https://ntrs.nasa.gov/search.jsp?R=19850020759 2018-10-04T23:04:23+00:00Z

F i na 1 Sumnary Report.

Prepared By:

30 Apri l 1985

I G N I T I O N CHARACTERIZATION OF

LOX/HYDROCARBON PROPELLANTS

Contract NAS 9-16639

Prepared For

NASA - Lyqdon B. Johnson Space Center Houston, Texas

Project Enginser

Approved By:

Program Manager

Aero j e t Techsystems Company P.O. Box 13222

Sacramento, Cal i f o r n i a 95813

RPT/AA018@

FOREWORD

This f i n a l sumnary report i s provided i n p a r t i a l f u l f i l l m e n t o f Con-

t r a c t MA? 9-16639, I g n i t i o n Characterization o f LOX/Hydrocarbon Propellants. The program was conducted for the NASA - Johnson Space Center by the Aerojet

Techsystems Company under the cognizance of W. C. Boyd, technical monitor. A t

Aerojet, the project engineers were B. R. Lawver and K. Y. Wong, and the program managers were R. W. Michel and 0. C. Rousar. performed during the period 6 July 1982 t o 31 July 1984.

The technical work was

Signi f icant contr ibutions t o t h i s program were made by the fo l lowing

Aero j e t personnel :

8. J . Anderson Y. G. Hooper 8. 0. Bates J. W. Juchau 6. W. Cathroe A. R. K e l l e r R. L. Ewen H. H. Mueggenburg J. C. Fang J. L. Pieper R. S. Gross R. J. Pruett R. Hernandez W. A. Thompson

This report, sumnarizes the resu l t s of an evaluation o f the i g n i t i o n

character is t ics o f the GOX/Ethanol propel lant combination. i g n i t i o n char- acter izat ion was accomplished through the analysis, design, fabr ica t ion and

test ing o f a spark i n i t i a t e d torch i g n i t e r and prototype 620 1bF th rus ter / ig - n i t e r assembly. The i g n i t e r was tested over a chamber pressure range o f 3.3 t o 262 ps ia and a mixture r a t i o range of 0.4 t o 40. The prototype th rus t - e r / i gn i te r assembly was tested over the chamber pressure range o f 74 t o 197

ps ia and mixture r a t i o range o f 0.778 t o 3.29. Cold (-92" t o -165°F) and ambient (44" t o 80°F) propel lant temperatures were used.

Spark i g n i t e r i g n i t i o n l i m i t s and thruster steady-state and pulse mode,

performance. cool ing and s t a b i l i t y data are presented. Spark i g n i t e r i g n i t i o n l i m i t s are presented i n terms o f co ld f low pressure, i g n i t i o n chamber diameter

and mixture ra t i o . Thruster performance i s presented i n terms o f vacuum speci f ic impulse versus engine mixture ra t i o .

GOX/Ethanol propellants were shown t o be ign i tab le over a wide range o f

Cold propellants were shown t o have a minor e f f e c t on i g n i t e r mixture rat ios. i gn i t i on l imi ts . Thruster pulse mode capab i l i t y was demonstrated w i th mul- t i p l e pulses o f 0.08 sec duration and less.

i v

TABLE CJF CONTENTS

I. I nttoduct 1 on 11. Program Objectives

111. Method of Approach I V . Signif icant Results V. Limitations

V I . Recamended Follow-on Work VII. Conclusions Ref etences

Page 1 3 6 17 26 27 28 29

V

LIST OF TABLES

Table No. I Task I Igniter Test Variables I 1 Task I GOX/Ethanol Ignit ion and Durabi l i ty Test

