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64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

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, zhengxiaoyong2001@163.com
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 LongMarch 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
Long-March 5 (CZ-5) is the next generation of the
Long-March launcher family, under study of the China
Academy of Launch Vehicle Technology (CALT).
Long-March 5 (CZ-5) is the first launcher utilizing
cryogenic and nontoxic propellants (liquid oxygen,
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.

The YF-77 engine is the first booster rocket engine
in China with cryogenic Oxygen/Hydrogen. The YF-77
engine utilizes a gas generator cycle and each engine
has a thrust rating of 700-kN in vacuum at an oxidizerto-fuel mixture ration (O/F) of 5.5. The cryogenic 5-mdiameter main-stage is shown on Fig. 2.

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

CZ-2C CZ-2D CZ-2E CZ-2F CZ-3 CZ-3A CZ-3B CZ-3C
LEO LEO LEO LEO
GTO LEO LEO
LEO
3800 3300 9500 8000
1450 2600 5000
3700

Fig. 1: Heritage of the Long-March family

CZ-5
GTO
14000

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
on pressure, a factor of 9 on mass-flow rate, and a
major increase in scale [1]. The scale of the YF-77
engine comparing with that of the YF-75 engine is
shown on Fig. 3.

* Beijing Aerospace Propulsion Institute, China, zhengxiaoyong2001@163.com
t
Beijing Aerospace Propulsion Institute, China, zhengxiaoyong2001@163.com

IAC-13- C4.1, 2×17679

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64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

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.
Fig. 3: YF-77 (left) and YF-75 (right) engine
The YF-77 engine is based on China’s 40-year
cryogenic engine development legacy and makes use of
the technical experiences acquired through prior
engines. Furthermore, three-dimensional modeling and
a wide array of numerical analysis and design tools are
implemented, which progressing the development
project and shortening the development time. The YF77 provides both high performance and high reliability
to meet the requirements of launcher.

II. ENGINE FEATURE
II.I Engine Main Characteristics
The requirements for low cost, high reliability and
moderate performance from an expendable launcher
had led to the choice of a gas generator cycle for the
YF-77 engine. Two YF-77 engines fly on the first stage
of the Long-March 5 (CZ-5), which joined together by a
flight-type thrust frame. Significant cost reduction and
development progress were achieved by developing two
identically moderate-thrust engines instead of a bigger
one.
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).
At the mid of 2002, the preliminary design of the
engine was accomplished. The first sets of components
had been manufactured and assembled in the first
quarter of 2003. The component and subsystem tests
started in 2003. On July 30, 2003, the gas generator was
successfully tested for the first time and three series of
tests were conducted subsequently. The first power-

IAC-13- C4.1, 2×17679

Figure 5: YF-77 engine schematic diagram
All the subsystems are fixed on the thrust chamber
by mean of supports, and linked together by articled
lines. Thrust vector control for vehicle steering is
achieved by gimbaling the entire engine. Each engine is
gimbaled in two orthogonal planes by two gimbal

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64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

actuators. The YF-77 engine general layout is given in
Fig. 6.

of liquid hydrogen pressurized by the two-stage
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.

Fig. 6: YF-77 engine layout model
The engine main features are:
-LOX/LH2 propellants
-Two engines joined together by flight-type thrust
frame
-Single LOX/LH2 gas generator
-Two separate turbines driven in parallel
-Thrust chamber with coaxial injector, regenerative
cooling and dump-cooled nozzle
-Pyrotechnic starter and igniters
-Pneumatic control system
-Thrust vector control by gimbaling the entire
engine
The major characteristics of the engine are presented
in Table 1.
Item
Nominal Value
Unit
Thrust (vacuum)
2×700
kN
Specific impulse
430
sec.
Mixture ratio
5.5
Chamber pressure
10.2
MPa
Weigh, dry
2700
kg
Expansion area ratio
49
Length
4200
mm
Maximum diameter
5000
mm
Flight burning
520
s
Reliability
0.999
Table 1: Major characteristics of YF-77 engine

Ignition and start system
It initiates combustion and spin-up the turbopumps.
The pyrotechnic igniters units are mounted on the
injector and supplied the initial energy source to ignite
propellants in the combustion chamber and gas
generator respectively. A solid propellant cartridge
provides the initial energy source to spin the propellant
turbopumps during engine start.
Engine pneumatic control system
It regulates the start and shut-down sequence of the
engine. The engine is completely independent and
carries its own helium supply for valve actuation.
Several helium tanks provide a helium pressure supply
to the system, which controls all pneumatically operated
engine valves and provides proper sequencing of engine
components during operation.
Flight instrumentation system
It contains sensors to measure selected engine
parameters for monitoring and evaluating the
operational characteristics of the engine.

II.II Systems
The YF-77 engine was comprised of five
operational systems. A description of each of these
operational systems is as follows.

