William M. Marshall and Julie E. Kleinhenz

Glenn Research Center, Cleveland, Ohio

Analysis of 100-lbf (445-N) LO2-LCH4 Reaction 

Control Engine Impulse Bit Performance

NASA/TM—2012-217613

June 2012
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William M. Marshall and Julie E. Kleinhenz

Glenn Research Center, Cleveland, Ohio

Analysis of 100-lbf (445-N) LO2-LCH4 Reaction 

Control Engine Impulse Bit Performance

NASA/TM—2012-217613

June 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors greatly acknowledge the entire facilities team which operated the Altitude Combustion Stand (ACS) facility and 

made testing possible. Additionally, the support received by the NASA Glenn Materials Branch is also appreciated, 

as well as discussion and insight provided by the engine manufacturer, Aerojet. This work was supported by the 

Propulsion and Cryogenic Advanced Development (PCAD) project, which was part of the 

NASA Exploration and Technology Development Program (ETDP).

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217613 iii 

Contents 

1.0 Introduction ............................................................................................................................................ 1

2.0 NASA Glenn Altitude Combustion Stand Facility ................................................................................. 2

3.0 Propellant Conditioning Feed System (PCFS) ....................................................................................... 3

4.0 Hardware Description ............................................................................................................................. 3

4.1 Exciter Sources ............................................................................................................................. 4

4.2 Injector .......................................................................................................................................... 5

4.3 Columbium Chamber ................................................................................................................... 5

4.4 Instrumentation ............................................................................................................................. 5

4.4.1 Mass Flow Rates ............................................................................................................... 6

4.4.2 Thrust Measurement ......................................................................................................... 7

4.5 Timing and Permissives ................................................................................................................ 8

4.6 Test Matrix ................................................................................................................................... 9

4.7 Hardware Issues During the Test Series ....................................................................................... 9

4.7.1 Restricted Abort Windows and Ignition Delay ............................................................... 10

4.7.2 Inoperative Methane Recirculation Pump ....................................................................... 10

5.0 Methodology ........................................................................................................................................ 11

5.1 Definition and Calculation of Impulse Bit (I-bit) ....................................................................... 11

5.2 Objectives of Pulse Testing ........................................................................................................ 12

6.0 Results .................................................................................................................................................. 12

6.1 Uncertainty Analysis of I-bit ...................................................................................................... 13

6.2 Impulse Bit Testing Results ........................................................................................................ 15

6.3 Compact Unison Exciter ............................................................................................................. 20

6.4 Anomalies Observed During Test Series .................................................................................... 20

6.4.1 Thrust Stand Ringing ...................................................................................................... 20

6.4.2 Thrust Stand Zero Drift ................................................................................................... 21

6.4.3 Double Pulse Phenomenon ............................................................................................. 22

6.4.4 Temperature Rise in Feed System ................................................................................... 26

6.4.5 Columbium Coating Damage .......................................................................................... 27

7.0 Summary and Conclusions ................................................................................................................... 30

8.0 Related Work ........................................................................................................................................ 30

References ................................................................................................................................................... 31

Appendix A.—Acronyms and Symbols ...................................................................................................... 33

Appendix B.—X ray and Scanning Electron Microscope Results.............................................................. 35

B.1 X ray Diffraction: Undamaged (Baseline) .................................................................................. 35

B.2 X ray Diffraction: Damaged ....................................................................................................... 36

B.3 Scanning Electron Microscope Undamaged ............................................................................... 37

B.4 Scanning Electron Microscope Damaged ................................................................................... 39

Appendix C.—Test Data Log ..................................................................................................................... 41

NASA/TM—2012-217613 1 

Analysis of 100-lbf (445-N) LO2-LCH4 Reaction Control Engine 

Impulse Bit Performance 

William M. Marshall and Julie E. Kleinhenz 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Summary 

Recently, liquid oxygen-liquid methane (LO2-LCH4) has been considered as a potential “green” 

propellant alternative for future exploration missions. The Propulsion and Cryogenic Advanced 

Development (PCAD) project was tasked by NASA to develop this propulsion combination to enable safe 

and cost-effective exploration missions. To date, limited experience with such combinations exist, and as 

a result a comprehensive test program is critical to demonstrating with the viability of implementing such 

a system. The NASA Glenn Research Center conducted a test program of a 100-lbf (445-N) reaction 

control engine (RCE) at the Center’s Altitude Combustion Stand (ACS), focusing on altitude testing over 

a wide variety of operational conditions. The ACS facility includes unique propellant conditioning feed 

systems (PCFS), which allow precise control of propellant inlet conditions to the engine. Engine 

performance as a result of these inlet conditions was examined extensively during the test program. This 

paper is a companion to the previous specific impulse testing paper, and discusses the pulsed-mode 

operation portion of testing, with a focus on minimum impulse bit (MIB) and repeatable pulse 

performance. The engine successfully demonstrated target MIB performance at all conditions, as well as 

successful demonstration of repeatable pulse widths. Some anomalous conditions experienced during 

testing are also discussed, including a double pulse phenomenon, which was not noted in previous test 

programs for this engine. 

1.0 0BIntroduction 

To enable future exploration of the Moon, Mars, and beyond, next-generation propellant systems are 

being developed. With an emphasis on nontoxic, “green” propellants, liquid oxygen-liquid methane (LO2-

LCH4) has risen to the forefront. LCH4 is an attractive propellant because it does not require the strict 

thermal storage requirements of hydrogen (due to its larger density and higher boiling point), nor does it 

require the rigorous handling protocols of toxic hypergolic propellants. It also has the potential, when 

paired with LO2, to produce higher specific impulse than either the existing hypergolic or LO2-kerosene 

systems. This higher specific impulse and improved storage capability could reduce vehicle mass when 

compared with hydrogen-based systems, since smaller propellant storage and management would be 

required. Not only is there a potential for decreased vehicle mass, these propellants can also be produced 

on Mars using local resources. Prior work with this propellant combination is limited, so a goal of the 

NASA Propulsion and Cryogenic Advanced Development (PCAD) project was to examine the feasibility 

and performance characteristics of these systems (Refs. 1 to 3). In particular, there is interest in 

demonstrating repeatable and reliable ignition of an engine over a wide range of valve inlet temperatures 

(from liquid-liquid operation to gas-gas operation), especially at vacuum conditions (Refs. 3 and 4). 

To facilitate this, a 100-lbf (445-N) LO2-LCH4 Reaction Control Engine (RCE) was developed by Aerojet 

Corporation (Ref. 5). In late 2009 and 2010, this engine underwent a series of tests in the Altitude 

Combustion Stand (ACS) at the NASA Glenn Research Center (Refs. 6 and 7). Two specially designed 

Propellant Conditioning Feed Systems (PCFS) were developed to enable precise propellant temperature 

control. The first test series (Refs. 6 and 7) at ACS examined specific impulse performance with burn 

durations up to 7 s. The engine met the Isp,vac goal, achieving an overall average Isp,vac of 305 s, ±4 percent. 
NASA/TM

Performan

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M—2012-21761

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M—2012-21761

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M—2012-21761

F

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M—2012-21761

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M—2012-21761

re 6: Feed sys

Mass Flow Ra

rate measurem

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were calculat

(Ref. 15). A s

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temperature 

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NASA/TM

Figure 7

mete

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behavior f

pulse that

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were perm

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4.4.2 23BT

The th

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Calibr

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thermal fl

applied to

verified th

M—2012-21761

7: Flow meter m

r and the turbin

s noted from p

ow meters un

pulse widths m

rate measurem

for the short p

t then decayed

Therefore, te

mitted to vary

time average v

Thrust Measu

hrust stand ut

d cell measure

th a 0- to 100

ration of the l

ut after the pr

(by introducin

lexure of the t

o the summed

he thrust stand

13 

measurements

ne flow meter r

previous testi

nder steady-sta

made it diffic

ments from th

pulse widths. 

d down, while

est conditions

y based upon t

values over th

urement 

tilized a tri-lo

ements, after a

0-lbf (0- to 445

load cells was

ropellant lines

ng cryogenics

thrust stand. T

d load to calib

d calibration 

 for a single 80

responses. 

ing that the tu

ate conditions

cult to accurat

e different flo

The turbine f

e the venturi f

s were based u

the transient b

he pulse durat

ad cell config

appropriate z

5-N) range (f

s accomplishe

s and associat

s up to the thr

The test pull t

rate the test th

did not signif

7 

0-ms pulse illus

urbine flow m

s. However, t

tely determine

ow meters, ill

flow meters h

flow meters s

upon steady-s

behavior of th

tion based up

guration. The 

ero offsets we

for a total thru

ed by perform

ted hardware 

ruster valves)

then provided

hrust measure

ficantly drift t

strating the diffe

meters showed

transient respo

e or adjust m

lustrating the 

have a very sh

showed a mor

state set point

he engine feed

pon the ventur

total thrust w

ere applied. E

ust range of 0

ming a test pu

had been chil

. This minimi

d a calibration

ements. A fin

throughout th

ference betwee

d good agreem

onse of both i

mass flow rates

distinctly dif

harp rise at th

re smoothed p

ts, and the flo

d system. The

ri flow meter 

was the summ

Each load cell

0- to 300-lbf (0

ull on the live 

lled down in 

ized thrust er

n factor (K-fa

nal test pull at

he test day. 

en the venturi fl

ment with the 

instruments d

s. Figure 7 sh

fferent transie

he beginning o

peak througho

ow rates and M

e reported flo

measuremen

mation of these

l was a dual b

0- to 1334-N)

bed prior to 

the vacuum 

rrors due to 

actor), which w

t the end of th

low 

during 

hows 

ent 

of the 

out 

MRs 

ow 

nts. 

e

bridge 

)). 

was 

he day 
NASA/TM—2012-217613 8 

4.5 13BTiming and Permissives 

In order to ensure proper conditions are met for safe and controlled tests in ACS, two different 

checks are conducted. The first checks are system permissives. These are conditions that must be met 

prior to a test and include “rig-ready” signals from the PCFS, proper power and electrical conditions to 

the test cell, and proper supply pressure for system pressurants. Once the permissives are met, operators 

are then permitted by the PLC to start a test. Timing for a test at ACS is handled in four zones, which 

are described in Reference 7. Unless otherwise noted, only Zone 2, the hot fire portion, is discussed in 

this document. In order to achieve pulse testing, the command timing of Zone 2 encompassed a single 

pulse. This zone was then repeated to achieve the desired number of consecutive pulses. The timing of 

a pulse included a 10-ms LO2 lead on startup, which was intended to bring both manifolds up to 

pressure at the same time based on cold propellant flow tests. A 5-psia (0.03-MPa) nitrogen gas purge 

was introduced through the igniter cavity pressure port (Pc,ign) and manifold pressure ports (PFJ and 

POJ) between pulses using a check valve. This purge ensured that no combustible gas was trapped in 

the manifolds between pulses, minimizing the chances for a hard start. The igniter sparks were initiated 

20 ms before the LCH4 flow. Table I lists the command timing with respect to the zone start, and the 

duration that the valve remains in that commanded state. Thus, a valve (or abort window) that opens at 

30 ms with a duration of 50 ms would mean the valve opened at 30 ms into zone 2 and closed at 80 ms 

into zone 2. The minimum timing resolution of the PLC was 10 ms. The implications of this timing 

schedule are discussed in Section 4.7.1 below. At the end of the pulse, the LO2 valve was closed with 

the LCH4 following 40 ms later. This fuel- rich shutdown was intended to prevent oxidation on the hot 

chamber walls. 

No attempt was made for a “pre-chill” (pre-flow of LO2 to lower the injector hardware temperature) 

of the engine prior to start of test. However, the propellant lines were maintained at chilled conditions 

up to the thruster valves during and between runs by means of trace cooling around the propellant lines. 

Since no pre-chill was conducted, the injector often started at near ambient temperature (≈500 °R 

(278 K)). Therefore, there was a considerable transient on startup as the liquid propellant phase changes 

to gas phase in the manifolds and the manifolds subsequently chill-in. 

The electric pulse width (EPW) is the time when both thruster valves are commanded open. In the 

current study, this is between the opening of the methane valve and when the oxygen valve closes. 

The duty cycle is the EPW divided by the total run time for a single pulse. Thus, for 80-ms EPW with a 

5-percent duty cycle, the dwell time between pulse starts would be 1600 ms (1.6 s). Figure 8 shows a 

basic timing sequence used for these series of tests. Unless otherwise noted, this timing was maintained 

for all tests described here. 

