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Enhanced Reliability Features of the RL10E 1 .pdf

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Title: PII: S0094-5765(98)00077-0

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C 1998 International

Acre Asrronautica Vol. 41, Nos 4 IO. pp. 197-207. 1997
Published by Elsewcr Scirnce Ltd

PII: SOO94-5765(98)00077-O

Printed tn Great Britain
Sl9.00 + 0.00

W. M. Van Lerberghe’, J. L. Emdee’, and R. R. Faust’
The Aerospace Corporation
El Segundo, CA


rocket engine has been partially

developed for the United States Air Force during the
Atlas Reliability Enhancement Program. This engine is
a 22,300 lbf thrust liquid hydrogen, liquid oxygen,
derivative of the RLlOA4- 1, and incorporates an
improved ignition system and new electromechanically
actuated (EMA)
valves controlled by an engine
mounted Digital Electronic Rocket Engine Controller.

15,000 lbf for the RLIOA-3
to 22,300 Ibf for the
The RLlOE-1 engine (Figure 1) is one tf
the latest RLlO derivatives and was partially developed
during the Atlas Reliability Enhancement Program
(AREP) for the United States Air Force (USAF).

electronic control system operates
the engine with only six valves (compared to the twelve
on the




boost phase engine

cooldown capability (to avoid cryogenic temperatures at
liftoff), and allows implementation of an improved torch
ignition system with redundant, modem electronics.
Other potential reliability improvements include active
engine health monitoring, derated thrust operation,
“sot?” transients, and improved minimum residual
shutdown capability.
Although development of the RLIOE-I

was not

completed, the engine development tests were highly
successful in demonstrating engine operation with the
full authority. digital electronic control system and
fewer valves. Many of the enhanced reliability features
have been demonstrated and are directly applicable to
other versions of the RL IO engines. Potential “hybrid”
engine conftgurations offer a lower weight and less
costly engine. while retaining much of the reliability
improvements and other benefits of the RL 1OE- I.

0 1998 International Astronautical Federation.
Published by Elsevier Science Ltd

RL IO rocket engines, manufactured by Pratt &
Whitney (P&W),
have been flying since the early
1960’s. These engines bum liquid hydrogen and liquid
oxygen (LOX). Engine thrust has been uprated from

aSenior Member of the Technical Staff.
’ Section Manager. Propulsion Dept.
’ RLIO Cons&ant

Propulsion Dcpt

to The Aerospace Corporation


Figure I: RLIOE-1 Engine (XR602).

The AREP program began in 1994 as a vehicle
reliability upgrade program for the Lockheed Martin
Astronautics (LMA)
Atlas Centaur vehicle’. The
program goal was to decrease the engine failure rate by a
relative 71% of the nominal failure rate, without
decreasing vehicle performance, such that vehicle would
experience a delta reliability increase of one and a half
percentage points (A = +1.5%)
from the baseline
demonstrated vehicle reliability of 92.7%. Major amas
selected for engine improvement included the ignition
system, the control system, and the chilldown system.


48th IAF Congress

The program was managed using the Integrated
Product Team (IPT) approach. Component Integrated
Product Teams (CIPT’s) were assigned one or more
areas for upgrade. The CIPT’s reported to a higher level
System Engineering Integration Team and a Program
Management Team. Team members included the
engine manufacturer (P&W), the vehicle manufacturer
(LMA), the USAF, The Aerospace Corporation, and
the National Aeronautics and Space Administration
The AREP program was terminated in April 1997
for the convenience of the USAF. An RLIOE-1
prototype test engine and two development engines
were tested prior to termination, while component
development testing was nearing completion. Many cf
Other potential enhanced reliability
features are available but have not yet been
demonstrated. Major reliability enhancing features of the
fU I OE- 1 will be discussed.
The RLIO engine family consists of multi-start
engines capable of mixture ratio (MR) control to
maximize vehicle propellant usage. These engines use
an expander cycle, in which heat absorbed tiom the
thrust chamber cooling jacket vaporizes hydrogen and
powers the turbine that drives the propellant pumps.

