid hydrogen, LH2, and Kerosene) in China, which is entirely clean and environmentally friendly. Long-March 5 (CZ-5) is powered by four Oxygen/Kerosene boosters, two YF-77 LOX/LH2 engines on the core stage, and two LOX/LH2 expender cycle engines on the second stage. Compare to the former Long-March launcher, CZ-5 has significantly more lift capability which can deliver a payload of 14,000 kg to Geosynchronous Transfer Orbit (GTO) and 25,000 kg to Low Earth Orbit (LEO). The evolution of the Long-March family is shown on Fig. 1.

Fig. 2: Cryogenic 5-m-diameter main-stage

The YF-77 engine is the first high-thrust cryogenic engine developed in China, which presented a big challenge. It takes a big technological step with respect to previous Chinese cryogenic Oxygen/Hydrogen engine, such as YF-75 which powers CZ-3A/3B’s upper stage, with a factor of 9 on thrust, a factor of 2.7 CZ-2C CZ-2D CZ-2E CZ-2F CZ-3 CZ-3A CZ-3B CZ-3C CZ-5 LEO LEO LEO LEO GTO LEO LEO LEO GTO on pressure, a factor of 9 on mass-flow rate, and a [1] 3800 3300 9500 8000 1450 2600 5000 3700 14000 major increase in scale . The scale of the YF-77 Fig. 1: Heritage of the Long-March family engine comparing with that of the YF-75 engine is shown on Fig. 3.

* Beijing Aerospace Propulsion Institute, China, [email protected] t Beijing Aerospace Propulsion Institute, China, [email protected]

pack test was made in December, 2003. Nine months later on September 17, a successful 50-seconds firing of prototype engine was achieved. In May 2013, the formal qualification test series began. At the end of September 2013, more than 70 tests and 24,000 seconds of steady state conditions have been accumulated with 12 engines. Today, a concept review of the YF-77 engine confirmed the performance goal and the need for launcher, which is intended to be ready of its maiden flight at mid 2015. Significant milestones have been reached.

II. ENGINE FEATURE Fig. 3: YF-77 (left) and YF-75 (right) engine II.I Engine Main Characteristics The requirements for low cost, high reliability and The YF-77 engine is based on China’s 40-year moderate performance from an expendable launcher cryogenic engine development legacy and makes use of had led to the choice of a gas generator cycle for the the technical experiences acquired through prior YF-77 engine. Two YF-77 engines fly on the first stage engines. Furthermore, three-dimensional modeling and of the Long-March 5 (CZ-5), which joined together by a a wide array of numerical analysis and design tools are flight-type thrust frame. Significant cost reduction and implemented, which progressing the development development progress were achieved by developing two project and shortening the development time. The YF- identically moderate-thrust engines instead of a bigger 77 provides both high performance and high reliability one. to meet the requirements of launcher. Its thrust chamber is fed by separate turbopumps with turbines in parallel and separate gas exhausts. The combustion chamber is regeneratively cooled and the nozzle is dump-cooled. The gas generator and the combustion chamber are ignited by pyrotechnic igniters and the turbopumps are started by a solid propellant cartridge. The pre-valves and main valves are helium actuated ball valves. The engine thrust and mixture ratio are calibrated with venturis and propellant utilization valve during engine tests on ground. The YF-77 also supplies gaseous hydrogen and oxygen for tanks pressurization by heat exchanger. The YF-77 engine schematic diagram is shown on Fig. 5.

Fig. 4: Virtual design and analysis

In January 2002, The Commission of Science, Technology and Industry for National Defense (COSTIND) approved the development of a new cryogenic engine—the YF-77, which was the most powerful cryogenic LOX/LH2 engine ever developed in China. The engine development program is under responsibility of Beijing Aerospace Propulsion Institute (BAPI), a division of the Academy of Aerospace Launch Propulsion Technology (AALPT). Figure 5: YF-77 engine schematic diagram At the mid of 2002, the preliminary design of the engine was accomplished. The first sets of components All the subsystems are fixed on the thrust chamber had been manufactured and assembled in the first by mean of supports, and linked together by articled quarter of 2003. The component and subsystem tests lines. Thrust vector control for vehicle steering is started in 2003. On July 30, 2003, the gas generator was achieved by gimbaling the entire engine. Each engine is successfully tested for the first time and three series of gimbaled in two orthogonal planes by two gimbal tests were conducted subsequently. The first power-

actuators. The YF-77 engine general layout is given in of liquid hydrogen pressurized by the two-stage Fig. 6. hydrogen pump cools the thrust chamber, and almost all of the heated hydrogen in the cooling path is injected into the combustion chamber.