TABLE I: TIMING FOR VARIOUS PULSE WIDTHS (EPWs) USED IN PULSE TESTING 

EPW, 

ms 

Zone 2,a

ms 

(% duty cycle) 

Spark 

on/duration,

ms 

Oxygen 

on/duration,

ms 

Fuel 

on/duration,

ms 

Pcav abort window 

start/duration, 

ms 

500 980 (50) 230/100 240/500 250/510 350/300 

100 380 (25) 30/100 40/110 50/140 120/20 

80 1580 (5) 30/90 40/90 50/120 120/10 

40 780 (5) 30/60 40/50 50/80 OFF a

Zone 2 is repeated for desired number of pulses. Zone 2 is left 20 ms short to account for 

programmable logic controller (PLC) delay in repeating the zone. 

 
NASA/TM

Fig

w

4.6 14BT

For th

conditions

shows the

small shad

discussed 

4.7 15BH

Durin

(RCE) bu

included d

LCH4. 

M—2012-21761

gure 8: Basic tim

width. A pulse t

Test Matrix 

he pulse test s

s explored in 

e planned targ

ded boxes are

in Reference

Hardware Is

ng the test seri

ut which contr

difficulty in r

13 

ming sequence

train is accomp

C

N

W

series, the inte

the prior dura

get temperatur

e the target co

e 7. 

ssues During

ies, some issu

ributed to the 

esolving very

e for individual 

plished by repe

TABLE II: PLA

INLET TEM

Condition 

Cold

Nominal 

Warm 

erest was testi

ation testing. 

re ranges (box

onditions, the 

g the Test S

ues were iden

quality of dat

y short EPW a

9 

pulses for puls

eating the abov

ANNED TEST MA

MPERATURES, 

Propella

LCH4

170 (94) 

204 (113) 

224 (124) 

ing the engine

Table II lists 

xes) with the 

larger open b

Series 

ntified that we

ta collected a

and the inabil

se tests. EPW r

ve sequence fo

ATRIX OF 

°R (K) 

ant 

LO2

163 (90) 

204 (113) 

224 (124) 

e over the sam

 the propellan

actual condit

boxes are the 

ere not directl

and ability to m

lity to reach th

refers to electro

or the given num

me matrix of 

nt target temp

tions shown. I

modified targ

ly related to th

meet test obje

he target cold

onic pulse 

mber of pulses

temperature 

peratures. Fig

In Figure 9, th

get conditions

he test engine

ectives. These

d temperature

. 

ure 9 

he 

s as 

e

e

es for 
NASA/TM

4.7.1 24BR

The ti

monitorin

of the win

above col

previous t

occur, esp

testing (R

random pu

chamber p

abort later

restricted 

series. Th

ignition, b

4.7.2 25BIn

Durin

the metha

further tem

This parti

increasing

number o

program. 

conditions

M—2012-21761

Figure 9: P

condition

Restricted Ab

iming resolut

ng. The PLC c

ndow. For 80-

d flow pressu

testing (Ref. 7

pecially at col

Ref. 7) illustra

ulse in a plan

pressure from

r would cause

abort window

hese aborts we

but rather wer

noperative M

ng the later po

ane PCFS bec

mperature ref

cularly impac

g trace coolin

f tests to be c

Note that due

s were actuall

13 

Propellant temp

ns based upon

bort Window

ion of the AC

checks to mak

-ms EPW test

ures was relat

7), and confir

lder propellan

ting observed

nned pulse trai

m rising above

e the abort to 

ws and the po

ere not ignitio

re a conseque

Methane Reci

ortion (Run 22

ame inoperat

finement was 

cted testing at

g bleeds, etc.

onducted. Fig

e to loss of th

ly closer to no

peratures at the

 the planned te

ws and Ignitio

CS facility PL

ke sure that pr

ts and shorter

ively short, o

rmed during th

nt temperature

d ignition dela

in. A delayed

e the abort lim

be tripped on

ossibility of ig

on failures, as

ence of the ab

irculation Pu

24 to end of s

tive. As a resu

lost, and only

t cold-cold co

, were emplo

gure 9 shows 

e methane rec

ominal condit

10 

e thruster valve

est matrix. 

on Delay 

LC was limited

ressure is abo

r, the period o

n the order of

he present ser

es. Figure 10

ay. During pu

d ignition, due

mits prior to th

n the shutdow

gnition delay, 

s the evidence

ove considera

ump 

series) of the t

ult, the capabi

y a single pas

onditions. Eff

yed to the ext

the plot of va

circulation pu

tions. 

e for all tests. T

d to 10 ms, in

ove a specifie

of time that th

f 10 to 40 ms

ries, that igni

shows a plot 

ulse testing, th

e to the short 

he window en

wn for non-del

several pulse

e from the dat

ations. 

testing, the m

ility to recircu

ss for tempera

forts to furthe

tent feasible w

alve temperat

ump, methane

The boxes show

ncluding abor

ed value throu

he chamber w

s. However, it

ition delays of

taken during

hese delays co

abort window

nd. Subsequen

layed ignition

e trains aborte

ta does indica

methane recirc

ulate the meth

ature conditio

r condition th

while trying t

tures for the p

e temperature

w target 

rt window 

ughout the dur

was sufficiently

t was noted du

f 10 to 40 ms

g specific imp

ould occur on

w, would prev

ntly, shifting 

ns. As a result

ed during the 

ate a successfu

culation pump

hane propella

oning was pos

he propellants

to maximize t

pulse testing 

es for cold 

ration 

y

uring 

 could 

ulse 

n a 

vent 

the 

t of the 

test 

ful 

p on 

ant for 

ssible. 

s by 

the 
NASA/TM

Figure 1

press

5.0 4BM

In this

objectives

5.1 16BD

I-bit i

until the c

(0.12 MPa

consistent

Data f

visual bas

was determ

A pressur

actively fl

decay, the

(10 percen

thrust (Fig

is the I-bi

M—2012-21761

10: Ignition dela

sure. 

Methodolo

s section, the 

s of the test se

Definition an

is defined as t

chamber press

a) for the RCE

t with Aerojet

from each tes

sic program. F

mined by exa

re rise in the m

flowing (Figur

e peak chamb

nt of the nom

gure 11(C)) w

t, in the units

13 

ay during Isp te

ogy

primary perf

eries are also 

nd Calculati

the area under

sure decays to

E, which has 

t’s calculation

st was importe

Figure 11 sho

amining the m

methane mani

re 11(A)). Sin

ber pressure w

minal rated val

was therefore 

 of lbf-s (N-s)

sting as observ

formance metr

discussed. 

ion of Impu

r the vacuum 

o 10 percent o

a 175-psia (1

ns from previ

ed into Micro

ows graphicall

methane manif

ifold indicated

nce the end of

was first identi

lue) after this 

integrated ov

). 

11 

ved in the (A) ig

ric for this tes

ulse Bit (I-bi

thrust curve 

of the rated no

1.2-MPa) nom

ious testing (R

osoft Excel, w

ly how I-bit w

fold pressure 

d that both va

f the integrati

ified. When th

peak, the inte

ver this period

gniter chambe

st program, Iit) 

from the time

ominal chamb

minal chambe

Ref. 5). 

where I-bit cal

was calculated

and the cham

alves were op

ion occurred d

he pressure re

egration ende

d, as represent

r pressure and

-bit, will be d

e both propell

ber pressure (

r pressure). T

lculations wer

d. The I-bit in

mber pressure

pen and prope

during the ch

eached 17.5 p

ed (Figure 11(

ted by the gra

d (B) main cham

defined. The m

lant valves op

(17.5 psia 

This definition

re performed 

ntegration tim

 data sets. 

ellants were 

amber pressu

psia (0.12 MP

(B)). The vac

ay box. The re

mber 

major 

pen 

n is 

in a 

me 

ure 

Pa) 

cuum 

esult 
NASA/TM

5.2 17BO

The o

pulsed-mo

propellant

(Ref. 5). 

(1) A

(2) A

(3) A

(4) O

6.0 5BR

Presen

explanatio

with a dis

M—2012-21761

Objectives of

objective of pu

ode operation

t temperature

Achieve a min

Achieve a min

Achieve a MIB

Operate over a

Results 

ntation of the

on of the expe

scussion of an

13 

Figure

f Pulse Test

ulse testing at

n and to quant

s. The follow

nimum impuls

nimum EPW o

B repeatability

a 0- to 100-pe

e results begin

ected uncertai

nomalies obse

e 11: Calculatio

ting

t NASA Glen

tify the I-bit p

wing technical

se bit (MIB) o

of 80 ms or le

y of ±5 perce

ercent duty cy

ns with a discu

inties. A thor

erved during t

12 

on regime for d

nn was to veri

performance a

l challenges h

of 4 lbf-s (17.8

ess 

nt after achie

ycle 

ussion of the 

rough discussi

the test campa

determining I-bi

ify the operati

at altitude con

had to be addr

8 N-s) 

eving stable te

uncertainty a

ion of results 

aign. 

it. 

ion of the eng

nditions for v

ressed by the 

emperature 

analysis utiliz

 is then provi

gine under a 

varying input 

delivered eng

zed with an 

ided, conclud

gine 

ing 
NASA/TM—2012-217613 13 

6.1 18BUncertainty Analysis of I-bit 

An uncertainty analysis was performed to qualify the I-bit values to fully meet Joint Army, Navy, 

NASA, Air Force (JANNAF) reporting standards (Ref. 16). The equations for I-bit and its uncertainty 

(UI-bit) are 

  F t

P

d

CH valve

10% high I - bit vac

4

cav (3) 

2 2 2 2 2 2 2 UI-bit  F BF  FPF  t Bt (4) 

In equation (3), Fvac is the vacuum thrust and t is the time. The thrust is integrated over the pulse 

width, as described above. In equation (4), B is the bias uncertainty that includes the instrumentation and 

calibration uncertainties, while PF is the precision uncertainty, which represents scatter in the data. Both 

also include a Student’s t-distribution factor for a 95-percent confidence interval, and will be discussed in 

more depth below. The θ values are the sensitivity terms, which are the partial derivatives of I-bit with 

respect to its dependent variables (F, t). Simply they are θF = t and θt = F. 

The bias uncertainty of time, Bt, is assumed to be negligible. The time stamp in the data files is taken 

from the data system clock. The bias uncertainty of the force term,	ܤி౬౗ౙ, is derived from the vacuum 

thrust equation: 





















  

cells

3 load

Fvac K F Pamb Ae (5) 

2 2

amb

2 2 2 2 2 2 2 2 2 2 2 2 2 2 2

vac amb amb e e Pamb Ae BF  F BF  K BK  P BP  A BA  K BF  F BK  Ae B  P B (6) 

Where F is the force measured by the load cells, Pamb is the pressure outside the chamber, Ae is the 

exit area of the nozzle, and K is the calibration factor (calibrations are done at the beginning and end of 

each test day). The bias values were determined in the previous test program and are displayed in 

Table III. The previous program encompassed 55 tests, so the Student’s t-distribution factor was 2 with a 

95-percent confidence interval. The only term in Table III that would change with the current pulse test 

program is the sensitivity of the load cell calibration factor (K-factor), θK. As shown in Equation (6), this 

term is equal to the average force exerted on the load cells over all tests. For the previous program 

(Ref. 7) it was ≈100 lbf (445 N) and for the pulse program it was ≈50 lbf (222 N) in each pulse. Using the 

latter would decrease the bias error, ܤி౬౗ౙ

ଶ , to a value of approximately 13.3. However, the uncertainty 

calculations were programmed prior to the pulse test program so estimates could be obtained in real time. 

Therefore, while the values in Table III are somewhat conservative, they were retained. 

TABLE III: THE BIAS UNCERTAINTY FOR THE VACUUM THRUST MEASUREMENTS 

 Sensitivity binst Bcali B2

 = Θ2 Σ(tb)2

Force (summation of three load cells) 1.033 = θF ±1 lbf each ±0.06 lbf each 12.862 

Force calibration term (K factor) 99.863 = θK ------ 0.005 0.999 

Thrust—area, exit 0.210 = θA ±0.001 in.2 ------ 1.757×10–7

Thrust—pressure ambient 13.710 = θP ±0.015 psi 0.0063 0.204 

ܤி౬౗ౙ

ଶ 14.065 

 
NASA/TM

The p

steady in 

great deal

standard d

polynomi

standard e

thrust data

in the load

Figure 12

445 N), p

of the I-bi

(89 to 178

The I-bit u

condition

propellant

portion of

interval is

for the no

Note t

Since the 

equation. 

uncertaint

reduction 

optimal pr

8.5 percen

accurate r

Figure 12

polyno

fluctua

M—2012-21761

precision unce

these short (8

l of ringing se

deviation calc

al curve fit w

error of the pr

a and the curv

d cell causes 

(A) was the w

eak-to-peak).

it was high. In

8 N), peak-touncertainties 

. A 95-percen

t run has the l

f the curve. H

s relaxed to 90

ominal data in

that the transi

chamber pres

Without the o

ty than its thru

in uncertaint

rediction of th

nt on I-bit and

representation

2: Thrust curve

omial curve fit u

ations. 