Figure 2: RLlOA41

The RLlOA-3-3A engine is currently flown on
Titan Centaur (TIC). It has a nominal vacuum thrust
and specific impulse (1s~) at MR=5.0 of 16,500 Ibf and
444 set, respectively. The higher thrust RLlOA-4 and
even higher thrust RLlOA4I
engines are currently
flown on Atlas Centaur (A/C). Both these engines use a
new LOX pump and turbine. The RL IOA4 I also uses
a modified main injector. An extendible, columbium
nozzle extension with increased expansion ratio can be
used to improve thrust (+300 Ibf) and Isp. With the
nozzle extension at MR=5.5, the RL10A-4 can achieve
a vacuum performance (thrust and Isp) of 20,800 Ibf and
449 set, while the RLlOA41 can achieve 22,300 lbf
and 451 sec. The RLlOA41 is the current production
engine. Each Centaur uses two RLlO engines.
In-flight cooldown, start, and shutdown for all
these engines are activated by three solenoid valves
which provide pneumatic actuation pressure to the five
primary propellant control valves (see Figure 2, for the
RL 1OA4 1). In all, twelve valves and solenoids enable
ground pre-chill, in-flight prestart cooldown, ignition,
acceleration to steady state, chamber pressure control,
mixture ratio control, and engine shutdown. The
weight of the RLlOA4I
engine, not including the
vehicle supplied propellant utilization motor and the
nozzle extension, is 320 Ibs. The nozzle extension,
when used, adds 55 Ibs to engine weight.


Flow Schematic.

48th IAF Congress

Figure e 3: RLIOE-I


Engine modifications
which would provide the
greatest reliability
and performance improvements with
the lowest risk and cost were selected based upon
detailed trade studies using analytical models, previous
program experience. and industry data on valves and
valve actuators. Reliability
improvement, with a 40%
trade weighting factor, was the most important factor in
the indrvtdual trades. The baseline configuration was
established at the conclusion of the trade studies in
September 199-I.
like the RLIOA-4-I.
is rated at
Ibf full thrust operation with a corresponding
6 IO psia chamber pressure. A change from the
nominally trimmed 5.5 oxidizer-to-fuel
mixture ratio for
to a 5.35 MR for the RLIOE-I
being developed to match vehicle propellant utilization
histoty. This modified trim condition provides slightly
higher thrust compared to the RLIOA-4-I
at a given
MR. The RL IOE-I engine has demonstrated throttle
capabili? down to 47% thrust. The minimum thrust
level is lrmited by the turbine bypass effective flow area
afforded by the mated gearbox housing and thrust
control valve combination. The RLIOE-I
includes the
same basic turbomachinery,
injector, thrust chamber
and nozzle extension as the RLIOA-4-I.
The number d
valves and solenoids has been reduced from twelve on
to six valves on the RLIOE-I
3). Engine weight without the nozzle extension is 384
Ibs, a 64 Ibs increase with respect to the RLIOA-4-1.

Flow Schematic.

Nevertheless, the delivered payload capability is not
projected to decrease, due to reduced propellant
consumption for engine chilldown and elimination d
the vehicle
pneumatic plumbing, and prechill system.
All of the valves are electromechanically actuated
(EMA) and controlled by a Digital Electronic Rocket
Engine Controller
valve control
eliminates the three pneumatic solenoids. Modified
engine operation enables elimination of one of the two
hydrogen cooldown valves. All six EMA actuators are
identical in design and are electrically redundant with
brushless. direct current motors. Each EMA
includes a position feedback system. The EMAs can
move the valves at a maximum slew rate of 360” per
second (whereas the valves open and close within 900).
is engine mounted
radiation hardened processors, redundant cables to the
EMAs and built-in fault detection and accommodation
logic to further enhance reliability features.
Every propellant control valve was newly designed
for the RLI OE- I. The oxygen and fuel inlet valves (OIV
and FIV) have similar
ball valve designs. The
cooldown valve (CDV) and oxidizer control valve
(OCV) are identical and have a partial ball, or visor,
valve design. The design of the fuel shutoff valve (FSV)
is similar to that of the OCV and CDV. The thrust
control valve (TCV) has a rotary sleeve valve which is