Matrix

111 Prototype Thruster Test Data Summary

Page

12 13

22

v i

LIST OF FIGURES

Figure No. 1 2 3 4

5 6 7

a

9 10

11

GOX/Ethano\ Thruster Task I I gn i t i on Test Sequence

Task I I gn i te r and Test Hardware Assembly GOX/Ethanol

Task I Ign i t io r l Test Sequence

gni ter Test 281

- Steady State Testing

He Actuator and Cold-Flow Pressure

Thruster I g n i t i o n Sequence

Thruster Pulse Sequence - Ef fec t o f Chamber Diameter on I g n i t i o n

Ef fect o f Spark Energy and

Thruster Cf and Isp,; E t f i c

Thruster B i t Isp

Cold h e 1 on I g n i t i o n ency

Page 7

8 9 10 11 14 16 20

21

23 25

v i i

I. INTRODUCTION

Recent studies indicate that future space transportat ion system (STS)

engine development and operational recurring costs may be reduced through the use of low cost oxygen/hydrocarbon propellants in the auxiliary propulsion system (orbital maneuvering, reaction control, and vernier engines). The storable propellants (ritrogen tetroxide/monomethyl hydrazine) used on the current STS are toxic and considerably more expensive than oxygenlhydrocarbon propellants. These propel lants were selected over the oxygen/hydrocarbon propellants because a more extensive storable propellant thruster data base existed at that time. Also, the storable propellants are hypergolic and this eliminated the need for external ignition sources which were considered to be less re1 iable.

During the past five (5) years, several studies have been conducted for the purpose of expanding the oxygen/hydrocarbon thruster data base. Previous studies have examined the combustion performance and cooling characteristics o f LOX/RP-1, LOX/propane, LOX/methane and LOX/ethanol. The objective or this study was to characterize the ignition and pulse mode capability of promising oxygen/hydrocarbon propel lants for long-1 ife, reusable, STS application. GOX/ethanol propellants were selected for testing on the basis o f their demon- strated clean burning characteristics (Ref. 1) and because of the system advantages indicated for this propellant combination i n the study reported in Reference 2.

In this prcgram, a prototype 620 1bF GOX/ethanol RCE thruster with spark-initiated torch igniter was desigved, fabricated and tested over a wide range of inlet conditions. The prototype igniter was used to determine the fundamental spark ignition characteristics of GOX/ethanol propellants prior to thruster testing. These ignition characteristics included the effects af ijni ter chamber diameter, cold-f low pressure level (inlet pressure), spark energy, a- ’ inlet temperature. An integral thruster/igniter assembly was used to evaluate thruster pulse mode capability with the GOX/ethanol propellant

1

I, Introduction (cont.)

combination. range o f inlet conditions. with no evidence of combustion instability. with no evidence of soot or carbon deposition. demonstrated with pulse durations down to 40 msec.

The thruster was demonstrated to be ignitable over the whole All o f the tests demonstrate smooth combustion

The exhausr plumes were clean Pulse mode capability was

Two hundred five (205) igniter ignition tests were conducted cwering a mixture ratio range of 0.4 to 40. Ignition limits were defined in cerms o f

igniter mixture ratio, cold-flow pressure, spark energy and propellar,t tee- perature.

Eight-four (84) thruster tests were run covering a mixture ratio range o f 0.78 to 3.29 and a chamber pressure range o f 74 to 303 psia. pulse mode capability was demonstrated with pulse durations as short as 0.040 SCC .

Thruster

Two (2) additional igniter injectors were designed and fabricated for hydrocarbon fuel ignition testing at NASA/JSC with LOX/methane and LOX/propane propellants.

2

11. PROGRAM OBJECTIVES

The objectives of this program were to evaluate and characterize the

Basic ignition char- ignition of promising low-cost 1 iquid oxygen/hydrocarbon (LOX/HC) propellants for long-life reusable spacecraft propulsion systems. acteristics of the GOX/Ethanol propellant combination were generated and evoluated over a range o f operating conditions applicable to an auxiliary propulsion system (orbital maneuvering, reaction control, and vernier engines) for the Space Shuttle Orbiter. ciated with engine ignition for these propellants were identified and assessed. ated.

The operational and design limitations asso-

Information necessary for specific hardware designs was also gener-

The program consists of five parts; Task I - Igniter Experimental Evaluation; Task I 1 - Preliminary Design; WBS 8.0 - Added Scope Testing; Task 111 - Thruster Experimental Evaluation; and WBS 9.0 - Added Scope Hard- ware Oesign and Fabrication.