13 

ertainty, PF, i

80-ms) pulse d

een in the thru

culation is not

was used to pre

redicted value

ve fit for two 

significant flu

worst in this p

 This resulted

n Figure 12(B

-peak), so the

for each puls

nt confidence 

lowest uncert

However the I0 percent the 

n Figure 13. 

ient nature of

ssure profile s

oscillations o

ust-based cou

y. The polyno

he transient e

d 3.5 percent 

n of the facilit

es for two pulse

used to estimat

s typically ba

durations, res

ust data (see S

t representativ

edict a smoot

es was used to

pulses during

uctuations in 

particular train

d in a poor cu

B) the amplitu

e standard erro

e are shown i

interval is us

tainties becau

-bit uncertain

uncertainties

f the short pul

should trend t

n the thrust st

unterpart. How

omial curve f

effects. The bi

on Pc-bit. Th

ty capabilities

es from a test a

te the thrust pr

14 

ased on the sta

sulting in a be

Section 6.4.1)

ve, since it ne

th, time-depen

o determine th

g a pulse train

the thrust, res

n with high-a

urve fit and hi

ude of the thru

or and the res

in Figure 13 f

sed to calculat

use the thrust f

nty is still 30 p

reduce to 20

lse durations i

the same as th

tand, the resu

wever, the im

fit used here to

ias uncertaint

ese are conse

s. 

at warm propell

recision uncerta

andard deviat

ell-shaped cur

). Using a sim

eglects this tim

ndent thrust r

he precision u

n with warm p

sulting in sign

amplitude fluc

igh error valu

ust fluctuation

sulting I-bit un

for a represen

te the Student

fluctuations t

percent even a

 percent, as re

is the primary

hrust, it was s

ulting “Pc-bit”

mprovement w

o estimate the

ty, which is ti

ervative uncer

lant conditions

ainty for (A) wo

tion. Howeve

rve over time

mple average t

me dependen

response for e

uncertainty. F

propellant con

nificant uncer

ctuations (≈40

ues. Thus, the 

ns were lowe

ncertainty we

ntative test at 

t’s t-distribut

tended to dam

at its best. If t

epresented by

y factor in the

substituted int

” should have

was only a 5- t

e precision un

ime independ

rtainties, but a

s. The solid line

orst case and (

er, the thrust i

. There is also

thrust value in

cy. Therefore

each pulse. Th

Figure 12 sho

nditions. Ring

rtainty values

0 to 100 lbf (1

total uncertai

er (≈20 to 40 l

ere also lower

each tempera

ion. The nom

mpen in the tai

the confidenc

y the dashed l

ese uncertaint

to the I-bit 

e a lower 

to 10-percent

ncertainty is n

ent, is typical

are a more 

e indicates the 

B) best case th

s not 

o a 

n the 

e, a 

he 

ws 

ging 

s. 

178 to 

inty 

lbf

r. 

ature 

minal 

iling 

ce 

line 

ties. 

t

not an 

lly 

hrust 
NASA/TM

Fig

s

6.2 19BIm

A tota

a non-igni

No conclu

spark data

result, it w

flow. Furt

With resp

respect to

over a wid

cycles fro

The fo

were used

changed t

the colum

but a low 

discussion

between d

Since

Glenn foc

Aerojet (R

4 lbf-s (17

pulses we

M—2012-21761

gure 13: I-bit un

shows uncertai

mpulse Bit T

al of 81 tests w

ition event, as

usive reason f

a (discharge v

was not possib

ther investiga

pect to the maj

 pulse width (

de variety of p

om 0 to 100 pe

focus of Glenn

d in the major

to a 25-percen

mbium chambe

duty cycle w

n of the coatin

duty cycle and

 there was a p

cused on this E

Ref. 5) showe

7.8 N-s). Sinc

ere conducted

13 

ncertainty for th

inty using a 90-

Testing Res

were conduct

s confirmed b

for the non-ig

voltage or cur

ble to definiti

ation of ignitio

jor objectives

(EPW) and du

pulse widths 

ercent, as def

n testing was 

rity of the test

nt and then a 5

er. It is unclea

was retained to

ng degradatio

d performanc

programmatic

EPW, althoug

ed that an EPW

ce this require

 at the 40-ms 

he pulse trains 

-percent confid

sults 

ted during the

by a lack of te

gnitions was e

rrent, spark pu

ively diagnose

on and spark 

s of the test se

uty cycle. Th

and duty cycl

fined in object

minimum I-b

ts. Testing ori

5-percent dut

ar whether th

o match previo

on is given in 

e, as shown in

c requirement

gh wider and 

W of 40 ms w

ed the ACS fa

pulse widths

15 

using a 95-per

dence band for 

e pulse test se

emperature ris

evident from t

ulses, etc.) wa

e the conditio

behavior is p

eries, Table IV

e chart shows

les, illustratin

tive (4) (see S

bit performanc

iginally starte

ty cycle after 

e duty cycle w

ous sea-level 

Section 6.4.5

n Figure 14.

t for a minimu

narrower pul

was required t

acility aborts t

s. Fewer pulse

rcent confidenc

r the nominal te

eries. Of those

se observed in

the flow cond

as not availab

on of spark en

lanned in a fu

V shows a tab

s that the eng

ng that it can o

Section 5.2 ab

ce, so narrow

ed at a 50-per

some coating

was a definiti

testing by Ae

5. However, n

um EPW of 8

lse widths we

to reach the m

to be turned o

es would limi

ce band. The d

emperature run

e 81 tests, onl

n igniter cavi

ditions or data

ble during this

nergy being d

uture test seri

ble of the test

ine has been 

operate over 

bove). 

w EPWs and s

rcent duty cyc

g degradation 

ive cause of th

erojet (Ref. 5

no correlation

80 ms, the ma

ere tested. Prio

minimum I-bit

off, only a lim

it the amount 

ashed line 

n. 

ly four resulte

ity (TCign). 

a, as individua

s test series. A

delivered to th

es. 

ted conditions

extensively te

the range of d

shorter duty cy

cle, but was 

was observed

he degradatio

). Further 

n was observe

ajority of testi

or testing by 

t (MIB) goal o

mited number 

of unburned

ed in 

al 

As a 

he 

s with 

ested 

duty 

ycles 

d on 

on, 

d 

ing at 

of 

of 
NASA/TM

methane t

vacuum Itemperatu

temperatu

tests that c

I-bit. This

to increas

M—2012-21761

that could bui

-bit versus pu

ure at the LO2

ure, with cold

consisted of 3

s is opposite o

es in specific

13 

TABLE IV: 

(AEROJET 

EPW, 

ms 5

40 

50 

60 X

80 

100 X

120 X

160 X

320 X

500 

960 

1000 

F

ild up in the te

ulse number fo

2 (A) and LCH

er propellants

30 pulses. The

of the trends o

 impulse and 

EXAMINED E

AND DUTY C

[SEA-LEVEL] =

5 10 15

 X

 X

X X

 X X

X X

X 

X X X

X X

 

 

 

Figure 14: Plot 

est capsule in

for 80-ms EPW

H4 (B) valve i

s having high

e trend shows

observed in sp

efficiency (R

16 

ELECTRONIC P

YCLE COMBIN

= X; NASA GLE

Duty cycle, per

5 25 50

X X 

X X 

X X 

X 

X O O

X X 

X X 

O

X X

X O

of I-bit versus 

n the event of 

W tests. The s

inlets. There i

her I-bits. Figu

s that increasi

pecific impul

Refs. 6 and 7)

PULSE WIDTH 

NATIONS 

ENN = O; BOT

rcent

0 60 75

 

 

X 

 

O 

 

 

 

O 

X X

O 

duty cycle. 

f a non-ignitio

symbols are c

is a dependen

ure 16 also de

ing propellan

se testing, wh

. 

(EPW) 

TH = ) 

90 

X 

on. Figure 15 

color-coded b

ncy of I-bit on

emonstrates th

nt inlet temper

here increasin

shows a plot 

based on prope

n valve inlet 

his observatio

rature decreas

ng temperatur

of 

ellant 

on for 

ses 

re led 
NASA/TM

Figure 1

80-m

M—2012-21761

15: I-bit for eac

ms EPW runs w

Figure 16: 

13 

ch pulse with pr

ith first pulse te

I-bit as a funct

ropellant inlet te

emperatures w

ion of (A) LO2 a

17 

emperature at 

with ±10 R of th

and (B) LCH4 i

the (A) LO2 an

he target are sh

nlet temperatu

d (B) LCH4 thr

hown. 

ure for 30 pulse

ruster valves. O

 tests. 

nly 
NASA/TM

Two f

thrust, bot

(a) and pe

propellant

trapped li

pressure t

“double p

Figure 17

values we

exists, it s

67-N) ban

445 N), p

the cause 

average, n

testing (R

With 

consistent

with the d

EPW tests

data run s

conditions

All of the 

prior Aero

M—2012-21761

factors, howe

th of which m

eak thrust (b) 

ts, resulting in

quid in the m

takes longer to

pulse” phenom

(B), has a sim

ere based upon

should be note

nd, while osci

eak-to-peak. 

of the I-bit tr

nominal temp

Ref. 5), sugges

an EPW of 4

tly, for a varie

data points co

s, warm tests 

shows such a 

s. This set of 

other 40-ms 

ojet testing (R

Figure 17: 

13 

ever, play a ro

may be influen

versus the LO

n a higher I-b

manifold, whic

o decay, whic

mena in some 

milar trend as 

n the curve fi

ed that the ma

illations obser

Thus, it is dif

rend can be pr

perature I-bit v

sting I-bit calc

0 ms, the eng

ety of propell

lor mapped to

generally hav

significantly 

data clearly d

EPW runs de

Ref. 5). 

The effect of L

ole in determin

nced by prope

O2 valve inlet

bit. The corres

ch would cont

ch can be seen

tests, which w

integration ti

it profile desc

ajority of the 

rved in the th

fficult to claim

rimarily attrib

values were c

culation meth

gine could ach

lant inlet temp

o propellant in

ve lower I-bit

larger I-bit (≈

defies trends o

emonstrate a M

LO2 inlet tempe

18 

ning I-bit: (1)

ellant tempera

t temperature.

sponding cold

tinue to vapor

n in Figure 18

will be discus

ime. Due to th

cribed in Secti

peak thrust d

hrust stand we

m a significan

buted to integ

consistent with

hods are consi

hieve the MIB

peratures. Fig

nlet temperat

t than cold tes

≈40 percent) t

observed, and

MIB of 4 lbf-s

erature on (A) in

) integration t

ature. Figure 

. Integration t

der hardware 

rize after valv

8. This effect 

ssed in Sectio

he thrust stand

ion 6.1 (see F

data falls with

ere as great as

nt contribution

gration time. I

th values prev

istent. 

B requirement

gure 19 shows

tures. From th

sts. Thus, it is

than colder te

d is considere

s (17.8 N-s) o

ntegration time

time and (2) i

17 is a plot o

time is highes

temperatures

ve closure. Th

is also believ

on 6.4.3. Thru

d ringing, the

Figure 12). W

hin a ±10- to 1

s ±40 to 100 l

n to I-bit base

It should also 

viously report

t of 4 lbf-s (17

s a plot of 40-

he previously 

s not clear wh

ests at otherw

ed here as a st

or less, which

e and (B) peak 

integration of

of integration 

st with colder

s may lead to 

hus, chamber 

ved to cause a

ust, shown in 

ese peak thrus

While a trend 

15-lbf (±45- to

lbf (±178 to 

ed upon thrus

be noted that

ted in Aerojet

7.8 N-s) 

-ms EPW test

observed 80-

hy a single wa

ise similar 

tatistical outli

h is consistent

thrust. 

f

time 

r

a

st 

o

st, and 

t

t

ts 

-ms 

arm 

ier. 

t with 
NASA/TM

Figure 1

are s

M—2012-21761

18: Chamber p

shown in the gr

Figure 19: I-bi

indicated by

13 

pressure of sev

aph inset. 

t for each 40-m

y color. The do

veral pulses (ov

ms EPW pulse w

tted line indica

19 

verlaid) of a 30-

with (A) LO2 an

ates the minimu

-pulse train. I-b

nd (B) LCH4 pro

um impulse bit

bits for the corr

ropellant inlet te

(MIB) target of

responding puls

emperatures 

f 4 lbf-s. 

ses 
NASA/TM

6.3 20BC

Durin

Unison. T

40-ms EP

to other ig

distinguis

exciter (80

40-ms EP

is possible

operated o

6.4 21BA

Throu

or which w

stand ring

nozzle col

6.4.1 26BT

As me

thrust sign

system, su

illustrates

(FFT) of t

(Run 099)

dampened

smaller 45

frequencie

type test w

addition, t

oscillating

M—2012-21761

Compact Un

ng the pulse te

The igniter pro

PW, at a 5-per

gniters and wa

shed from the 

0-ms EPW) w

PW, non-ignit

e these were l

otherwise. Fig

Anomalies O

ughout the tes

were notable,

ging, thrust sta

lumbium coa

Thrust Stand 

entioned in th

nal during pu

uggesting the 

 thrust stand 

the data indic

) from prior te

d significantly

5 Hz oscillati

es of the thru

was conducte

this behavior 

g thrust, albei

13 

nison Excite

est series, the 

oved reliable 

rcent duty cyc

as not expecte

other test dat

were delayed 

ion events oc

late ignition e

gure 5 is a ph

Observed Du

st series, seve

, unexpected e

and zero drift

ting. 