48th IAF
greatly simplified compared to the mechanical fMback
system of the RL 1OA4 I thrust control valve.
An improved, dual spark plug, continuous burning
torch ignition system (Figure 4) with a redundant,
modem electronics exciter box permits elimination d
the igniter oxidizer supply valve (recall Figure 2). The
exciter provides 0.2 joules to each spark plug at a rate
of 40- 100 sparks per second. The torch igniter includes
a single shear coaxial oxygen-hydrogen element
injector. The torch combustion chamber is cooled by


Figure 4: Torch igniter.

The RL IOE-I control system enables ambient
temperature litlofT and boost phase cooldown. This
eliminates the need for the launch site ground chill
system and the engine prelaunch cooldown valve, either
of which potentially can contaminate the engine. Boost
phase cooldown increases reliability and also conserves
oxidizer compared to current RLIOA4I methods.

For flight operation, the engine lifts otTwith pump
near 450”R. Oxygen side chilldown is
initiated shonly after liftoff by opening the OIV to allow
liquid oxygen to fill the pump to the DCV under
pressure and gravity forces. This cooling phase is
termed percolation chilldown, since oxygen vapor
bubbles evolve from the engine and rise up through the
oxygen feedline to the tank where the gas is vented.
Percolation chilldown lasts about 280 set during the
ascent phase. Fuel side chilldown is initiated after the
vehicle reaches an altitude of 150.000 ft. The FIV is
fully opened and theCDV is opened to 1.15 sq. in. to
allow hydrogen to enter the engine, flow through the
CDV, and then be expelled out an overboard vent
system. Hydrogen Row is stopped temporarily I5
seconds prior to A/C separation to prevent the presence

of hydrogen gas during the pyrotechnic sepamtion
event. Shortly after separation, the OCV is opened to
0.35 sq. in., the FIV is fully reopened, and the CDV is
reopened to 1.15 sq. in. This 8 second prestart process
flushes any saturated propellant out of the feedlines in
preparation for the first main-engine-start (MES I).
Second burn chilldown is accomplished by frrst
using a CDV flow area of 0.24 sq. in. and an OCV area
of 0.035 sq. in. to provide low oxygen and tire1 flow
rates. This chilldown process is termed trickle
chilldown. Then just prior to start, the valves a~
opened to their prestart positions to again flush any
saturated propellants Out Of the feedlimes. The
significant reduction in propellant consumption required
fbr both fvst and second start chilldown more than
of&s the higher RLlOE-I engine weight, while
eliminating potential fkhure modes associated with air
ingestion and moisture contamination during ascent.
Engine startup is accomplished in a manner similar
to the RLIOA41 startup process, but with the
following differences. At the engine start command the
DCV closes to the ignition area. Shortly themafter, the
FSV onens and the CDV closes to a bleed area. This
sequence with the OCV lead minimizes the ignition
pressure rise and improves torch igniter operation as
described later. The torch igniter is lit by redundant
spark plugs and lights the main chamber. Two pressure
sensors, each electronically redundant, are used by the
DEREC to trigger the valve positions during the start
transient. One sensor measures oxygen pump d&charge
pressure and higgen the opening of the DCV to its
steady state flow area after ignition. The other sensor
measures fuel venturi upstream pressure and triggers the
CDV closed (to control pump stall) and the TCV open
(to control acceleration).
At?er ignition, the engine accelerates open loop to
the 55% thrust level and briefly pauses at that point.
Then the control loop is closed, with the control
parameter being the FSV upstream pressure (FSVUP),
and the engine is ramped up to full power by reducing
the TCV flow area (i.e., reducing the turbine bypass
flow). Mixture ratio is maintained during the ramp by
scheduling the DCV position relative to an FSVUP
control pressure request. The schedule for each engine is
based on development test data and/or acceptance test
data. Once the steady state operating level is achieved,
the closed loop cotmol system maintains a constant
FSVUP using integral and proportional control. Since
FSVUP is exposed only to hydrogen gas, it was
selected as the control parameter to eliminate any
possibility of control sensor moisture contamination
!?om the main chamber combustion products (i.e.,
steam). The RL I OA4- I usesconstantchamber pressure
control. The impact of controlling FSVUP. instead d