The objectives of Task I were to determine the fundamental spark igni- tion characteristics o f GOY/Ethanol propellants. These characterist'cs incltide the effects o f design and operating variables such as ign'ter chamber size, cold-flow pressure level, spark energy, igniter cooling requirements, pulse mcde capability and oxidizer manifa;3 fuel contamination. A secondary objective was to evaluate the carbon formation potential of fuel rich gas generator mixtures. This was accomplished by designing a gas generator cham- ber with an internal secondary injector to fit the igniter injector. Nine (9) tests were run over a mixture ratio range o f 0.189 to 0.441. carbon formation was observed. Detailed results are reported in the Task I data dump, Reference 3,

No evidence o f

The objective o f Task I 1 was to analytically evaluate performance and cooling requirements o f GOX/Ethanol igniters auxiliary propulsion application;. The results of these eva

the and uat

ignition, thrusters for ons were used

3

11, Program Objectives (cont.)

t o guide the Task I and I11 i g n i t e r and thruster design a c t i v i t i e s . These r e s u l t s are reported i n the Task I 1 data dump, Reference 4.

The object ive of the added scope tes t ing was t o get an eab evnlu; on

These tes ts were conducted using res idual hardware fro% the mid The r e s u l t s o f these evaluations were an a i d t o the

of combustion performance o f GOX/Et.hanol thrusters f o r auvi 1 i a r y propulsion

applications. Pc (NAS 9-15958) program. Task I11 thruster design a c t i v i t y . 8.0 data dump, Reference 5.

Complete d e t a i l s are presented i n the WBS

The primary ob ject ive of the Task I 1 1 t e s t i n g was t o experimental ly evaluate GOX/Ethanol thruster i g n i t i o n and pulse-mode operation. A secondary ob ject ive was t o evaluate th rus ter steady-state performance, cool ing and s t a b i l i t y character is t ics . in jec to r , and th rus t chamber was tested over a wide range o f propel lant i n l e t temperatures and pressures.

A prototype thr l rster consist ing o f an i g n i t e r ,

The tes t ing was conducted i n two parts. The f i r s t par t checked out the Nineteen (19) tes ts Task I GOX/Ethanol i g n i t e r i n the th rus ter t e s t f a c i l i t y .

were run t o evaluate the ign ' ter feed system, valve timing, and propel lant i n l e t condi t ion effects.

The second p a r t addressed f u l l th rus ter operat "m. S i x t y - f i v e (65) thruster tes ts were Zonducted. The f i r s t t h i r t e e n (13) t e s t s were run wi th a heat sink chamber t o evaluate i g n i t i o n and i n l e t condi t ion e f f e c t s on per for - mance. f i lm-coolant effect iveness and pulse mode performance over a range o f i n l e t conditions. Detai led data are presented i n the Task I11 data dump, Reference 6.

The remaining tests were run w i th a th in-wal l chamber t o evaluate

4

I I , Program Object i ves (cont . ) The object ive of the added scope design and fabr ica t ion t e s t s was t o

design and fabr ica te two (2) addit ional i g n i t e r in jec tors for i g n i t i o n tes t ing with LOX/Methane and LOX/Propane propellants. The i g n i t i o n tes t ing w i l i be

conducted a t NASA/JSC.

5

I 1 I. METHOD OF APPROACH

The prototype 620 1bF GOX/Ethanol RCS thruster with integral spark- initiated torch igniter, shown in Figures 1 and 2, !,as designed, fabricated and tested over a wide range of inlet conditions. in Figures 3 and 4 was used to determine the fundamental spark ignition char- acteristics o f GOX/Ethanol propellants prior to thruster testing. These ignition charatteribtics included the effects of igniter chamber diameter, cold-flow pressure level (inlet pressure), spark energy, and inl;t tempera- ture. ratio range of 0.4 to 40 and a cold-flow pressure range of 3.3 to 49 psia. Chamber diameters of 0.15, 0.20 and 0.30 were tested. Propellant temperatures ranged from 80" to -165°F.

The prototype igniter shown

Two hundred five (205) ignit. ,. hot fire tests were run over a mixture

The ignition tests were conducted using the start sequence shown in =igure 5. The oxygen flnw was started 10 msec ahead of the Ethanol flow. The spark discharge was delayed about 110 nsec to establish the cold-flow pres- sure. Ignition or non-ignition was indicated by the chamber pressure and exhaust temperature response. The total test duration was 0.4 seconds.