Ringing 

he uncertainty

lses. This osc

oscillation w

oscillations d

cated these oc

esting (Ref. 7

y towards the 

on. Since the 

st stand and t

d on the thrus

was also seen

it for a differe

Figure 20: T

r 

engine was a

and capable o

cle for a train 

ed to change 

ta. The only a

ignition even

ccurred, but gi

events that fel

hotograph of t

uring Test S

ral anomalies

events. They 

t, double puls

y analysis, Se

cillation was n

was caused by

during pulses 

curred at app

7), similar osc

latter half of

oscillations d

the impulsive 

st stand to det

n in prior pul

ent serial num

Thrust stand rin

20 

also tested wit

of successfull

of 30 pulses 

I-bit perform

aborts observe

nts described i

iven the lengt

ll outside the 

the engine ope

Series 

s were observ

are discussed

ing, feed syst

ction 6.1, sign

not observed 

 the natural fr

from a single

roximately 45

cillations were

f a 5-s burn, w

dampen durin

nature of star

termine the n

se testing by 

mber injector. 

nging for two re

th the compac

ly igniting the

(at 80-ms EP

mance, thus the

ed during test

in Section 4.7

th of an igniti

EPW window

erating with t

ved that either

d here for com

tem temperatu

nificant oscill

in chamber p

frequency of t

e run (Run 24

5 and 100 Hz

e noticed dur

with the shutd

ng the burn, th

rtup transient

natural ring-ou

Aerojet (Ref.

A rough estim

epresentative p

ct exciter unit

e engine throu

PW). Perform

e Unison data

ting with the 

7.1 above. Du

ion delay (see

w, given how

the compact U

r affected the 

mpleteness, an

ure rise, and d

lations were o

pressure traces

the thrust stan

3). A fast Fou

z. In Isp,vac perf

ing the startu

down transient

hey are likely

ts. To date, no

ut frequency o

. 5), who pres

mation of this

pulses. 

t developed b

ugh 80- and 

ance was sim

a was not 

Unison comp

uring testing o

e Section 4.7.

w reliably the e

Unison excite

quality of res

nd include thr

damage to the

observed in th

s (Pcav) or the

nd itself. Figu

urier transform

formance data

up transient. T

t showing onl

y due to natura

o impact ham

of the test stan

sented a plot o

s frequency 

by 

milar 

pact 

of the 

1), it 

exciter 

er. 

sults 

rust 

e

he 

e feed 

ure 20 

m

a

These 

ly the 

al 

mmernd. In 

of the 
NASA/TM

suggests a

injector se

pulse run 

to filter pu

compariso

6.4.2 27BT

Over 

approxim

effects. Th

loads, as s

loads due 

contracted

load cells

relax, retu

(thrust val

the overal

M—2012-21761

a 100-Hz com

erial number 

discussed in 

ulse data for t

on, noting the

Thrust Stand 

the course of

ately 4 to 6 lb

he injector as

seen in Figure

to the chill-in

d, thus pulling

, resulting in 

urning the eng

lues were rez

ll trends are m

Fig

z

13 

mponent. An F

indicated a 50

Figure 20 (Ru

thrust stand ri

e oscillation s

Zero Drift 

f longer (>20)

bf (17.8 to 26.

ssembly is hel

e 22. During t

n of the engin

g on the engin

the zero offse

gine to the ori

zeroed betwee

maintained (F

gure 21: Thrust

zero drift towar

Figure 

coup

FFT of some u

0-Hz oscillati

un 243). Sinc

inging, it was

imilarities be

) pulse trains, 

.7 N) (Figure 

ld in place by

testing, the en

ne. These load

ne assembly. 

et drift. As the

iginal zero. Ien each pulse)

igure 23). 

t measurement

rds end of train

22: Photograp

pler (indicated)

21 

unpublished A

ion, during a p

ce Aerojet did

s decided to p

tween data se

the zero offs

21). The cau

y a threaded co

ngine mountin

ds could have

This would h

e assembly w

-bit calculatio

). The drift alt

ts over a 30-pu

n (inset). 

h of engine sho

used to reduce

Aerojet data (

pulse test und

d not make an

present the cur

ets. 

et for the thru

use for this is a

oupler, which

ng bracket ex

e translated to

have changed 

warmed up bet

ons were adju

ltered I-bit res

ulse train, illustr

owing threaded

e cantilever loa

(Ref. 17) for 

der similar co

ny indication (

rrent data for 

ust levels shif

attributed to t

h reduces som

xperienced sig

o the threaded

the force dist

tween tests, th

sted to accou

sults by ≤0.5 

rating thrust 

d

ads.

this specific 

onditions to th

(Ref. 5) of eff

direct 

fted down by 

thermal expan

me of the cant

gnificant therm

d coupler, whi

tribution on th

his load woul

unt for this eff

lbf-s (2.2 N-s

he 

fforts 

nsion 

tilever 

mal 

ich 

he 

ld 

fect 

s), but 
NASA/TM

6.4.3 28BD

A uni

as indicate

shows a s

image 7 th

representa

Because t

correlated

The e

pulses in a

for all the

pressure p

The press

higher I-b

when the 

indicator, 

One p

manifold 

remains c

over to a s

oxygen m

value betw

closed. Si

Thus, a fl

noted that

temperatu

manifold 

the manifo

M—2012-21761

Double Pulse 

ique phenome

ed visibly by 

equence of vi

he flame rega

ative of the be

the video fram

d with the 100

effect on perfo

a test at nomi

e pulses. The I

plot, where th

ure decay is s

bit. However, 

second pulse 

but this also 

possible cause

for both fuel 

onsistent thro

saturated state

manifold press

ween pulses. T

ince the metha

ame could co

t the double p

ure is only a p

system the te

fold instrumen

13 

Figure 23: Effe

Phenomenon

enon was obse

a second flam

ideo images f

ains strength. 

ehavior. Seco

me rate is only

00-Hz data. 

ormance can b

inal propellan

I-bits are high

he dotted line 

slower when a

the chamber 

begins or end

occurs for pu

e of the doubl

and oxidizer 

oughout the te

e in the highl

sure. When th

This could be

ane valve clo

ontinue, or rei

pulse region d

point measure

mperature ma

nts. 

ect of thrust ze

n

erved whereb

me pulse. Thi

from a burn. T

While the du

ondary pulses 

y 30 Hz, the v

be seen in Fig

nt temperature

hest during th

at 17.5 psia (

a double puls

pressure data

ds. For examp

ulse 29, when 

le pulse is sug

are shown in 

est, but at 12 

ighted region

here is liquid i

e a result of L

ses 40 ms aft

gnite, after th

does not exact

ement taken at

ay be such tha

22 

ero drift on indiv

by the pulse re

s double puls

The nominal, 

uration of the s

are not fully 

visualization o

gure 25, wher

es are overlaid

he double puls

0.12 MPa) in

se occurs, lead

a fails to revea

ple, the inflec

no double pu

ggested in Fig

Figure 26(A)

s (pulse 8), th

n of the graph

in the manifol

LO2 vaporizing

er the oxygen

he oxygen val

tly correlate w

t the entrance

at propellant 

vidual pulse I-b

egained stren

se is best desc

main pulse is

secondary pu

expanded an

of the phenom

re the chambe

d. The inset s

ses. The reaso

ndicates the en

ding to a long

al any feature

ction point at 

ulse was obse

gure 26. The p

). The methan

he oxygen tem

. Figure 26(B

ld, the pressu

g in the mani

n, both reactan

lve was comm

with the satura

e to the manif

exists in a dif

bit results. 

ngth after the v

cribed using F

s shown in im

ulse varies, thi

d have a mor

mena could no

er pressure tra

shows the cor

on is evident 

nd of the I-bit

ger integration

e that definitiv

around 0.15 s

erved. 

propellant tem

ne temperatur

mperature plu

B) shows the c

ure fails to retu

ifold after the

ants are presen

manded closed

ation region i

fold, and elsew

fferent state th

valves had clo

Figure 24, wh

mages 1 to 6, b

is figure is 

e disperse fla

ot be directly

aces for sever

rresponding Iin the chambe

t integration t

n time and thu

vely indicates

s may be an 

mperatures in

re behavior 

ummets, cross

corresponding

urn to its zero

 oxygen valv

nt in the cham

d. It should b

in Figure 26. 

where in the 

han indicated

osed, 

hich 

but in 

ame. 

ral 

-bits 

er 

time. 

us a 

s

n the 

sing 

g

o

e has 

mber. 

e 

This 

d by 
NASA/TM

Fig

g

l

M—2012-21761

gure 25: Chamb

graph inset. Th

onger integrati

13 

Figure 24: Im

phenomen

the second

ber pressure of

he pressure dec

on time for I-bi

mage sequence

na. The main p

dary double pu

f several pulses

cay for the dou

it. 

23 

e illustrating the

pulse is shown 

ulse begins in im

s (overlaid) of a

uble pulses (pu

e double pulse

in images 2 to 

mage 7. 

a 30-pulse trai

lses 15 and 20

6, and 

n with I-bits sh

) is slower, lea

own in the 

ding to a 
NASA/TM

Figur

indicated 

tests. The 

temperatu

pulse phen

showed d

in the man

as Run 28

infrequen

state in th

experienc

Anoth

weaker af

early in th

propellant

displayed 

video fram

weak puls

propellant

between n

The in

image wa

downstrea

each video

measurem

luminositi

M—2012-21761

Figure 2

and (B

e 27 illustrate

in the title of

highlighted r

ure traces indi

nomenon is m

ouble pulsing

nifold for mu

85, which may

t during tests 

he manifold, th

cing different 

her feature wa

fter the double

he train (prior

t tests and cou

a visible flam

mes for each p

ses lasted two

t flow was ac

normal and w

ntensity of the

as converted to

am of the noz

o frame of the

ments were do

ies from Run 

13 

6: Saturation in

B) LO2 manifold

es the effect o

f each plot. Th

region is whe

icate saturatio

most prominen

g. This is supp

ch of these te

y suggest a tw

at warm prop

he temperatur

bulk conditio

as observed in

e pulsing had

r to the double

uld be quantif

me, and secon

pulse revealed

o to three fram

tive for 80 m

weak pulses is 

e flame was q

o 8-bit gray s

zzle exit, and 

e main pulse 

one for the sec

265 at warm 

n the LO2 mani

d pressure. Do

of the propella

he oxygen ma

re double pul

on (calculated

nt in the cold

ported by the 

ests. There is f

wo-phase, or b

pellant condit

re shifts are m

ons than the lo

n conjunction

d subsided. Vi

e pulsing). Th

fied in two w

nd, by the inte

d that most st

mes. However

ms. Therefore,

not clear by f

quantified by 

cale. A line p

the mean lum

flame, and th

condary flame

propellant co

24 

ifold is indicate

uble pulses (bo

ant temperatu

anifold tempe

lsing was obs

d using the ma

d and nominal

temperature t

fluctuation be

borderline con

tions. While t

more severe, p

ocalized therm

n with the dou

isually, these

his phenomen

ways using the

ensity of the f

tandard pulse

r, the video fr

given the tem

frame count a

measuring th

profile tool (R

minosity acros

he values were

e. Figure 28 (

onditions. In t

ed in the (A) ma

ox) overlap this

ure conditions

erature is show

erved. The da

anifold pressu

l tests, where 

traces that ind

etween liquid

ndition. In co

the data sugge

perhaps indic

mocouple ind

uble pulsing; t

appeared to b

non was most 

e video: first, b

flame while p

s lasted appro

rame rate was

mporal resolut

alone. 

he luminosity 

Ref. 18) was p

ss the line wa

e summed. Fo

(right axis) sh

this test, the l

anifold tempera

s saturation reg

s on this phen

wn for all nin

arker colored 

ure at each po

the majority 

dicate that the

d and gas for s

ontrast, double

ests the liquid

ating that the

dicates. 

the pulses beg

be shorter live

noted during

by the numbe

present. Count

oximately fou

s 30 Hz (33 m

tion of the vid

in each fram

positioned 10 

s determined.

or double pul

hows the main

luminosity see

ature 

gion. 

nomenon as 

ne 30-pulse tra

portions of th

oint). The dou

of the 30 puls

e oxygen is li

several tests, 

e pulses are 

d is in a satura

 propellant is

gan to grow 

ed than pulse

g the warm

er of frames w

ting the numb

ur frames whi

ms/frame) and

deo, the diffe

e of video. Th

pixels 

. This was do

lses, separate 

n pulse 

ems to have a

ain 

he 

uble 

ses 

iquid 

such 

ated 

s

s 

which 

ber of 

ile 

d the 

erence 

he 

one in 

a
NASA/TM

Figure 27

highlig

Figure 2

press

satur

M—2012-21761

7: LO2 manifold

ghted areas, an

28: The summe

sure traces ind

ration.) 