48th IAF Congress

chamber pressure, on the engine thrust versus mixture
ratio is insignificant’.
Engine shutdown is initiated with a closed loop
ramp down to 55% power by opening the TCV. Again,
mixture ratio is maintained during the ramp by
scheduling the OCV. The engine can pause at the 55%
power level until the shutoff command is given, if
desired. Shutdown is completed similar to the RLIOA4-1 by closing the FSV and inlet valves and opening
the CDV. Valve timing was optimized for RLIOE-I
operation to reduce pressure loads and backflow d
combustion products into the chamber supply systems.
The RLIOE-I test program was highly successful
in demonstrating improved engine operation with the
new control system and reduced number of valves’.
Demonstrated areas of improvement included more
efftcient chilldown and reduced propellant consumption
for engine start, reliable torch ignition, reduced start and
shutdown loads enabled by part power transients,
reduced inlet pressure drops at start due to slower flow
acceleration, improved start and shutdown repeatability,
higher precision steady state thrust control, improved
fault tolerance, and greater control system flexibility.
The RL I OE- I development test program included
component testing and engine testing. Both occurred
with information learned l?om each
program benefiting the other. Component testing was
based on MIL-STD-IS4OB’ requirements tailored to
engine component testing. Engine testing was planned
as a three stage process: a risk reduction phase, a
development phase. and final qualification.

The risk reduction phase took place in 1995 and
early 1996 prior to the System Critical Design Review.
These tests included early hot fire testing of a prototype
engine (designated XR601) simulating the RLIOE-I
configuration, vibration/acoustic testing of an engine
using mass simulators, and pump chilldown coldflow
A primary goal of the XR601 testing was to
demonstrate engine operability using electrohydraulic
valves and other modified valves which actuated at the
slew rates proposed for the RLIOE-I . The XR601
engine started as an RLIOA-4-I configuration engine,
but was progressively
component to a simulated RLIOE-I contiguration.
Gradual changes were made to reduce the risk to the
engine and to determine the impact of each change. The
electrohydtaulic valve timing was controlled to mimic
EMA actuated valves by a facility controller, and the
controller software allowed convenient changes in
control logic and control parameters to quickly respond


to information gained during the testing.
innovative approach was extremely successful and
provided valuable infotmation needed to complete the
preliminary design’.
The progressive component changes to a simulated
RLIOE-1 configuration was successful and led to a
simulated RLIOE-1 configuration by the 7th firing.
Accomplishments during the later ftigs
included the
following: a design of experiments for control valve
trigger levels; chilldown concept demonstrations; torch
ignition operation; fuel pump stall margin and safe
shutdown with a single CDV; part power start and
shutdown; and steady state MR and thrust control with
control gain, frequency response, and control parameter
testing. In addition, two modes of open loop steady
state operation were successfully demonstrated on the
prototype engine. Both methods showed a reduction in
chamber pressure oscillations as compared to the
RLlOA-4-I. XR601 hot fire testing included a total d
48 firings.
Test Prw
The second phase of the engine test program was
the development test phase. The scope of the program
was based on the number of engine samples required to
validate new operating characteristics and a review d
past RLlO development
development testing per Reference 3. In addition,
consideration was given to the extent of the changes in
the basic engine configuration. In particular, the thrust
level remained essentially constant and the basic
powerhead, including turbopump, injector, and chamber
were unchanged. Successes of the early engine tests
were also considered.
It was determined that three development engines
XR602, XR603, and XR605),
qualification engine (designated XR604), and a total aF
I35 development fuings were required to achieve a
minimum acceptable level of confidence that infant
and/or random failures would be identified’. The first
two development engines were to be new engines while
the third would be a rebuilt engine with limited new
components (including the torch igniter, TCV, CDV
and OCV). The DEREC was to be common to all three
development engines, since it and the associated
software would be adequately exercised by a rigorous
bench level validation series.
A type of accelerated mission testing (AMT) was
proposed to satisfy the 135 ftig
goal without
exceeding program budget constraints. AMT is used
extensively in the testing of gas turbine engines to
increase the high stress cycle time on the engines. The
new RLIOE- 1 components primarily a&coed engine
start and shutdown. Therefore, testing focused on start