The gaseous oxygen inlet pressure was set to achieve the desired cham- ber cold-flow pressure a!d the fuel inlet pressure variad up or down to shift the mixture ratio. determine its effect on ignition limits.

The spark enerc_;y was varied at fixed inlet CGnditions to

The ignition test variables and the range o f conditions tested are listed in Table I. the ignition response (i.e., ignition or non-ignition) on the fltme quench parameter (PD) versus igniter core mixture ratio coordinates. quench parameter i s the product of the cold-flow pressure before ignition and the chamber diameter.

The effects of these variables zlere determined by plotting

The flame

6

#--

0 t tu A= c, w \ x 0 a

7

% n E aJ in ul

c

a aJ L Q 3 -0 L Q I c, v) aJ c U c

CLI

L a, c1

c n,

- H

H

Y in CLI c

m aJ L 7 m

LL. *r

9

!

W

v) v) W ot

t

!3 t

n ! ot i W 1 z I u

i

. , ! : , ' I

. I I . .

m al L 0,

LL

a .p

11

TABLE I

TASK I IGNITER TEST VARIABLES

0

0

0

0

0

0

0

Propellant Temperature O knb 44 t o 80°F

* Cold -92 t o -165OF

Cold Flow Pressure O 3.3 - 49 psia

Chamber Diameter O 0.15 in . O 0.30 in.

Mixture Ratio

O 0.4 - 40

Spark Energy O 10 - 50 mJ

Spark Rate O 300 SPS (Fixed)

Spark Gap O 0.025 in. (Fixed)

12

111, Method of Approach (cont.)

The integral thruster/igniter assembly was used to evaluate thruster pulse node capability with the GOX/Ethanol propellant cornbination. Thirteen heat sink chiunber and 52 thin wall chamber tests were run over the mixture ratio range of 0.778 to 3.29 and chamber pressure range of 74 to 197 psia. Propellant terperatures ranged from ambient to -165°F.

The thruster assembly was instrumented to measure thrust, propellant flow-rates, inlet pressures and temperatures, conbustor wall temperatures, and combustion pressures. These data were recorded on both digital and analog recording instruments .

Thruster tests were run to evaluate ignition, performance, thermal compatibility, stability and pulse m d e operation. The primary test variables were propellant inlet pressures and temperature. Secondary variables were film coolant injection scheme (tangential or swirl) and valve actuation gas (GNp or GHe). Most of the tests were run with ambient temperature propellants and the thin-wall chamber.

An oscillograph trace of a typical thruster start sequence for a steady-state firing is shown in Figure 6. The igniter fuel valve was signaled 10-20 milliseconds ahead of the igniter oxidizer valve to provide a momentary core fuel rich condition. The igniter-alone testing showed this to be a more reliable start condition than siimltaneous igniter valve actuation. The spark was signaled on with the fuel valve so that ignition would occur immediately upon introduction o f the oxidizer. The thruster bipropellant valve was signaled on about 45 milliseconds from the igniter fuel valve signal to permit time for the computer to test the igniter chamber p-essure. chamber pressure had not achieved a preselected level, then the test would have been terminated.

If the igniter

Ignition was achieved on all of these tests.

13

THRUST

PCIGN 7

FUEL VALVE OX VALVE SIGNAL SIGNAL

Figure 6. Thruster I g n i t i o n Sequence - Steady State Testing

14

111, Method o f Approach (cont.)

An osci l lograph trace o f a t y p i c a l thruster pulse sequence I s shcwn i n Figure 7. The pulse t e s t var iables were valve actuat ion (GN2 o r GHe), pulse width (valve open t o valve close), i n l e t pressures and I n l e t temperature (-lent or cold). The first four (4) pulse tes ts were run w i t h a GN2 dr iven

actuator. The remaining tes ts used a GHe driven actuator t o speed the valve

opening and closing times.