II.III Major Components
Thrust combustion chamber
The thrust chamber is composed of an injector with
a central solid propellant igniter tube. The coaxial
injector elements include baffle elements which extend
beyond the injector face to prevent high frequency
combustion instability. The main combustion chamber
consists of a LH2 regeneratively cooled inner liner
which made of copper alloy and an outer
electrodeposited nickel shell. The dump-cooled gyroidal
tubular nozzle extension is attached to the main
combustion chamber at an area ratio of 5:1. The nozzle
area ratio of 49:1 is selected considering not separate at
a sea level operation. The thrust chamber and nozzle are
shown in Fig. 7.

Propellant feed system
It supplies pressurized propellants for thrust
chamber and gas generator. Liquid oxygen from the
oxidizer tank is pressurized by the oxidizer turbopump,
and approximately 3.4% of liquid oxygen flow is split
and supplied to the gas generator. Approximately 84%

Item
Nominal Value
Chamber pressure
10.2
Mixture ratio
6.4
Nozzle area ratio
49
Dump-cooling flowrate
5%
Table 2: Thrust chamber characteristics

IAC-13- C4.1, 2×17679

Unit
MPa

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64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

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.
The fuel turbopump is show on Fig. 9.

Fig. 7: Thrust chamber configuration
Gas Generator
The gas generator consists of a non-cooled
combustion chamber and an injector assembly. The
nominal mixture ratio of the gas generator is 0.9. The
high energy gases produced by the generator are
directed to the fuel and oxidizer turbine respectively
before being exhausted. The solid propellant cartridge is
mounted vertically at the gas generator exhaust. The gas
generator is show on Fig. 8.
Item
Nominal Value
Combustion pressure
8.5
Mixture ratio
0.9
Gas temperature
900
Table 3: Gas generator characteristics

Item
Nominal Value
Pump discharge Pressure 16.5
Pump efficiency
0.75
Bearings D×N
2.1×106
Shaft speed
35000
Turbine pressure ratio
15.5
Turbine efficiency
0.52
Table 4: Fuel turbopump characteristics

Unit
MPa
mm×rpm
rpm

Unit
MPa
K

Fig. 9: Fuel turbopump configuration

Fig. 8: Gas generator 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
purged by helium supply during turbopump operation to
separate the cryogenic bearing coolant in pump and the
fuel-rich hot gas in the turbine section. Five dynamic
seals in series located between the turbine section and
the pump section prevent the coolant from leaking into
the turbine. The oxygen turbopump works in the region
of 1st and 2nd critical rotate speed. The oxygen
turbopump is show on Fig. 10.

Turbopumps
The fuel turbopump is composed of a two-stage
centrifugal pump with an inducer and a two-stage
impulse turbine. The two fully shrouded impellers have
identical flow passages and made of Titanium alloy
with powder metallurgy process. The inner and outer
pump housings are manufactured in Titanium alloy
Item
Nominal Value
casting.
Pump discharge pressure 14
The turbine is a supersonic axial turbine which
Pump efficiency
0.74
consists of two shrouded blisk (blades and shroud
Shaft speed
18000
integrated to the disk).The rotor blades are machined
Turbine pressure ratio
14
from a monolithic disk forging using electro discharge
Turbine efficiency
0.35
machining (EDM) processes. Three honeycomb seals
are used to providing closer operating clearance Table 5: Oxygen turbopump characteristics
between the turbine blade tip seal lands and the

IAC-13- C4.1, 2×17679

Unit
MPa
rpm

Page 4 of 7

64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

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

Fig. 10: Oxygen turbopump configuration
Valves
All valves are pneumatically actuated by helium
from the bottle. The main fuel and oxidizer valves are
ball-type valves located in the propellant high pressure
duct between the turbopumps and the combustion
chamber. The oxidizer valve is a two-stage valve. The
first-stage actuator positions the main oxidizer valve at
the 10-deg position to obtain initial thrust chamber
ignition; the second-stage actuator ramps the main
oxidizer valve full open to accelerate the engine to
main-stage operation. The propellant prevalves are also
ball-type valves located in the low pressure ducts
interfacing the stage and the engine, retaining propellant
in the stage until being admitted into the engine.

Fig. 13: Propellant utilization valve configuration
III. ENGINE DEVELOPMENT PROGRAM
YF-77 components, subsystems, as well as the
entire engine had conducted various rigorous testing to
verify performance and reliability. To progress the
development project and shorten the development time,
each of the developmental phases were overlapped. The
development history of YF-77 engine has demonstrated
the robustness and reliability of this engine.
A summary of the YF-77 engine testing can be
found in Table 6.
30000
Engine-vehicle
static hot-fire test

Total Time, sec.

25000
20000

Engine dev. ended
Cert. started

15000
Prototype engine
50-seconds firing

10000

First flightduration test

5000

Fig. 11: Main valve configuration
The pneumatically operated gas-generator valves
and bleed valves control supplied pressurized
propellants for gas generator, and provide pressure
relief for the boiloff of propellants trapped interior at
engine shutdown. An auxiliary function of the bleed
valves is to provide propellants bleed and chilldown
circuit through the pump.