48th IAF Congress

resulted from the Boost Phase Chilldown Initiative, the
Modem Electronic Igniter Initiative, and the Engine
Control Initiative. Table I shows the contributions due
to each. Note that the vehicle reliability is estimated to
increase by 1.55%. Reliability improvements were
calculated using a baseline reliability model, with
component failure rates calculated using MIL-HDBK2 l7F6, NPRD-95’, and other available data.

and shutdown cycles, for which the multiple relight
capability of the RLIO could be used to reduce test
stand time and cost, rather than long duration steady
state operation. An average of three firings per test setup
were planned. As many as five firings per setup were
tested with XR60 I and XR602.
Many of the fuings were planned for part power
operation only. Part power tests are representative
ftings because 95% of the engine control system is
utilized during engine start and shutdown to 55%
power including
start and
shutdown valve movement, and closed loop control.
The goal was to have 90 full power ftigs out of the
total 135 firings. Based on the RLIOA-3-3A experience,
starts to part power were expected to extend chamber
life and result in fewer chambers having to be allocated
to the program. The engine thrust chamber is qualified
for 28 starts to the full power 22,300 Ibf. thrust level.
Although this is more than adequate for the two bum
mission requirement, extending chamber life is clearly
beneficial for development test purposes.
The first RLIOE-I development engine, XR602,
was hot f& from 30 August to 27 September 1996,
and experienced 34 firings with 3055 seconds of run
time‘. Primary accomplishments
included start and
shutdown design of experiments testing, demonstration
of chilldown. torch ignition, thrust levels from 52 to
104%, mixture ratio control. and shutdown from lOO%,
70% and 55% power. Engine test instimentation
the most extensive ever for an RLIO, with over 900
parameters recorded during each firing.
The second development engine, XR603. was hot
fired from I5 March to 26 April 1997, and experienced
20 firings with 3376 seconds of run time. Primary
beyond those demonstrated
XR602 included chilldown margin testing, shutdown
from high and low MR. EMA removal and replacement
capability, operation at minimum controller voltage,
and thrust overshoot control.
The AREP program was terminated by the USAF
for its convenience in April 1997. Thus the third
development engine and the qualification engine were
never tested. Nevertheless, the development engine test
results for XR602 and XR603 have demonstrated many
of the significant reliability enhancing features of the
RL I OE- 1 engine.


The reliability improvement goal of the AREP
program was to provide the vehicle with a delta
reliability increase of one and a half percentage points
and to decrease the engine failure rate by a relative 7 I %.
These goals were met. Reliability

Table 1: Vehicle reliability improvement
breakdown by initiative.