F igu re 7. Thruster Pulse Sequence - GHe Actuator

16

I V . SIGNIFICANT RESULTS

The ignition and steady-state performance characteristics of spark ignited GOX/Ethanol propellants were defined and pulse mode capability was demonstrated. The prototype igniter was tested over the range of conditions listed in Table 11. A total of one hundred thirty-eight (138) igniter tests were run to define the ignition limits in terms of cold flow pressure, igni- tion chamber diameter, mixture ratio, spark energy and propellant temperature.

The effect of cold flow pressure (P) and chamber djameter ( D ) on the flame quench ignition limits were correlated as shown in Figure 8. (-125°F) fuel was found to only slightly reduce the ignition limits. oxidizer had no measurable effect on ignition. 50 MJ/spk to 10 MJ/sph ha> a more significant effect on the ignition limit as indicated in Figure 9.

Cold Cold

Reducing the spark energy from

Sixty-seven additional igniter tests were run to verify igniter cool- Ignition was achieved at all ing, C* performance and pulse mode capability.

conditions where ignition was predicted.

A total of eighty-four thruster tests were made. Nineteen added scope tests were made using residual hardware (Contract NAS 9-15958) and sixty-f ive tests were made using the prototype thruster (Figure 1) designed and fabri- cated on this program. The results of the added scope tests were used to confirm the performance of a swirler/like doublet OF0 injector f type thruster.

The previous thruster test data are summarized in Table I tion and Isp efficiencies are shown in Figure 10 for ambient and

r the proto-

I. Combus- cold propel-

lant. of mixture ratio tested. 6% at the design point mixture ratio. Twenty five percent (25%) fuel film collant was used on all of the prototype thruster testing.

The ambient temperature performance is nearly flat over the whole range The cold propellant reduces Isp efficiency by about

17

W

3 U rn 0, a E F - 0

s u 0 .- -0 C u o m v OIL E O v u

. X I O L V a l

I .-

k % Lv) o m - ~ m

a c m

m

E

f c

n Q

I

0

0 V --. 0

0 V

c

c

(u

*

0

m c u * u c u o w OI - u al U W

a

V

c

B

C 0

U * hl

.C

.,- E

a - U

0

rn E

0 0 V

.- c

& c c I&

0 m H

m

n m I-

C)

I- m

* 0

P P \

rl

0

a

Y U

W c

E V

W c a .C

U v -

2 0 c 9

al U

3 0

L U u L u m -u c n PIP) - L

c

0 m d

.4

d

s I-

Q m I-

e C

9 P

rl

rl

=l

W U

u c

t V

r c,

x ._ U

Q U In 0 L

0 c u m u c w m VI- cc a w o n 0 L L W C L U - v

&Z - u

0 m d

d

c'

0 c aa

n m I-

* 0

U

0 V -. 0

0 V

c

c

m

W

>

0 U

u c

s" V

0 0 c

k er L m U ul W L

0 C m

U ul * 0 U

W m f

L 01

0 L c Y

c3 E rn 0 0

L

- c

n

a

- 8 C 0

U

c WI

z v, 3 0 L s U

U W C

CI

- c

c

c m

.c

t n C 0 U

0 c 0 V

W W

U

U

m 4, c,

U 0)

0 U

c

c

a c

c

s

H

.. W t- 0 2

3

U W

L al v)

L 0 1c

W a ul n aJ L

W 4

* U

m I

h -0 (0

W U ul

0 U

w C 0

-

a

I

c 0, a 0 L

4:

U W L 0, U

L 0 c

0 L

n U 0) L

Y

c

a

c c 0

0 V

0 U

U L

4, (c

W L

u 3 ul ul 4, L .

c

if * L

A = = 0 0

d

0, L

U m L

a

2 E P U

c N W u L c

al 3 *-

al c e 0) * * LL 0

0 v)

a.t

al Y

c

::

c

t c I: 0 L 0 CL Q

W L Q

U W

Y m L al

a

e Y

cr c m c - f a V c

8 n

..

f 0 10 a

Q Q d

' 0 ° O8 * 'I a 0

D 7 B

Q

20

3 L aJ t W

Y L a 0. v,

re 0 c, V aJ re

re W

21

0 4

rn c

0 -

U \ L

U \ L

ru I m d

0

2 N 0,

m N I

E 0

- - U C al rn c w t-

c

N r a

Y c v)

0 m al I

-

U E al ?