13 

d temperatures

nd the darker c

ed luminosity fr

icate gas in the

s are plotted for

olor of the trac

rom each pulse

e manifold at p

25 

r each 30-pulse

e indicates the

e video frame i

ulse 27. (Note 

e train run. Dou

e temperature is

ndicates the w

that a measure

uble pulses we

s below the sa

weak pulse phe

ed luminosity o

ere observed in

turation condit

nomena. The 

of 100 is white 

n the 

ion. 
NASA/TM

cyclic fea

manifold 

represent 

measured 

valve pres

weak puls

6.4.4 29BT

Since

the locatio

valve. Thi

propellant

performan

Figure 30

fluctuatio

shallow sl

the locatio

others, an

temperatu

be due to 

(Refs. 6 a

removed a

While

factor in t

engine no

such as th

Secondly,

experienc

there is no

Figure 

M—2012-21761

ature, where th

(Figure 27). T

the oxygen v

valve temper

ssure meets th

ses. 

Temperature 

 the weak pul

on of the heat

is warming tr

ts are flowing

nce is represe

 with the loca

ns before stea

lopes during t

on between th

nd the increase

ure and valve 

recirculation 

and 7), observ

after the pulse

e engine heat 

the temperatu

ozzle. If soak b

he valves, sho

, in the later p

cing as much h

o indication fr

29: Propellant 

13 

he non-monot

The last nine 

valve pressure

ratures). In th

he saturation 

Rise in Feed

lse phenomen

t source. As s

rend is counte

g through the 

ented in Figur

ations indicate

adily rising. T

the valve inst

he two turbine

e occurs here 

temperature.)

or poor rotati

ved pressure tr

e testing, but 

soak back ha

ure rise for two

back was a fa

uld see heat r

pulses, the eng

heating from 

from the temp

valve inlet tem

tonic behavio

pulses show a

e and the corre

he last four pu

curve. This su

d System 

non seems to b

hown in Figu

erintuitive; on

system. (Not

re 16.) Tempe

ed in the sche

The two upstr

tability, and th

e flow meters

first. (The ve

) Thus, it app

ion of the turb

ransients wer

subsequent te

as not been de

o reasons. Fir

actor, compon

rise first as ex

gine throat do

combustion (

perature plots 

mperatures for e

26 

or at pulses 10

a weakening 

esponding sat

ulses, where th

uggests transi

be related to p

ure 29, the wa

ne would expe

te that the effe

erature measu

ematic. The v

ream locations

hen begin to r

s is unique. Th

ertical lines in

ears that the h

bine flow me

e also traced 

ests for valida

efinitively rule

rst, the flow m

nents in direct

xpected from 

oes not radiate

(either from c

that the heati

each pulse indi

0 and 20 corre

flame. The ot

turation press

he luminosity

ition to gas in

propellant wa

arming trend i

ect the hardw

ect of this tem

urements alon

valve temperat

s, venturi and

rise with the v

he temperatur

ndicate the slo

heat rise initi

eter. During p

to this locatio

ation of this th

ed out as a co

meters are not

t line of sight

radiative or c

e as brightly 

cooler flame o

ing of the line

icate a tempera

espond to pha

ther traces on

sure (calculate

y decrease is m

n the valve ma

arming, the qu

is not isolated

ware to cool w

mperature rise

ng the inlet lin

atures show so

d flow meter 1

valve tempera

re is higher th

ope rise of thi

ates at this lo

revious speci

on. This flow

heory have no

ontributor, it i

t in direct line

t or connected

conductive pr

suggesting th

or better wall 

es has subside

ature rise durin

ase changes in

n this plot 

ed using the 

most notable,

ay be causing

uestion becom

d to the oxyge

while cryogeni

e on I-bit 

nes are plotted

ome initial 

1, show stead

ature. Howev

han any of the

is flow meter 

ocation. This m

ific impulse te

w meter was 

ot been done.

is an unlikely

e of sight of t

d to the engin

inciples. 

he chamber is 

cooling). Yet

ed. 

ng the pulse tra

n the 

, the 

g the 

mes 

en 

ic 

d in 

dy or 

ver, 

e

may 

esting 

. 

y

the 

ne, 

not 

t, 

ain. 
NASA/TM

Figure 30

in the 

Anoth

lines. The

condition 

that trace 

portion of

maintainin

propellant

flow mete

begin in th

against pr

6.4.5 30BC

Two c

some degr

inspection

shown in 

surface co

(Figure 31

oxidation 

chamber i

M—2012-21761

0: Feed system

above schema

her source of 

ese lines circu

the propellan

cooling may 

f the propellan

ng propellant

t during the d

ers were large

his vicinity. A

ropellant warm

Columbium C

columbium n

ree of damag

ns of the cham

Figure 31 for

oating was pre

1(B)), the pitt

of columbium

if not address

13 

m temperatures

atic. 

potential war

ulated liquid a

nt prior to a ru

have been on

nt could also 

t line tempera

dwell portions

e masses relat

Additional ins

ming. 

Coating Dam

ozzles were u

e to the inner

mber, which o

r both chambe

evalent throug

ting was isola

m (niobium), 

ed. 

s for (A) LO2 an

rming in the p

argon or liquid

un. This cooli

n or off for a g

be vented jus

ature. These v

s of zone 2 wo

tive to the rest

sulation was i

mage 

used during th

r chamber sur

occurred after

ers. For nozzl

ghout the com

ated only to a 

so a coating 

27 

nd (B) LCH4 line

propellant line

d nitrogen aro

ing system wa

given run. Us

st upstream of

valves were cl

ould be susce

t of the prope

installed follo

he course of th

face. The dam

r every test da

le SN–3 (Figu

mbustion cham

few small sp

failure could 

es. Correspond

es was the tra

ound the prop

as manually a

sing bleed val

f the thruster 

losed for zone

eptible to line 

ellant lines, it 

owing this tes

his test progr

mage was rev

ay. The result

ure 31(A)), pi

mber. In the s

ots. The coati

ultimately lea

ding sensor loc

ace cooling an

pellant lines to

adjusted to m

lves in the pro

valves as a m

e 2, and thus t

warming. Sin

is likely that 

st series as a f

ram, both of w

vealed in routi

ting boroscop

itting and spa

second nozzle

ing is meant t

ad to mechan

cations are sho

nd propellant 

o thermally 

meet such test n

opellant lines

means of 

the non-flowi

nce the turbin

heating woul

further precau

which experie

ine visual 

pe images are 

alling of the 

e, SN–1 

to prevent the

nical failure o

own 

bleed 

needs 

, a 

ing 

ne 

ld 

ution 

enced 

e

f the 
NASA/TM

Figure 3

dama

The fi

temperatu

burn time

chamber i

the damag

(Figure 32

The d

benign tes

than seven

temperatu

three prop

indication

Even 

Table V s

(Aerojet) 

the contra

damage, s

The m

breakdow

heat and o

M—2012-21761

32: Sparks app

age. 

first chamber, 

ure conditions

. (Prior to pul

inspection can

ge. However, 

2). 

damaged SN–

st set was cho

n tests per day

ures. The cham

pellant temper

ns in the plum

though the da

ummarizes so

and materials

actor’s facility

so many of th

most probable

wn of the R512

oxygen, a pro

13 

Figure 31: Co

pearing in the p

SN–3, was d

s. This test da

lse testing the

n only be don

sparks were 

–3 chamber w

osen: a 25-per

y. Minor dam

mber had und

rature conditi

me. The nozzle

amage to SN–

ome of the id

s experts at G

y as well as N

hese ideas rela

e cause of the 

2E silicate co

tective silica 

ating damage 

plume during tw

damaged durin

ay consisted o

e nozzle had u

ne at the end o

observed in th

as then replac

rcent duty cyc

mage was foun

dergone a tota

ions. Unlike t

e itself had no

–1 was minor

eas that were 

Glenn (Ref. 19

NASA White S

ate to the diffe

damage, as s

ating in a ma

layer is form

28 

on (A) SN–3 a

wo tests with th

ng the first pu

of 95 pulses at

undergone 13

of the day, it i

he final two t

ced with a pre

cle with no m

nd on the third

al of 55 pulses

the previous o

ot yet been di

r, testing was 

explored afte

9). Earlier test

Sands Test Fa

erences in the

uggested by t

atter similar to

ed, and volati

nd (B) SN–1 c

he SN–3 chamb

ulse test day t

t a 50-percent

31 s of specifi

is difficult to 

tests, at the en

eviously unus

more than five 

d test day tha

s with 5 s of a

occurrence, th

iscolored by h

halted until a

er consulting 

ting of identic

acility (WSTF

e test conditio

the materials 

o “pesting.” A

ilizes. If the o

hambers. 

ber are likely re

that used prop

t duty cycle to

ic impulse tes

pinpoint whi

nd of the 20-p

sed nozzle, th

pulses per te

at used cold pr

accumulated b

here were no s

heat. 

a cause could 

with the engi

cal nozzles by

F)) did not ex

ons. 

experts, is a c

As the coating

oxygen or tem

elated to the co

pellants at nom

otaling 38 s o

sting.) Since t

ich test caused

pulse trains 

he SN–1. A m

est and no mor

ropellant 

burn time at a

sparks or othe

be identified

ine manufactu

y Aerojet (bot

xperience this

chemical 

g is exposed to

mperature 

oating 

minal 

of 

the 

d 

more 

re 

all 

er 

d. 

urer 

th at 

s

o
NASA/TM—2012-217613 29 

TABLE V: POTENTIAL CAUSES OF DAMAGE TO COLUMBIUM CHAMBER 

Potential cause Proof Result 

Nozzles used here were newer than 

ones used in previous testing. The 

coating batch on the new nozzles 

could have been flawed. 

Test the old nozzle in the current 

facility in an attempt to cause the 

same damage. 

Nozzle had been damaged by other means in 

previous testing and could not be reused. 

Operating temperature exceeds the 

rated coating temperature. 

No direct chamber temperature 

diagnostics are available; 

however, the color of the nozzle 

during burn is an indicator. 

Nozzles discolor after heat exposure, but SN–1 had 

not yet been discolored. Yet the damage still 

occurred. Also, during duration testing, much 

higher temperatures were reached with no damage. 

Molten material from a degrading of 

the spark plug damaged the nozzle. 

Spark plug was removed and 

investigated. 

No damage to the spark plug was found. No 

damage was observed in the igniter cavity nor at 

the converging section of the nozzle. 

High local mixture ratios (MRs) near 

the wall were creating hot spots. 

Igniter pins inspected for 

alignment. Rough calculations 

were done to determine local MR

at the injector. 

Pins are inspected as standard leak check 

procedures and were never found out of alignment. 

Calculations indicated a lower MR at the injector. 

Propellant flow timing could be 

causing hard starts or oxidation on 

the walls (if directly exposed to 

LO2). 

Timing was reviewed. 

The oxygen valve was shut first so 

that the hot walls would not be 

exposed directly to oxygen. 