Boost Phase Chilldown


Engine Control System


Modem Electronic Igniter




Boost Phase
Boost phase chilldown, made possible by the RLIOE-1
control system,
increases vehicle reliability
eliminating the need for a ground chill system. A
ground chill system makes the engines susceptible to
moisture contamination until the vehicle passes out d
the atmosphere. A failun in the original prelaunch
cooldown system was determined to be responsible for
the flight failures of A/C-70 and A/C-7 I. Although that
flight failure mode was eliminated prior to return to
flight through the use of redundant pyrotechnic isolation
valves (with positive indications of closure prior to liftoft) and “shallow” chill (i.e., higher lift& engine
temperatures), ground prechill remained as part of the
vehicle operations. Boost phase chilldown eliminates
the potential for moisture contamination. In addition, it
reduces requirements for engine flight purges, since the
fuel system cooldown valve remains closed during
lower altitude portions of vehicle ascent.
The effectiveness of boost phase chilldown to
adequately chill down the engines was examined during
development testing. The engine test stand propellant
plumbing was modified to simulate vehicle geometry
and allow percolation of oxygen vapor &otn the engine.
A flight-like engine radiation shield was also installed.
Oxygen pump boost phase percolation chilldown and
fuel pump boost phase trickle chilldown were tested at
both minimum and maximum pump inlet pressures.
The vehicle external ascent pressure profile and the tank
pressure profile were also simulated.
Boost phase chilldown development test results fcr
the oxygen and fuel pumps at minimum inlet pressures

48th IAF Congress

are shown in Figure 5. Adequate chilldown
accomplished well within the available ascent period, as
by proper propellant pumping and
successful engine starts. The tests also indicated that
the flight ascent pressures tested had little cffea on the
required chilldown
time, although more fuel is
consumed at the higher pressures.


simulated long coast missions, the chilldown tests used
a 52 set oxygen ttickle chill (0.035 sq. in. OCV flow
area) and a 54 KC tire] trickle (0.24 sq. in. CDV flow
area). These trickle flows were followed by a 10 set
oxygen prestart (0.35 sq. in. OCV area) and an 8 set
fuel prestart (I.15 sq. in. CDV flow area). Figure 6
shows test results for a simulated long coast second
bum chilldown.
Tests showed that higher tank
pressures decreased the chilldown time required, but
also decreased
chilldown time margin would be used for flight.

As described previously, the electronic engine
control system includes six newly designed valves, the
EMA actuators for these valves, and the DEREC. This
new control system increases reliability and provides
many advantages with respect to the current system.
The RLIOE-I control system is much less
complex than the RLIOA control system, since the
former uses only six valves (compamd to the twelve)
and the designs of these new valves are much simpler.
The simple valve designs should increase producibility
and reduce acceptance test requirements for these
components. Also, clever exploitation of the DEREC
and electronic control system may help reduce engine
hot tire acceptance test requirements. Furthermore,
electronic control will help ensure high reliability by
providing greater checkout capability prior to launch.

Figure 5: Simulated boost phase pump chilldown
at minimum inlet pressures (XR603 Run 17.01).
Trickle (i.e., low flowmte) chilldown, also made
possible by the RLIOE-I control system, does not
provide any significant reliability increase of its own,
but it is desirable because it reduces the amount d
propellants consumed during chilldown compared to
the current process. Thus, it increases the payload
capability of the vehicle, which could be used to
provide margin against a low performing engine (an
indirect reliability improvement).


1.- .-



_ _.











__~_ /’






Figure 6: Simulated
chilldown at minimum
I 14.01).

long coast second burn pump
pressure (XR603

One of the major reliability enhancements inherent
to the RLIOE-I control system is its extensive
redundancy and significant fault tolerance capability.
The engine controller is capable of accommodating any
single failure in the control system. A single failure in a
control sensor, valve position control, or a DEREC
driver board results in an automatic switch from channel
A to channel B. Furthermore, even if both redundant
sensors fail at a particular location, the engine software
will reconstruct the missing parameter based on the
other measured control parameters. This engine
robustness was demonstrated during a simulated dual
failure of the fuel venturi upstream pressure control
sensor, which provides the start triggers for the CDV
and TCV. After failun detection,
the missing
measurement was simulated by the DEREC sofhvare
using input from the other coneol pressures, and the
CDV and TCV were triggered successfully.

Runs 12.01

The effectiveness of trickle chilldown was also
examined during development testing. Prior to each
test, the oxygen and fuel pumps were thermally
conditioned to their predicted maximum end-of-coast
temperatures with a helium prechill system. For

The baseline start (Figure 7) was initially
developed using engine analytical models. It was later
modified based on development test results from an L-8
Taguchi design of experiments @OX), in which the
pressure levels, or trigger values, at which the CDV,
OCV. and TCV moved to their part power position
were varied during eight relight tests.
parameters optimized during the DOX testing included

48th IAF Congress


igniter backflow margin, time to accelerate, chamber
pressure overshoot, fuel pump stall margin, and oxygen
pump suction pressure drop during acceleration.