P U c 0) c

Q U E u - s

U c 0 .C

P U E al - 9

V

0 V 7

U

0 V c

3 r r x

22

8

!

U

I - zp o,

atc

Ql

CD

0.

a

L Q, c1 lA 3 t r c

23

I V , S ign i f icant Results (cont.)

The th in-wal l chamber thermocouple data w i th 25% FFT indicate that 40%

FFC w i l l be required t o cool a columbium rad ia t ion cooled f y igh t engine. Additional tes t ing w i l l be required t o define performance For 40% FFC.

Thruster pulse performance was determined over a pulse width range of 40 msec t o 500 msec. The B i t Isp i s shown i n Figure 11.

A l l o f the thruster tests were smooth w i th no evidence o f combustion The exhaust plumes were observed t o be clear and clean. i n s t a b i l i t y .

evidence o f carbon format'on or deposit ion was found.

No

24

c c 0 0

0 0 0

0

7 @\

\

00

-

0 cu cu 0 0 cu

0 a c

0 w c

0 e c

0 N c

V. L I M I T A T I O N S

The applicab l i t y of the GOX/Ethanol th rus ter combustion data obtainnd on t h i s program are l i m i t e d t o the 0-300 ps ia pressure range and the 0.778 t o 3.29 mixture r a t i o range.

t o 40 mixturs r a t i o range. t o -156°F. The fue l f i l m coolant data are l i m i t e d t o 26% FFC. No other data

l i m i t a t i o n s are known.

The spark i g n i t i o n data are appl icable t o the 0.4

Propellant temperatures ranged from ambient down

26

V I . RECOMMENDED FOLLOW-ON WORK

The following additional work i s recommended.

1. Further GOX/Ethanol thruster testing should be conducted to evaluate axial fue? film coolant injection and performance using up t o 45% FFC

and columbium chambers to finalize FFC requirements f o r R C S thrustetb.

2. Extreme fuel rich (MR <0.1) GOX/Ethanol igniter testing should be conducted to completely define the fuel-rich ignition limit.

3. Ignition testing with other hydrocarbon fuel (i.e., Propane and Methane) should be done to expand the hydrocarbon fuel data base.

27

V I I . CONCLUSIONS

It is concluded that spark ignition o f the GOX/Ethanol propellant combination i s reliable and that thruster pulse mode capability was demon- strated. The spark ignition limits were defined and a GOX/tthanol thruster data base generated.

28

REFERENCES

1. Michel, R. W., "Combustion Performance and Heat Transfer Characteriza-

t i o n o f LOX/Hydrocarbon Type Propellants", Final Report, NAS

9-15958, September 1983. Aerojet TechSystems Company.

2. Orton. G. F., Mark, T. D., and Weber, D. D., "LOX/Hydrocarbon Auxi l iary

Propulsion System Study", F'nal Report NAS 9-16305. Douglas Corp. Report M!JC E 2548, July 1982.

McOonnel

3. Lawver, 6. R., "1gnit':n Lharacterization o f LOX/Hydrocarbon Propellants" T. :'.; I data Dump, Contract NAS 9-16639, June 1983,

Aerojet TechS; tems Company.

4. Lawver, B. R., " I g n i t i o n Characterization of LOX/Hydrocarbon

Propellants'', Task I1 Data Dump, Contract NAS 9-16639, FeFuary 1983, Aero j e t TechSystems Company.

5. Lawver, 6. R., 'I Ign i t ion Characteri tat iov of LOX/Hydrocarbon

Propellants", WBS 8.0 Data Dump, Contract NAS 9-16639, May 1983, Aero j e t Techsystems Company.

6 . Lawver, 6. R., " I g n i t i o n Characterization o f LOX/Hydrocarbon Propellants", Task I11 Data Gump, Contract NAS 9-16639, August 1984, Aerojet Techsystems Company.

7. LaBotz, R. J., Rousar, 0. C., and Valler, H., "High Density Fuel

Combustion and Cooling Investigation'l, Final Report NASA CR 16577, Aerojet TeThSystems Company, September 1980.

29