Aerojet testing used a variety of timing conditions 

including oxygen flow on shutdown. 

environment is insufficient, the rate of volatilization will exceed the formation, and the protective layer 

will fail. The oxygen may then leach into the coating and cause internal damage (scale formation). Small 

samples of the damaged coating were scraped off the chamber and chemically analyzed (x ray and 

scanning electron microscope (SEM)). The results (see Appendix B) indicated the presence of niobium 

oxide indicating oxidation of the substrate, but lacked the silica that would be indicative of the protective 

coating. The timing of the propellants had been oriented to avoid concern over too much oxygen near the 

hot walls, but this may have caused oxygen starvation of the coating. This, combined with the transient 

thermal conditions of a high duty cycle, could have led to break down of the coating. However, 

destructive testing and a more detailed materials study would be required to confirm this theory. 

Various solutions were discussed including preconditioning of the nozzles in an oxide atmosphere or 

application of a barrier coating. However, the path forward involved a simple change in test conditions in 

an attempt to avoid or reduce the occurrence of the pesting phenomenon. After weighing the risks, testing 

was resumed with the SN–1 chamber (since damage was minimal) with the following changes: 

(1) The duty cycle was decreased to 5 percent. The majority of the previous Aerojet testing had been 

done at this condition. This would also help mitigate the transients in the chamber, allowing more 

time between pulses for gas dispersion. 

(2) The lower duty cycle also allows more time for the coating and substrate to reach thermal 

equilibrium between pulses. A thermal mismatch between the coating and substrate may also 

have contributed to the damage. 

(3) The existing damage to the nozzle was well documented. When it was reinstalled, the nozzle was 

rotated 90 to see if the damage was related to the injector pattern (i.e., flow irregularities). 

Following the above changes, chamber SN–1 was tested for over 700 pulses with videoscope 

inspections performed at the end of each test day. No further degradation of the coating was observed 

beyond expected normal wear. Since the nature of anomaly is beyond the scope of the testing program, 

and since no further damage was observed, it was concluded not to pursue a more definitive explanation 

of cause at this time. 

 
NASA/TM—2012-217613 30 

7.0 6BSummary and Conclusions 

The 100-lbf liquid oxygen-liquid methane (LO2-LCH4) reaction control engine (RCE) developed by 

Aerojet was tested to determine impulse bit (I-bit) performance. All experiments took place in a reduced 

pressure environment using the Altitude Combustion Stand at the NASA Glenn Research Center. 

Propellant temperature was varied between three set point conditions to simulate operation limits of an 

in-space system. Pulse train length (number of consecutive pulses), pulse duration (electronic pulse width 

(EPW)), and duty cycle were all adjusted to meet test needs. 

The performance fully met three of the four objectives set forth: It was operated at, and below, the 

required minimum pulse duration of 80 ms; it achieved a minimum impulse bit (MIB) of less than 4 lbf-s 

(at 40-ms pulse duration); and it was operated at a range of duty cycle conditions. The fourth requirement 

stated that the MIB repeatability should be within ±5 percent after achieving stable temperature. 

However, stable temperatures were not achieved due to warming in the lines, believed to be associated 

with one of the turbine flow meters, though this has yet to be confirmed. This was a facility issue and not 

related to the performance of the engine. As a result, propellant temperatures increased gradually over the 

pulse train, which involved as many as 30 consecutive pulses. The I-bit values were shown to be strongly 

dependent on propellant inlet temperature with warmer propellants showing shorter I-bits. Thus the I-bit 

values dropped off toward the end of longer pulse trains. The number of consecutive pulses demonstrates 

reliable and repeatable ignition and pulse performance, bolstering the feasibility of this propellant 

combination for future exploration endeavors. 

Several issues with this particular engine system did occur. A unique phenomenon was observed 

whereby the engine reignited immediately after the valves were closed, resulting in a “double pulse.” 

Manifold temperature and pressure conditions indicate that this may have been due to a two-phase flow 

condition. Vaporization of liquid in the oxygen manifold could have reacted with the lagging methane 

flow to cause a short re-ignition. This double pulse phenomenon also led to higher I-bits (due to longer 

integration times). Another issue was damage to the chamber surface coating that occurred early in the 

pulse test program. Since the coating prevents oxidation of the columbium substrate, this posed a concern 

to the safety of the engine. Chemical analysis of some damaged coating material indicates this may have 

been caused by a phenomena likened to pesting. Insufficient oxygen at these temperature conditions 

prevented the formation of silicon oxide on the coating surface. The lack this protective layer caused 

internal damage to the coating (scaling). The thermal cycling of a high duty cycle could have contributed 

to this issue. Testing was resumed at lower duty cycle conditions (5 percent) with no further damage. 

8.0 7BRelated Work 

Testing with this engine continued with a focus on ignition phenomena. Since one of the perceived 

drawbacks of this propellant combination is high ignition energy, this program sought to determine 

energy limits. A well instrumented exciter unit was used to obtain high-fidelity voltage information of the 

spark discharge. By coupling this to the existing pressure instruments, it was possible to determine the 

energy of the ignition spark. A variable energy exciter unit was developed to explore these limits. 
NASA/TM—2012-217613 31 

8BReferences 

1. Melcher, J.C. and Allred, J.K., "Liquid Oxygen/Liquid Methane Testing of the RS-18 at NASA 

White Sands Test Facility", AIAA 2008-4843, 44th AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Hartford, CT, July 21–23, 2008. 

2. Melcher, J.C. and Allred, J.K., "Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar 

Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility," AIAA 2009–

4949, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Denver, CO, Aug. 2–

5, 2009. 

3. Over, A.P., Klem, M.K., and Motil, S.M., "A Successful Infusion Process for Enabling Lunar 

Exploration Technologies," AIAA 2007–6196, AIAA Space 2007 Conference and Exposition, Long 

Beach, CA, Sept. 18–20, 2007. 

4. Stone, R., Tiliakos, N., Balepin, V., Tsai, C.-Y., and Engers, R., "Altitude Testing of LOXMethane Rocket Engines at ATK GASL," AIAA 2008–3701, 26th AIAA Aerodynamics 

Measurement Technology and Ground Testing Conference, Seattle, WA, June 23–26, 2008. 

5. Robinson, P.J., Veith, E.M., Hurlbert, E.A., Jimenez, R., and Smith, T.D., "100-lbf LO2/LCH4 - 

Reaction Control Engine Technology Development for Future Space Vehicles", 59th International 

Astronautical Federation, Glasgow, Scotland, United Kingdom, September 29–October 3, 2008. 

6. Marshall, W.M. and Kleinhenz, J.E., "Hot-Fire Testing of 100 lbf LOX/LCH4 Reaction Control 

Engine at Altitude Conditions", JANNAF 57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee 

Meeting, Colorado Springs, CO, May 3–7, 2010. 

7. Marshall, W.M. and Kleinhenz, J.E., Performance Analysis of Specific Impulse Tests of a 100-lbf 

(445-N) LO2/LCH4 Reaction Control Engine at Altitude Conditions, NASA/TM—2011-217131, 

National Aeronautics and Space Administration, Cleveland, OH, 2011. 

8. The Historic Rocket Engine Test Facility, Produced by: NASA GRC, Available from: 

http://retf.grc.nasa.gov/, Last Accessed 2011. 

9. Smith, T.A., Pavli, A.J., and Kacynski, K.J., "Comparison of Theoretical and Experimental Thrust 

Performance of a 1030:1 Area Ratio Rocket Nozzle at a Chamber Pressure of 2413 kN/m2

 (350 

psia)," AIAA–87–2069, AIAA/SAE/ASME/ASEE 23rd Joint Propulsion Conference, San Diego, CA, 

June 29– July 2, 1987. 

10. Pavli, A.J., Kacynski, K.J., and Smith, T.A., Experimental Thrust Performance of a High-AreaRatio Rocket Nozzle, NASA TP–2720, National Aeronautics and Space Administration, Cleveland, 

OH, 1987. 

11. DIAdem: Getting Started with DIAdem, National Instruments, Austin, TX, 2009. 

12. Grasl, S.J., Skaff, A.F., Nguyen, C., Schubert, J., and Arrington, L., "Liquid Methane/Liquid 

Oxygen Propellant Conditioning and Feed System (PCFS) Test Rigs—Preliminary Test Results," 

JANNAF 57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee Meeting, Colorado Springs, CO, 

May 3-7, 2010. 

13. Skaff, A., Grasl, S., Nguyen, C., Hockenberry, S., Schubert, J., Arrington, L., and Vasek, T., 

"Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs," 

JANNAF 3rd Spacecraft Propulsion System Joint Subcommittee Meeting, Orlando, FL, December 8-

12, 2008. 

14. Stiegemeier, B.R. and Marshall, W.M., "Sea-Level Testing of a 100 lbf LOX/Methane Reaction 

Control Engine", JANNAF 57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee Meeting,

Colorado Springs, CO, May 3–7, 2010. 

15. Lemmon, E.W., Huber, M.L., and Mclinden, M.O., NIST Standard Reference Database 23: 

Reference Fluid Thermodynamic and Transport Properties-REFPROP, Version 8.0, National 

Institute of Standards and Technology, Standard Reference Data Program, Gaithersburg, MD, 

2007. 
NASA/TM—2012-217613 32 

16. JANNAF Rocket Engine Performance Test Data Aquisition and Interpretation Manual, CPIA 

Publication 245, Chemical Propulsion Information Agency, Silver Spring, MD, 1975. 

17. Aerojet, Telecon, January 20, 2010. 

18. Klimek, R.B. and Wright, T.W., Spotlight-16 Image Analysis Software, Software Manual Version 

2004.4.27, 2004. 

19. Aerojet, Telecon, April 28, 2010. 
NASA/TM—2012-217613 33 

Appendix A31B.—Acronyms and Symbols 

The acronyms and symbols found in this report are presented here. 

Acronyms 

ACS Altitude Combustion Stand 

EPW electronic pulse width 

ETDP Exploration and Technology Development Program 

FFC fuel film cooling 

FFT fast Fourier transform 

JANNAF Joint Army, Navy, NASA, Air Force 

MIB minimum impulse bit (lbf-s (N-s)) 

MR mixture ratio 

NASA National Aeronautics and Space Administration 

NIST National Institute of Standards and Technology 

PCAD Propulsion and Cryogenic Advanced Development (project) 

PCFS propellant conditioning feed system 

PLC programmable logic controller 

RCE reaction control engine 

RETF Rocket Engine Test Facility 

RTD resistance temperature detector 

SEM scanning electron microscope 

SN serial number 

TFV Fuel temperature measured just upstream of thruster valves, °R (K) 

TOV Oxidizer temperature measured just upstream of thruster valves, °R (K) 

WSTF White Sands Test Facility 

Symbols 

A area, in.2

 (mm2) 

B bias error 

b uncertainty value 

Cd discharge coefficient 

c* characteristic exhaust velocity, ft/s (m/s) 

F thrust force, lbf (N) 

I-bit impulse bit (lbf-s (N-s)) 

Isp specific impulse, s 

L′ characteristic chamber length, in. (mm) 

m mass flow rate, lbm/s (kg/s) 

MR mixture ratio, oxidizer/fuel (O/F), lbm/s oxidizer / lbm/s fuel (kg/s oxidizer / kg/s fuel) 

P pressure, psia (MPa) 

࣪ precision uncertainty 

T time, s 

Q volumetric flow rate, gal/min (L/min) 

t factor from Student’s t-distribution 
NASA/TM—2012-217613 34 

Δ differential 

 fluid density, lbm/ft3

 (kg/m3) 

Subscripts 

amb ambient 

c chamber 

calc calculated uncertainty 

cali calibration uncertainty 

cav chamber acoustic cavity 

e nozzle exit 

F thrust force 

f fuel 

ign igniter 

inst instrument uncertainty 

meas measurement uncertainty 

o oxidizer 

therm thermal expansion effects 

throat throat

total total 

vac vacuum condition 

 
NASA/TM—2012-217613 35 

Appendix B32B.—X ray and Scanning Electron Microscope Results 

Small samples of the coating from one of the damaged nozzles, serial number (SN)–3, were provided 

to the NASA Glenn materials group for analysis. The nozzle was scraped in both damaged and 

undamaged areas for comparison. Kapton® tape (DuPont) was used to collect and contain the sample. The 

following information is the raw data that led to the conclusions provided by Dongming Zhu of the 

NASA Glenn materials group, who performed this analysis. 

X ray diffraction data was gathered using a D8 Advance Diffractometer using Cu K radiation. Data 

was also gathered on the D8 Discover Diffractometer (area detector) to check for grain size effects. Full 

even-intensity Debye rings were observed indicating good sampling statistics. Samples of light and dark 

Kapton tape were also run to determine the contribution of the tape to the background. The scan of the 

light tape corresponded well with both datasets. 