Figure 7: Baseline start and shutdown chamber
pressure profile (XR602

Run 9.01).

Figure 8 shows the chamber pressure during the
start transient for the eight DOX tests. The e&t d
varying valve triggers was significant as expected, but
all eight start attempts were successful. The results
demonstrate the robusmess of the RLIOE-I system. In
addition, the engine proved to be insensitive to
variations that might result if a sensor or valve faihae
caused a delay in valve trigger prior to control system
channel switch to the backup channel.

Engine start Design
Figure 8:
chamber pressure (XR602).



Development testing demonstrated successful
engine starts at all combinations of fire1and oxidizer
pump extreme inlet conditions. The part power pause at
the 55% thrust level was shown to reduce the inertial
pressure losses in the facility oxygen feedliie at SM
from 13.5 to 9.5 psi. This benefits ground testing d
minimum inlet pressures. Maximum current draw
during the start transient was 24 amps for less than
0.005 seconds which was well under the transient 50

amps peak draw available from the vehicle.

The baseline shutdown is also shown in Figure 7.
Upon receiving a ramp down command &nn the
vehicle. the engine DEREC ramps thrust down fiun
100% power to 55% power along a 5 second closed
loop ramp. The engine dwells at this part power
position until the vehicle commands engine shutdown.
The 5 second ramp was utilized for ground testing due
to the test stand exhaust diSuser limitations. Faster
ramps could be tested during short duratioo tests, but
this was never attempted befbm the program was
terminated. Additional DOX testing was conducted to
optimize the valve command sequence at shutdown.
Optimization pammeters included peak fuel system
characteristics, DEREC cumnt draw, and shutdown
Development testing demonseated successful
engine shutdowns from power levels of 55 to 104% and
6-om MR of 4.9 to 5.8. The engine shutdown thrust
transient was very repea@le, with run-to-run variation
in shutdown time to 10% thrust less than 0.030 set for
all cases. Part power dwell times from zero to over 30
seconds were also demonstrated with no measurable
diffemnce in engine shutdown transients. Maximum
measured current draw during shutdown was 48.8 amps
fa less than 0.005 seconds (the actual cumnt draw is
less, since the measurement included electronic noise
that could not be eliminated).
The RLIOE-l’s baseline start to 55% thrust,
followed by a brief pause and controlled ramp to 100%
thrust, and its baseline shutdown, initiated with a
controlled ramp to 55% thrust and followed by a brief
pause before fml shutdown, provide significant
reliability enhancements by themselves. These “soft
transients” generate reduced loads at start and shutdown
for the engine, vehicle, and payload. The so!? start
transient also offas the vehicle benefit of reduced start
differential between the two engines, which is expected
to improve guidance, lessee tank slosh, and themby
reduce subsequent tank pressurization gas usage. The
soft shutdown transient. on the other hand, also
The electronic control system appears to provide
benefits for steady state thrust control as well.
Development testing showed that the engine control
valves, EMAs. sensors, DEREC, and software all
perform as designed during steady state operation.
Mixture ratio control at all propellant inlet condition
extremes and with tank pressurizAoo bleed flow waz
demonstrated. No control system hardware changes
were identified. Soffware gains were modified based on
the initial test results.
Steady state thrust repeatability was better than
HL5% of full power thrust. Thrust oscillations wae