Gas generator system It produces turbine drive gases for the turbopumps. The fuel-rich gas from the gas generator is divided into 70% and 30% to the fuel turbine and the oxidizer turbine respectively before being exhausted.

Ignition and start system It initiates combustion and spin-up the turbopumps. The pyrotechnic igniters units are mounted on the Fig. 6: YF-77 engine layout model injector and supplied the initial energy source to ignite propellants in the combustion chamber and gas generator respectively. A solid propellant cartridge The engine main features are: provides the initial energy source to spin the propellant -LOX/LH2 propellants turbopumps during engine start. -Two engines joined together by flight-type thrust frame Engine pneumatic control system -Single LOX/LH2 gas generator It regulates the start and shut-down sequence of the -Two separate turbines driven in parallel engine. The engine is completely independent and -Thrust chamber with coaxial injector, regenerative carries its own helium supply for valve actuation. cooling and dump-cooled nozzle Several helium tanks provide a helium pressure supply -Pyrotechnic starter and igniters to the system, which controls all pneumatically operated -Pneumatic control system engine valves and provides proper sequencing of engine -Thrust vector control by gimbaling the entire components during operation. engine The major characteristics of the engine are presented Flight instrumentation system in Table 1. It contains sensors to measure selected engine parameters for monitoring and evaluating the Item Nominal Value Unit operational characteristics of the engine. Thrust (vacuum) 2×700 kN Specific impulse 430 sec. II.III Major Components Mixture ratio 5.5 Thrust combustion chamber Chamber pressure 10.2 MPa The thrust chamber is composed of an injector with Weigh, dry 2700 kg a central solid propellant igniter tube. The coaxial Expansion area ratio 49 injector elements include baffle elements which extend Length 4200 mm beyond the injector face to prevent high frequency Maximum diameter 5000 mm combustion instability. The main combustion chamber Flight burning 520 s consists of a LH2 regeneratively cooled inner liner Reliability 0.999 which made of copper alloy and an outer Table 1: Major characteristics of YF-77 engine electrodeposited nickel shell. The dump-cooled gyroidal tubular nozzle extension is attached to the main II.II Systems combustion chamber at an area ratio of 5:1. The nozzle The YF-77 engine was comprised of five area ratio of 49:1 is selected considering not separate at operational systems. A description of each of these a sea level operation. The thrust chamber and nozzle are operational systems is as follows. shown in Fig. 7.

Propellant feed system Item Nominal Value Unit It supplies pressurized propellants for thrust Chamber pressure 10.2 MPa chamber and gas generator. Liquid oxygen from the Mixture ratio 6.4 oxidizer tank is pressurized by the oxidizer turbopump, Nozzle area ratio 49 and approximately 3.4% of liquid oxygen flow is split Dump-cooling flowrate 5% and supplied to the gas generator. Approximately 84% Table 2: Thrust chamber characteristics

stationary honeycomb seal face. The choice of material for the most of turbine parts is In 718. The rotating assembly is supported by duplex angular contact ball bearings, which are mounted in flexible damper carriers in order to allow adaptation of the stiffness and to limit the effects of shaft vibration. The bearings, equipped with ceramic balls, are cooled by hydrogen and lubricated by their own retainer material. The fuel turbopump is operating between the second and third critical speed. The maximum DN value is over 2.1×106 mm×rpm. The fuel turbopump works in the region of 2nd and 3rd critical rotate speed. Fig. 7: Thrust chamber configuration The fuel turbopump is show on Fig. 9.

Item Nominal Value Unit Gas Generator Pump discharge Pressure 16.5 MPa The gas generator consists of a non-cooled Pump efficiency 0.75 combustion chamber and an injector assembly. The Bearings D×N 2.1×106 mm×rpm nominal mixture ratio of the gas generator is 0.9. The Shaft speed 35000 rpm high energy gases produced by the generator are Turbine pressure ratio 15.5 directed to the fuel and oxidizer turbine respectively Turbine efficiency 0.52 before being exhausted. The solid propellant cartridge is Table 4: Fuel turbopump characteristics mounted vertically at the gas generator exhaust. The gas generator is show on Fig. 8.