B.1 34BX ray Diffraction: Undamaged (Baseline) 

 

Position [°2Theta] (Copper (Cu))

10 20 30 40 50 60 70 80 90 100

Counts

0

10000

40000

90000

 Peak List

 Nb Si2; Hexagonal; 04-003-6241

 Fe0.998 O; Cubic; 04-002-3667
NASA/TM—2012-217613 36 

Phase ID sample: 

Chemical formula Compound name Crystal system Reference code SemiQuant [%]

Nb Si2 Niobium silicon Hexagonal 04–003–6241 93

Fe0.998 O Wüstite, syn Cubic 04–002–3667 7

Possible matches to single unidentified peaks (not shown in graphics): 

30.1° Fe3O4, Fe(Nb2O6), or Cr2FeO4

40.9° NbCr2

73.2° FeO 

B.2 35BX ray Diffraction: Damaged 

 

Position [°2Theta] (Copper (Cu))

10 20 30 40 50 60 70 80 90 100

Counts

0

10000

40000

90000

 Peak List

 Nb Si2; Hexagonal; 04-003-6241

 Nb O; Cubic; 00-042-1125

 Cr Nb4 Si3; Tetragonal; 00-054-0368
NASA/TM

Phase iden

Ch

Nb

Nb

Cr 

Possible m

10.9° a 

36.1° F

36.5° N

42.3° F

39.0° F

B.3 36BS

The b

indicating

neither ha

M—2012-21761

ntification sam

hemical formula 

b Si2

b O 

Nb4 Si3

matches to sin

host of zeolit

eO (different 

NbO 

eO (different 

e2O3

canning Ele

black spots in 

g the niobium

as oxidized. 

13 

mple: 

Compo

Niobium sili

Niobium oxi

Chromium n

ngle unidentif

te (SiO2) phas

lattice param

lattice param

ectron Micr

the image ind

substrate, an

ound name 

con 

ide 

niobium silicon 

fied peaks (no

ses 

meter than in t

meter than pos

roscope Und

dicate pore fo

nd a high Si fo

37 

Crystal sys

Hexagonal 

Cubic 

Tetragonal 

ot shown in gr

table) 

ssible at 36.1°

damaged 

ormation, whi

or the coating

stem Refer

04–0

00–0

00–0

raphics): 

° and table) 

ich is limited.

g. The low O (

rence code Se

003–6241 

042–1125 

054–0368 

. There is a hi

(oxygen) peak

emiQuant [%] 

----- 

----- 

----- 

igh Nb peak, 

k indicates that 
NASA/TM—2012-217613 38 
NASA/TM

B.4 37BS

Area 

substrate o

M—2012-21761

canning Ele

L shows the c

oxidation. 

13 

ectron Micr

coating with v

oscope Dam

ery little pro

39 

aged 

otective oxidation. Areas J and K show 

Area L 

evidence of 
NASA/TM—2012-217613 40 

Area K 

Area J 
NASA/TM—2012-217613 41 

Appendix C3B.—Test Data Log 

The following tables list all tests that were performed as part of this study. The measurements are 

averages taken over the I-bit integration time. (Note that for high duty cycle (e.g., 50 percent) the 

chamber pressure never decays to 17.5 psi, which is the end point of the I-bit integration defined in 

Section 5.1. Therefore, the integration time ends when the next pulse begins). The information is 

abbreviated with respect to JANNAF standards to facilitate easier viewing of the parameters referenced in 

this document. The full table is available upon request. 

Column descriptions: 

(1) Run #: The number of the test run as referenced to the first facility tests in the 100-lbf RCE test 

program. 

(2) Test Description: The target conditions and/or purpose for the run is given. 

(3) All successful hot fire tests are indicated by their target propellant temperature (nominal, warm, 

or cold). 

(4) With the exception of a few tests, the flow rates were set to achieve a mixture ratio (MR) of 2.5 

based on the previous specific impulse test series. There are a few conditions designated “MR 

Low” which used the set point for an MR=2.0. 

(5) “Thrust Cal” indicates a calibration of the thrust load cell. All calibrations were performed at 

altitude conditions. These calibrations were performed at the beginning and end of each day 

(6) “Late ignition” indicates that for one of the pulses in a given spark train, ignition occurred late. 

This triggered an abort, as described in section 4.7.1, which caused an incomplete (aborted) 

pulse train. 

(7) “No-ignition” indicates that a given pulse in the train failed to ignite, resulting in an incomplete 

pulse train. 

(8) “Abort” indicates a facility abort. This could be triggered by a sensor reading out of limit. In 

some cases, this was due to an overly conservative limit set in the PLC (e.g. for the manifold 

pressures POJ and/or PFJ). 

(9) Pulse Duration: The time between ignition and the closing of the propellant valves for each 

pulse. This is also known as the electronic pulse width (EPW). 

(10) Number of Pulses achieved/planned: The first number specifies the number of successful pulses 

achieved, while the second number indicates the number of pulses intended for the pulse train. 

(11) Duty Cycle: This indicates the frequency of pulsing. It is the ratio of the pulse duration to the 

total between pulses (pulse start to pulse start). For example, a test with a 5-percent duty cycle 

and 80-ms pulse durations would mean there is 1600 ms between pulses (80 ms of hot fire and 

1520 ms of down time). 

(12) Run Tank Pressures: The pressure in the propellant holding tanks. This is a pressure-driven 

flow system, so these pressures govern the mass flow rate of the propellants, and thus the MR. 

(13) Run Tank Temperature: The temperature of the propellant in the holding tanks. The desired 

conditions (cold, nominal, and warm) were judged based on the temperature at the thruster 

valves. Thus the run tank temperatures were chosen to account for line losses. Note the 

temperatures are given in Fahrenheit, since this was the convention of the propellant 

conditioning systems (PFCS). 

(14) Thrust parameters: The thrust measurements from the load cell. The first column (Thrust, lbf) is 

a summation of the three loads, corrected using the calibration coefficient or “K factor”. The 

“vacuum thrust” is corrected to vacuum relative conditions. 

(15) Mass Flow: The mass flow rates and corresponding mixture ratio, or MR, of the propellants 

calculated at the venturi. During these short pulse times, these measurements are highly 

transient (Section 4.4.1) and these numbers are rough averages. 

(16) Chamber pressure: The pressure inside the combustion chamber, “Pcav” averaged. 
NASA/TM—2012-217613 42 

(17) Valve temperature: The temperature at the propellant thruster valve. TOV is the oxidizer side, 

TFV is fuel. 

(18) Test capsule pressure: The pressure in test cell. Therefore, the ‘ambient’ pressure of the engine. 

(19) Average I-bit: The I-bits for each individual pulse are averaged to obtain this number. 

(20) Note: Test specific notes. 
RCE ACS Test summary: Pulse testing

Run # Test description Pulse Duration, s

Number of 

pulses 

achieved / 

planned

Duty Cycle LOX CH4 LOX CH4 Thrust (lbf) Vac Thrust (lbf) K Factor LOX Venturi (lbm/s) CH4 Venturi (lbm/s) MR venturi P

Columbium Nozzle SN#3, Injector 6, Champion Exciter (contunation from steady state tests)

Interim activity: Prepare for pulse testing, software modified.

3/5/2010 Day Plan: Begin pulse test program at nominal conditions. Start with 1s duration and low number of pulses, gradually decrease duration and incorperate more pulses. Target is to get 80ms duration. 

End of Day: Very sucessful day.PCIGN broke at test 193. Took time to work out facility timing setting & did not have enough propellant to try for the 80ms case. Orange color in flame during 20s pulse... boroscope next d

190 Thrust cal191 Nominal 1 5/5 50% 325 265 43.93 47.11 1.028 0.121 0.054 2.24 

192 Nominal 1 5/5 50% 325 265 44.75 47.91 1.028 0.118 0.055 2.15 

193 Nominal 1 5/5 50% 325 265 44.01 47.33 1.028 0.112 0.056 1.99 

194 No ignition-Nominal 1 0/10 50% 325 265 -262 -265 1.028

195 Nominal 1 10/10 50% 325 265 -262 -266 44.04 46.96 1.028 0.111 0.057 1.97 196 Nominal 0.5 5/5 50% 325 265 -262 -264 45.62 49.55 1.028 0.120 0.058 2.06 

197 Nominal 0.5 5/5 50% 325 265 -262 -265 46.51 50.02 1.028 0.120 0.057 2.10 

198 Nominal 0.5 5/5 50% 325 265 -262 -262 46.06 49.21 1.028 0.117 0.058 2.01 

199 Facility abort- Nominal 0.1 0/5 50% 325 265 -261 -263 1.028

200 Nominal 0.1 5/5 50% 325 265 52.29 56.07 1.028 0.142 0.074 1.93 201 Facility abort- Nominal 0.1 0/10 50% 325 265 -262 -261 1.028

202 Nominal 0.1 10/10 50% 325 265 -260 -262 55.32 59.02 1.028 0.142 0.076 1.87 203 Facility abort- Nominal 0.1 0/20 50% 325 265 -260 -260 1.028

204 Nominal 0.1 20/20 50% 325 265 55.77 58.32 1.028 0.144 0.078 1.85 

205 Nominal 0.1 20/20 50% 325 265 54.55 57.59 1.028 0.139 0.076 1.83 

206 Thrust cal

NEW Columbium Nozzle SN#1

Interim Activity: The SN#3 nozzle was removed, inspection confirmed spalling seen with boroscope. Virgin spare Nozzle SN1 was put on the stand. Pcign transducer was replaced with the more rugged Tabor sensors.

3/12/2010 Day Plan: Run at very benign conditions since we don't know what caused damage. No more than 10pulses. Goal is to get some tests at 80ms

End of Day: Lost Pcign on 1st test. Late ignitions triggering aborts... cannot set abort limits any shorter. Methane lag diminishing as train proceeds... could this be a cause of spalling?

207 Thrust Cal208 Late ignition - Nominal 0.1 4/5 25% 325 265 -263 -257 35.33 39.59 1.028 0.119 0.065 1.82 

209 Nominal 0.08 1/1 25% 325 265 -264 -258 38.12 42.25 1.028 0.166 0.086 1.94 

210 Late ignition - Nominal 0.08 2/5 25% 325 265 -264 -257 38.17 41.86 1.028 0.151 0.071 2.13 211 Nominal 0.08 5/5 25% 325 265 -256 -264 37.87 41.80 1.028 0.139 0.067 2.07 

212 Nominal 0.1 5/5 25% 325 265 42.50 46.58 1.028 0.128 0.071 1.81 

213 Thrust cal

Interim Activity: Looked at data, still loosing methane lag. Don't know what is causing. Replace tabor pressure transducer and add higher range transducer to try to catch the problem

3/18/2010 Day Plan: Move to warm conditions, operate in 'safe' mode since we still don't know cause. Goal is to get SOME test for milestone progression and see effect of temp on lag

End of Day: There was difficultly with the Pcav abort, the window is so small at these durations that a late ignition (at the tail end of window) may mean Pcav has not met the abort condition. Easiest solution is to shift the ab

214 Thrust cal215 Late ignition - Warm 0.1 0/1 25% 370 265 -238 -237 216 Late ignition - Warm 0.1 4/5 25% 370 265 -238 -236 40.49 44.48 1.029 0.131 0.065 2.00 217 Warm 0.08 5/5 25% 370 265 -238 -235 38.48 42.65 1.029 0.136 0.064 2.11 218 Late ignition - Warm 0.08 2/5 25% 370 265 -238 -235 38.73 42.52 1.029 0.146 0.067 2.17 219 Late ignition - Warm 0.08 1/5 25% 370 265 -238 -235 30.91 33.77 1.029 0.159 0.076 2.10 220 Late ignition - Warm 0.08 3/5 25% 370 265 -238 -236 41.74 44.70 1.029 0.135 0.068 2.00 221 Warm 0.1 5/5 25% 370 265 -238 -236 40.86 44.38 1.029 0.125 0.066 1.91 

222 Late ignition - Warm 0.1 4/5 25% 370 265 -238 -235 43.52 46.80 1.029 0.130 0.066 1.97 223 Thrust cal

3/23/2010 Day Plan: Move to Cold conditions, still use relatively benign settings to avoid spalling. The methane recirculation pump broke prior to testing. Decided to proceed with test day even though methane will not meet cold cond

End of Day: Light spalling was found on post-test inspection. Stopped testing to explore potential causes more thoroughly

224 Thrust cal225 Late ignition - Cold 0.1 1/5 25% 325 285 -303 -280 27.45 30.80 1.032 0.118 0.064 1.84 

226 Abort -Cold 0.1 1/5 25% 325 285 -305 -280 39.24 42.27 1.032 0.123 0.066 1.87 227 Late ignition - Cold 0.1 2/5 25% 325 285 -305 -280 41.46 45.08 1.032 0.169 0.109 1.55 228 Abort -Cold 0.1 4/5 25% 325 285 -305 -280 43.97 47.46 1.032 0.165 0.089 1.86 

Run Tank Pressure, 

psig

Run Tank Temperature, 

Deg F Thrust paramaters Average Venturi Mass flow 

NASA/TM—2012-217613 
Test # Test description Pulse Duration, s

Number of 

pulses 

achieved / 

planned

Duty Cycle LOX CH4 LOX CH4 Thrust (lbf) Vac Thrust (lbf) K Factor LOX Venturi (lbm/s) CH4 Venturi (lbm/s) MR venturi P

229 Late ignition - Cold 0.08 0/5 25% 325 285 -305 -279 

230 Late ignition - Cold 0.08 2/5 25% 325 285 -305 -280 38.96 42.83 1.032 0.200 0.137 1.46 231 No ignition -Cold 0.08 0/5 25% 325 285 -305 -281 232 Abort - Cold 0.08 4/5 25% 325 285 -305 -278 41.57 45.64 1.032 0.146 0.087 1.68 

233 Thrust cal

Installed Unison Exciter

Interim activity: Anaysis of previous test data shows we need longer pulse trains to resolve pulse behavior. No definitive answer on the coating damage. Materials group analysis shows possible Pesting issue. Previously o

6/3/2010 Day Plan: Run at nominal conditions. Chamber has been clocked to monitor damage locations. Reduce Duty cycle to 5%, run longer pulse trains (10pulses).