48th IAF Congress

f l%, a factor of three better than previous RLlOA
engines. These reduced oscillations should benefit the
spacecraft. Testing at mixture ratios ranging from 4.7 to
6.0 demonstrated run-to-run mixture ratio repeatability
of kO.08 MR units, and single-run repeatability d
iO.03 MR units. Closed loop control demonstrated a
maximum rate of 0.062 MR units/set with good
control system accuracy and response.
Another major advantage of the electronic control
system is the considerable flexibility it provides. This
flexibility allows for continuous optimization of engine
operation as appropriate. Engine conditions can be
adjusted for new mission requirements or to address
undesirable operating conditions. The capability to
operate the engine at reduced thrust level for extended
duration offers unique possibilities for future missions
planning. The flexibility of the control system was
exploited with great success to reduce backflow in the
torch, as will be described later. Other modifications are
possible, as well. For example, minor changes to
control valve areas could be made to reduce shaft
speeds. provide larger critical speed margins, or allow
greater mixture ratio operating ranges. The full
implications of having complete control of valve timing
and slew rates were still being explored when testing
was halted.

As described previously, the RL I OE- I engine uses
an improved. dual spark plug, continuous burning torch
igniter with redundant. modem electronics. The torch
igniter eliminates the need for the igniter gaseous
oxidizer (GOX) required by the RLIOA ignition
system. The GOX valve is prone to freezing after the
first bum and staying closed for the remainder of the
mission. Thus, the oxidizer required for reliable
ignition during RLIOA engine restartz must be
provided by extended duration prestart cooldown (with
the associated excessive propellant usage).
Development testing has demonstrated that the
torch igniter provides more reliable ignition and
increased ignitability compared to the RLIOA ignition
system. The torch ignition system successfully lit the
main engine for all runs, even though the majority of
the tening was performed with only one spark plug (the
other plug port generally was used for instrumentation).
In addition. the torch ignited reliably with prestarf
cooldown times well below that required for engine
A major concern prior to development testing was
the effect of a momentary backflow of combustion
products (steam) into the oxidizer circuit during the
rapid chamber pressure rise at start. Moisture deposited
during backflow could l?ecze and block the oxidizer flow
for the subsequent start. It is such oxygen side backflow


that prevents implementation of the torch on the
RLIOA model engines. However, engine testing has
demonstrated that tbe RL IOE-I electronic control
system can eliminate this backflow by starting to part
power. In fae a part power pause as short as 0.5 set
provides the necessary margin against backflow. A 0.75
second pause at 55% thrust was selected for the baseline
start to allow additional time for conditions to stabilize
prior to closing the control loop.
In all of the development testing to-date, fuel side
blockage has never been observed. Nevertheless, test
series were run to optimize the ignition timing to
minimize fuel side backflow at start. The best timing
was determined to be a 0.400 set lead in the OCV
closure at start and a 0.350 set delay in the ignition
timing. This resulted in a blowdown of the oxygen
pressure and a lower chamber pressure rise at ignition
which minimized (but did not eliminate) the fuel side
reverse pressure and backflow.
Development testing did identify two significant
anomalies with the torch igniter, however. One
anomaly was that the ceramic insulator on the igniter
spark plugs was cracking. Although the cracks resulted
in multiple sparks to light the engine, the engine still
ignited on every test. Thermal gradients at start and
shutdown were identified as the most likely cause of the
cracks. An insulator redesign with a tougher ceramic
material was under development at the time the AREP
program was canceled.
The other anomaly was ice formation on the
oxygen injector. Instrumentation indicated that the
igniter was overcooled resulting in moisture condensing
on the internal wails during steady state operation. The
torch combustion chamber is cooled by hydrogen gas in
parallel flow with the hot gas. Bench testing and a
unique engine video of the igniter revealed that the
moisture would flash off at shutdown when the engine
pressures dropped rapidly. The moisture was observed
to condense and fiuzc on the cold oxygen injection
post. The ice was observed to remain for as long as
three minutes
before melting/sublimating.
phenomena is suspected to be the cause of oxygen
supply line blockage indicated in one run for XR601.
The igniter lit the engine for this run despite the
blockage using the residual oxygen t?om the prestart
period. A torch redesign to increase the wall
temperatures above moisture sustaining conditions by
using a counterflow hydrogen coolant scheme was under
and tested with promising

Other potential reliability enhancing features are
available with an electronically controlled RLIO engine.
Some are described in detail below.

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