Item Nominal Value Unit Combustion pressure 8.5 MPa Mixture ratio 0.9 Gas temperature 900 K Table 3: Gas generator characteristics

Fig. 9: Fuel turbopump configuration

The oxygen turbopump consists of a single-stage centrifugal pump with a helical inducer driven by a two-stage turbine. The turbine side bearings are cooled and lubricated by oxygen. A cavity is continuously Fig. 8: Gas generator configuration purged by helium supply during turbopump operation to separate the cryogenic bearing coolant in pump and the Turbopumps fuel-rich hot gas in the turbine section. Five dynamic The fuel turbopump is composed of a two-stage seals in series located between the turbine section and centrifugal pump with an inducer and a two-stage the pump section prevent the coolant from leaking into impulse turbine. The two fully shrouded impellers have the turbine. The oxygen turbopump works in the region st nd identical flow passages and made of Titanium alloy of 1 and 2 critical rotate speed. The oxygen with powder metallurgy process. The inner and outer turbopump is show on Fig. 10. pump housings are manufactured in Titanium alloy casting. Item Nominal Value Unit The turbine is a supersonic axial turbine which Pump discharge pressure 14 MPa consists of two shrouded blisk (blades and shroud Pump efficiency 0.74 integrated to the disk).The rotor blades are machined Shaft speed 18000 rpm from a monolithic disk forging using electro discharge Turbine pressure ratio 14 machining (EDM) processes. Three honeycomb seals Turbine efficiency 0.35 are used to providing closer operating clearance Table 5: Oxygen turbopump characteristics between the turbine blade tip seal lands and the

valve gate is equipped with nozzle, the size of which is determined during engine calibrating firing test.

Fig. 10: Oxygen turbopump configuration

Valves Fig. 13: Propellant utilization valve configuration All valves are pneumatically actuated by helium from the bottle. The main fuel and oxidizer valves are III. ENGINE DEVELOPMENT PROGRAM ball-type valves located in the propellant high pressure YF-77 components, subsystems, as well as the duct between the turbopumps and the combustion entire engine had conducted various rigorous testing to chamber. The oxidizer valve is a two-stage valve. The verify performance and reliability. To progress the first-stage actuator positions the main oxidizer valve at development project and shorten the development time, the 10-deg position to obtain initial thrust chamber each of the developmental phases were overlapped. The ignition; the second-stage actuator ramps the main development history of YF-77 engine has demonstrated oxidizer valve full open to accelerate the engine to the robustness and reliability of this engine. main-stage operation. The propellant prevalves are also A summary of the YF-77 engine testing can be ball-type valves located in the low pressure ducts found in Table 6. interfacing the stage and the engine, retaining propellant in the stage until being admitted into the engine. 30000 Engine-vehicle 25000 static hot-fire test

20000 Engine dev. ended Cert. started 15000 Prototype engine First flight- 10000 50-seconds firing duration test Total Time, sec.

5000 Entire engine gimbaling test Fig. 11: Main valve configuration 0 1 10 100 Accumulative Number of Tests The pneumatically operated gas-generator valves and bleed valves control supplied pressurized Fig. 14: Total time Vs. accumulative number of tests propellants for gas generator, and provide pressure relief for the boiloff of propellants trapped interior at III.I Sub-scale thrust chamber tests engine shutdown. An auxiliary function of the bleed The first sub-scale thrust chamber test was made on valves is to provide propellants bleed and chilldown June 6, 2003. A total of 8 hot-fire tests were conducted circuit through the pump. on this program to evaluate the combustion performance and instability. The cryogenic gaseous hydrogen was delivered to the sub-scale injectors through a mixer, which was designed to mix the cryogenic liquid hydrogen and room temperature gaseous hydrogen to achieve appropriate temperature. Recalculated cooling water was used to cool the combustion chamber. Be enslaved to qualification of the facility, the full-scale thrust chamber test was canceled.

Fig. 12: Gas generator valve configuration

The propellant utilization valve is mounted on the oxidizer turbine inlet to vary engine mixture ration. The

Fig. 17: Workhorse gas generator testing Fig. 15: Sub-scale thrust chamber testing III.IV Prototype engine tests III.II Gas generator tests At the end of the component and subsystem The first gas generator was assembled and tested on development phase, the first prototype engine tests with July 30, 2003 and extended over the 2003~2006 period. a short nozzle began in June 2004. To mitigate the risks, Total 9 tests had been carried out to demonstrate the first hot test with ignition of the chamber was made compliance with the specifications. The operational on June 10, 2004 to knowledge of the turbopumps and domain of the gas generator had been explored over a thrust chamber behavior during chill-down and ignition. range from 2.4 to 8.6 MPa, mixture ratio from 0.63 to After that, a short duration (10 s.) test of steady state 1.08. Special test was made without pyrotechnic condition was achieved on June, 18. Three months later igniters to check the ignition by the solid propellant on September 17, 2004, successful 50-seconds firing of cartridge. The performance, reliability and stability of prototype engine was achieved. These tests helped to gas generator had been verified. verify compatibility of each component, and confirm the start-up and shut-down sequence. The prototype engine tests had accumulated experience and prepared for the engine flight-duration test.