End of Day: After facility abort resulting in a long duration (10s) test, we ended testing to check for damage. Good data from 1st two tests. Double pulse phenomena in run 235

234 Thrust cal235 Nominal 0.08 10/10 5% 325 285 33.94 38.89 1.019 0.161 0.086 1.88 

236 Nominal 0.08 10/10 5% 325 285 36.58 41.71 1.019 0.184 0.086 2.14 237 Faciity abort - Nominal 0.08 0/10 5% 325 285

Interim activity: Checked facility and engine. Facility had some heat damage from e-stop blow back. Engine had some discoloration on the bell... some blue that we haven't seen before. Boroscope does not indicate any sp

6/10/2010 Day Plan: Run Warm/Warm, 3 tests at 10pulse train.

End of Day: Got 3 good tests (good temp, no abort) at condition, and one more at higher temp. Did a timing test to see what would take to get a 40ms pulse duration. 

238 Thrust Cal239 Warm 0.08 10/10 5% 370 265 -239 -236 36.33 40.88 1.026 0.149 0.067 2.24 

240 Warm 0.08 10/10 5% 370 265 -238 -238 37.59 42.62 1.026 0.149 0.076 1.96 

241 Late ignition - Warm 0.08 6/10 5% 370 265 -238 -238 37.11 42.14 1.026 0.160 0.078 2.06 

242 Warm 0.08 10/10 5% 370 265 -238 -236 39.04 43.75 1.026 0.146 0.077 1.90 

243 Warm 0.08 10/10 5% 370 265 -238 -237 37.91 42.77 1.026 0.147 0.078 1.90 

244 Cold flow to test 40ms timing 0.04 10/10 5% 370 265 245 Thrust cal

Interim activity: Checked data and looked at double pulse phenomena. May be liquid collecting in manifold, then off gasing and reacting (autoignite) in the methane lag period

6/15/2010 Day Plan: Cold/Cold with 3 good tests of 10 pulse trains. Methane reciculation pump sill non-operational, so temperatures will likely not reach target

End of Day: First test was warm. Increased bleeds to get colder. Several aborts on POJ, reduced abort limit to 100psi to correct, then 2 good tests. Last test we reduced methane lag to see if effects double pulse.

246 Thrust Cal247 Cold 0.08 10/10 5% 325 285 -302 -280 41.39 46.52 1.027 0.171 0.083 2.07 

248 Cold 0.08 10/10 5% 325 285 -301 -280 42.48 48.04 1.027 0.160 0.083 1.92 249 Abort - Cold 0.08 7/10 5% 325 285 -301 -279 43.02 48.39 1.027 0.165 0.086 1.92 250 Abort - Cold 0.08 6/10 5% 325 285 -300 -278 43.46 48.64 1.027 0.169 0.088 1.93 251 Cold 0.08 10/10 5% 325 285 -301 -278 42.27 47.51 1.027 0.162 0.083 1.95 252 Cold 0.08 5/5 5% 325 285 38.61 43.93 1.027 0.183 0.088 2.09 

253 Thrust Cal

Interim activity: 

6/17/2010 Day Plan: Longer pulse trains and possibly 40ms pulse duration

End of Day: Successful day

254 Thrust Cal255 Late ignition - Nominal 0.08 12/15 5% 325 265 -262 -268 36.81 41.85 1.030 0.146 0.083 1.74 

256 Nominal 0.08 15/15 5% 325 265 -261 -265 37.40 42.92 1.030 0.137 0.080 1.70 

257 Nominal 0.08 20/20 5% 325 265 -260 -268 38.62 44.03 1.030 0.132 0.077 1.72 

258 Nominal 0.08 20/20 5% 325 265 -261 -269 38.92 44.22 1.030 0.132 0.078 1.70 

259 Nominal 0.08 30/30 5% 325 265 -260 -270 38.97 44.46 1.030 0.132 0.077 1.72 

260 Nominal 0.04 5/5 5% 325 265 -260 -268 25.22 31.15 1.030 0.151 0.081 1.87 261 Nominal 0.04 5/5 5% 325 265 -260 -271 24.81 30.44 1.030 0.155 0.083 1.88 262 Thrust cal

Compact Exciter Installed

Interim activity: Installed the compact exciter, Data system recalibrated, new data channels for compact exiciter data signals

7/7/2010 Day Plan: Warm/Warm with the compact exciter. Up to 30 pulse cycles. Nominal mixture ratio. 1 test at low MR at conclusion of day. 40 ms pulse train at end

End of Day: Exciter performed well, there were no dropouts or ignition failures. No issues encountered.

263 Thrust cal264 Warm 0.08 10/10 5% 370 265 -242 -239 32.44 38.04 1.038 0.145 0.071 2.03 

265 Warm 0.08 30/30 5% 370 265 -241 -238 35.63 41.35 1.038 0.132 0.071 1.86 

266 Warm 0.08 30/30 5% 370 265 -240 -238 36.51 42.05 1.038 0.124 0.072 1.73 

267 Warm 0.08 30/30 5% 370 265 -241 -237 36.48 42.17 1.038 0.127 0.072 1.77 

Run Tank Pressure, 

psig

Run Tank Temperature, 

Deg F Thrust paramaters Average Venturi Mass flow 

NASA/TM—2012-217613 
Test # Test description Pulse Duration, s

Number of 

pulses 

achieved / 

planned

Duty Cycle LOX CH4 LOX CH4 Thrust (lbf) Vac Thrust (lbf) K Factor LOX Venturi (lbm/s) CH4 Venturi (lbm/s) MR venturi P

268 Warm - Low MR 0.08 10/10 5% 330 265 -241 -247 35.24 41.30 1.038 0.136 0.082 1.66 

269 Warm - Low MR 0.08 10/10 5% 330 265 -240 -240 35.40 41.62 1.038 0.148 0.082 1.81 270 Warm 0.04 5/5 5% 370 265 -240 -240 27.40 33.83 1.038 0.157 0.076 2.07 

271 Warm 0.04 5/5 5% 370 265 -240 -240 29.45 36.04 1.038 0.146 0.076 1.92 

272 Thrust cal

Interim activity:

7/9/2010 Day Plan: Cold/Cold with the compact exciter. Up to 30 pulse cycles. Nominal mixture ratio. 1 test at low MR at conclusion of day. 40 ms pulse train at end

End of Day: Temperature conditions were too warm, but continued testing to prove out exciter. Test 279 almost out of Nitrogen, so we went to 40ms tests instead of repeating low MR. Double pulse observed in almost all 3

273 Thrust Cal274 Late ignition - Cold 0.08 6/10 5% 325 285 -302 -277 35.96 42.11 1.031 0.199 0.089 2.24 275 Cold 0.08 10/10 5% 325 285 -302 -278 37.31 43.83 1.031 0.153 0.080 1.90 

276 Cold 0.08 30/30 5% 325 285 -303 -278 37.51 43.55 1.031 0.147 0.073 2.03 277 Cold 0.08 10/10 5% 325 285 -302 -277 40.12 46.38 1.031 0.162 0.086 1.89 

278 Cold 0.08 30/30 5% 325 285 -302 -278 40.07 46.25 1.031 0.149 0.075 1.98 

279 Late ignition - Cold - Low MR 0.08 3/10 5% 290 285 -302 -277 34.44 40.76 1.031 0.195 0.115 1.70 

280 No ignition -Cold 0.04 4/5 5% 325 285 -302 -275 23.71 30.12 1.031 0.173 0.097 1.79 

281 No ignition -Cold 0.04 3/5 5% 325 285 25.24 31.51 1.031 0.179 0.100 1.80 

282 Thrust cal

Interim activity:

7/13/2010 Day Plan: Nominal with the compact exciter. Up to 30 pulse cycles. Nominal mixture ratio. 1 test at low MR and a few 40 ms pulse trains

End of Day: Traces left on during all tests to try to maintain temperature better (reduce line heating issues)

283 Thrust Cal284 Nominal 0.08 10/10 5% 325 265 -269 -265 34.70 37.99 1.025 0.177 0.077 2.30 

285 Nominal 0.08 30/30 5% 325 265 -266 -266 38.38 41.95 1.025 0.142 0.067 2.11 

286 Nominal 0.08 30/30 5% 325 265 -266 -265 40.56 43.95 1.025 0.142 0.069 2.06 

287 Late ignition -Nominal 0.08 20/30 5% 325 265 -265 -263 36.63 40.36 1.025 0.139 0.079 1.75 

288 Nominal 0.08 30/30 5% 325 265 -265 -265 38.23 41.94 1.025 0.135 0.073 1.84 289 Nominal 0.04 5/5 5% 325 265 -265 -264 25.56 29.75 1.025 0.149 0.080 1.88 290 Late ignition -Nominal - Low MR 0.08 3/30 5% 275 265 -265 -264 32.12 36.36 1.025 0.155 0.099 1.57 291 Nominal - Low MR 0.08 30 5% 275 265 -265 -264 30.95 34.71 1.025 0.117 0.077 1.52 292 Thrust cal

Run Tank Pressure, 

psig

Run Tank Temperature, 

Deg F Thrust paramaters Average Venturi Mass flow 

NASA/TM—2012-217613 
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4. TITLE AND SUBTITLE

Analysis of 100-lbf (445-N) LO2-LCH4 Reaction Control Engine Impulse Bit Performance 

5a. CONTRACT NUMBER

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5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Marshall, William, M.; Kleinhenz, Julie, E. 

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WBS 253225.04.02.01.06.01 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-18198 

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Subject Categories: 12, 20, and 28 

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This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Recently, liquid oxygen-liquid methane (LO2-LCH4) has been considered as a potential “green” propellant alternative for future exploration 

missions. The Propulsion and Cryogenic Advanced Development (PCAD) project was tasked by NASA to develop this propulsion 

combination to enable safe and cost-effective exploration missions. To date, limited experience with such combinations exist, and as a result 

a comprehensive test program is critical to demonstrating with the viability of implementing such a system. The NASA Glenn Research 

Center conducted a test program of a 100-lbf (445-N) reaction control engine (RCE) at the Center’s Altitude Combustion Stand (ACS), 

focusing on altitude testing over a wide variety of operational conditions. The ACS facility includes unique propellant conditioning feed 

systems (PCFS), which allow precise control of propellant inlet conditions to the engine. Engine performance as a result of these inlet 

conditions was examined extensively during the test program. This paper is a companion to the previous specific impulse testing paper, and 

discusses the pulsed-mode operation portion of testing, with a focus on minimum impulse bit (MIB) and repeatable pulse performance. The 

engine successfully demonstrated target MIB performance at all conditions, as well as successful demonstration of repeatable pulse widths. 

Some anomalous conditions experienced during testing are also discussed, including a double pulse phenomenon, which was not noted in 

previous test programs for this engine. 

15. SUBJECT TERMS

Methane; Liquid oxygen; Propellants; Vacuum; Specific impulse 

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