Entire engine
gimbaling test

0
1

10

100

Accumulative Number of Tests

Fig. 14: Total time Vs. accumulative number of tests
III.I Sub-scale thrust chamber tests
The first sub-scale thrust chamber test was made on
June 6, 2003. A total of 8 hot-fire tests were conducted
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

IAC-13- C4.1, 2×17679

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64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

Fig. 15: Sub-scale thrust chamber testing
III.II Gas generator tests
The first gas generator was assembled and tested on
July 30, 2003 and extended over the 2003~2006 period.
Total 9 tests had been carried out to demonstrate
compliance with the specifications. The operational
domain of the gas generator had been explored over a
range from 2.4 to 8.6 MPa, mixture ratio from 0.63 to
1.08. Special test was made without pyrotechnic
igniters to check the ignition by the solid propellant
cartridge. The performance, reliability and stability of
gas generator had been verified.

Fig. 17: Workhorse gas generator testing
III.IV Prototype engine tests
At the end of the component and subsystem
development phase, the first prototype engine tests with
a short nozzle began in June 2004. To mitigate the risks,
the first hot test with ignition of the chamber was made
on June 10, 2004 to knowledge of the turbopumps and
thrust chamber behavior during chill-down and ignition.
After that, a short duration (10 s.) test of steady state
condition was achieved on June, 18. Three months later
on September 17, 2004, successful 50-seconds firing of
prototype engine was achieved. These tests helped to
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%

Fig. 16: Gas generator testing

Chamber Pressure

100%

C7701
Short duration

80%

C7702
50-seconds

60%
40%

C7701-L
Chamber Ignition

20%

III.III Powerpack tests
Before the subsystem level test, bearings, dynamic
seals, inducers, pumps and other key components of the
turbopumps were extensively tested at special facilities.
Full-scale powerpack test was conducted to validate
turbopumps and gas generator behavior near the
nominal point. Two oxygen turbopump (OTP) with GG
tests and one hydrogen turbopump (FTP) with GG test
were carried out in December, 2003, March, 2004 and
January 2004 respectively. The first OTP with GG test
and FTP with GG test were made with liquid nitrogen
in pump. Testing results verified proper behaviors of
the turbopump during transition to steady-state
operation.

IAC-13- C4.1, 2×17679

0%

Duration Seconds

Fig. 18: Prototype engine transient test phase
III.V Developmental and limited engine tests
After the prototype engine test, YF-77 engine began
extensively tests to verify the reliability and
performance. Engine durability, performance, thrust,
and operational limits was thoroughly demonstrated in
developmental and limited engine testing. Most of
problems had been encountered during this phase, and
been analyzed, understood, and corrected.
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

Page 6 of 7

64rd International Astronautical Congress, Beijing, China. Copyright ©2013 by the International Astronautical Federation. All rights reserved.

seconds and 15 mission cycles. The endurance, lifetime
and reliability of the YF-77 engine were well validated.

remarkably. A total of 4 aggressive hot-fire tests for
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
verified a good behavior in a wide range.
7

Engine Mixture Ratio

Envelops and limits testing of the YF-77 engine
verified the good behavior and reliability of the Engine
in a wide range, covering the operation domain taking
into account sufficient margin for ambient and
environmental scatter.

Design Point
Test Points

6.5

Fight Envelope
6

5.5
5

4.5
4
9

9.5

Fig. 21: Captive firing test lab.

10

10.5

11

Combustion Chamber Pressure, MPa

11.5

Fig. 20: Domain tested during development
Identified Limits Verified
The limits of YF-77 engine were checked and
indentified during development, such as chill-down
temperature of bearing before spin-up, pump critical
NPSH, start & shut down sequences, and etc. Major
component operation appeared satisfactory, exhibiting
wide operation margin and no evidence of failure or
damage were encountered on any of these tests.
As an instance, special cavitation hot-fire tests with
entire engine were made to characterize the cavitation
behavior of the inducer and impeller. Pump inlet
pressure was decreased gradually during engine steadystate till the pumps reached a critical cavitation point.
The firing was terminated by the engine safety cutoff
system before the performance of the pump deteriorated

IAC-13- C4.1, 2×17679

IV. CONCLUSION
During the past 40 years, the cryogenic engines of
China have developed from the YF-73 (44kN) to YF-77
(700kN). The YF-77 engine makes use of existing state
of the art and provides both high performance and high
reliability to meet mission requirements of vehicle.
Today, 1.5 years before the maiden flight, the
development of the YF-77 engine remains perfectly
inside the schedule and the objectives assigned.
Being the first large cryogenic LOX/LH2 engine in
China, the YF-77 program is a key element of China
access to space in future, and it gives the potential to
perform a broad array of missions. The YF-77 engine
not only contributes to the high launch capability of the
CZ-5, but also opens a door to future more powerful
engines.
REFERENCES
[1] Gu Mingchu, 2000, “Speed up the Development of
LOX/LH2 Rocket Engine to Greet the 21st Century”.
MISSILES AND SPACE VEHICLES, No. 1, 2000.

Page 7 of 7


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