120% C7701 Short duration 100%

80% C7702 50-seconds 60%

40% C7701-L Fig. 16: Gas generator testing Chamber Ignition Chamber Pressure Chamber 20% III.III Powerpack tests 0% Before the subsystem level test, bearings, dynamic Duration Seconds seals, inducers, pumps and other key components of the turbopumps were extensively tested at special facilities. Fig. 18: Prototype engine transient test phase Full-scale powerpack test was conducted to validate turbopumps and gas generator behavior near the III.V Developmental and limited engine tests nominal point. Two oxygen turbopump (OTP) with GG After the prototype engine test, YF-77 engine began tests and one hydrogen turbopump (FTP) with GG test extensively tests to verify the reliability and were carried out in December, 2003, March, 2004 and performance. Engine durability, performance, thrust, January 2004 respectively. The first OTP with GG test and operational limits was thoroughly demonstrated in and FTP with GG test were made with liquid nitrogen developmental and limited engine testing. Most of in pump. Testing results verified proper behaviors of problems had been encountered during this phase, and the turbopump during transition to steady-state been analyzed, understood, and corrected. operation. Equivalent Mission Validated Numbers of cycles and cumulated tests duration are key factors for reliability estimation. The YF-77 engine is a single-burning engine, and its design life is 520 seconds at one mission duration. In the development program, one of the engines demonstrated over 5300

seconds and 15 mission cycles. The endurance, lifetime remarkably. A total of 4 aggressive hot-fire tests for and reliability of the YF-77 engine were well validated. cavitation were achieved without anomaly, and the disassembly of the FTP and OTP showed a very good behavior of all the parts in turbopumps. The cavitation performance and cavitation specifications of the pump were verified. To verify compatibility of the 1st Booster Core with two YF-77 engines, static hot fire test series will be conducted on a new launch pad in the near future. During the series, engine-vehicle communication, operational sequences will be demonstrated and verified. Fig. 19: Development and limits testing

Envelops and Domain Verified The thrust and mixture ratio of YF-77 engine are drifting according to the evolution of the inlet pressure and temperature during flight. The deviations and scatters of modeling processes, component performances and manufacturing tolerances to the flight domain have also to be taken into account at mission duration. The YF-77’s performance, security and reliability were demonstrated widely at nominal and off-design conditions. The domain tests of YF-77 engine during development is show in Figure 20, which Fig. 21: Captive firing test lab. verified a good behavior in a wide range. Envelops and limits testing of the YF-77 engine 7 Design Point verified the good behavior and reliability of the Engine 6.5 Test Points in a wide range, covering the operation domain taking o Fight Envelope into account sufficient margin for ambient and 6 environmental scatter.

5.5 IV. CONCLUSION 5 During the past 40 years, the cryogenic engines of

Engine Mixture Rati Mixture Engine 4.5 China have developed from the YF-73 (44kN) to YF-77 (700kN). The YF-77 engine makes use of existing state 4 of the art and provides both high performance and high 9 9.5 10 10.5 11 11.5 reliability to meet mission requirements of vehicle. Combustion Chamber Pressure, MPa Today, 1.5 years before the maiden flight, the Fig. 20: Domain tested during development development of the YF-77 engine remains perfectly inside the schedule and the objectives assigned. Identified Limits Verified Being the first large cryogenic LOX/LH2 engine in The limits of YF-77 engine were checked and China, the YF-77 program is a key element of China indentified during development, such as chill-down access to space in future, and it gives the potential to temperature of bearing before spin-up, pump critical perform a broad array of missions. The YF-77 engine NPSH, start & shut down sequences, and etc. Major not only contributes to the high launch capability of the component operation appeared satisfactory, exhibiting CZ-5, but also opens a door to future more powerful wide operation margin and no evidence of failure or engines. damage were encountered on any of these tests. As an instance, special cavitation hot-fire tests with REFERENCES entire engine were made to characterize the cavitation behavior of the inducer and impeller. Pump inlet [1] Gu Mingchu, 2000, “Speed up the Development of pressure was decreased gradually during engine steady- LOX/LH2 Rocket Engine to Greet the 21st Century”. state till the pumps reached a critical cavitation point. MISSILES AND SPACE VEHICLES, No. 1, 2000. The firing was terminated by the engine safety cutoff system before the performance of the pump deteriorated

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