NASA Technical Memorandum 103166

AIAA-90"2241 ......

A Comparison of Analytical Results for

20 K LOX/Hydrogen Instabilitie_ ...................

J--: :-:: ..... Mark D. Klcm and Kevin J, Breisacher

.................... Lewis Research Center _ _ - __L: _:-_-:__ _ _=7

Cleveland, Ohio

(NASA-TM-lO_I66) A COHPA_ISON nF ANALYTICAL

RrS!JLTS F_R 20 K LOX/HY_R_GFN INSTAR[LITIES

(NASA) 17 p CSCL ?IH

G.3/20

N?O-251R6

Unclas

0290T34

Prepared for the

26th Joint Propulsion Conf_e_r¢_nce .......

' cosponsored by the AIAA.SAE, ASME, and ASEE

....... _ ....... Orlando, Florida, July 16-18, 1990 _ _

7

Z_

A CO_ARISON OF ANALYTICAL RESULTS FOR

20K LOX/HYDROGEN INSTABILITIES

Mark D. Klem I and Kevin J. Brelsacher I

NASA Lewis Research Center

Cleveland, Ohio

OO

u_

u_

!

Abstract

Test data from NASA Lewis' "Effect of

Thrust Per Element on Combustion

Stability Characteristics of HydrogenOxygen Rocket Engines" test program are

used to validate two recently released

stability analysis tools. The first tool

is a design methodology called ROCCID

(ROCket _ombustor Interactive Resign).

ROCCID is an interactive design and

analysis methodology that uses existing

performance and combustion stability

analysis codes. The second tool is

HICCIP (High frequency Injection _oupled

_ombustion Instability _rogram). HICCIP

is a recently developed combustion

stability analysis model. Using a matrix

of models, results from analytic

comparisons with 20K LOX/H 2 experimental

data are presented.

Introduction

In order to provide a convenient tool for

the analysis and design of liquid rocket

engine combustors, NASA Lewis initiated

"LOX/Hydrocarbon Rocket Engine Analytical

Design Methodology and Validation," NAS

3-25556. The purpose of this program was

to compile and evaluate existing codes,

select the most appropriate codes for an

interactive program, and provide a

modular framework that would make these

codes readily usable. The product of

this effort is ROCCID (ROCket _ombustor

Interactive Resign). ROCCID is currently

capable of analyzing mixed element

injector patterns containing impinging

llke-doublet, impinging triplet,

showerhead, shear coaxial, and hydraulic

hollow-cone swirl coaxial elements. Real

propellant properties for liquid and

gaseous oxygen, hydrogen, methane,

propane, and liquid RP-I are currently

available. ROCCID also contains the

logic to interactlvely create a combustor

design which will meet input performance

and stability goals.

Currently, ROCCID has been released to

industry for their use and their

assistance in identifying any problems

that may still exist in the code (i.e.,

BETA test). As part of this "BETA test"

process, the authors selected a set of

instability data and ran ROCCID through

its matrix of available models both as an

operational test of ROCCID and as a quick

comparison of the range of results

obtained by running existing models for

the same problem. It should be made

clear that it is not the intent of the

authors to indicate one model as being

superior to another. Rather, our intent

is to illustrate the qualitative nature

of the state of the art of combustion

instability analyses and to demonstrate

the convenience of using a framework such

as ROCCID for making model comparisons.

HICCIP (High frequency Injection Coupled

_ombustion Instability Program) (Ref. I),

a recently developed combustion stability

model, was also used to analyze the same

set of data. HICCIP is in its validation

phase and further refinements to the code

are planned in the future. The analysis

of these tests and comparisons to

existing models provide a means of

validating HICCIP at this stage in its

development.

Description of _he Models

The matrix of models utilized in ROCCID

to analyze the stability of the 20K

IMember AIAA

Copyright © 1990 by the American Institute of Aeronautics

and Astronautics, Inc. No copyright is asserted in the

United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under

the copyright claimed herein for Governmental purposes.

All other rights are reserved by the copyright owner.
combustors is shown in Table I.

HIFI (High Frequency Intrinsic Stability

Analysis) (Ref. 2) utilizes a

concentrated (at a single axial plane)

combustion model and the sensitive timelag approach. In the concentrated

combustion approach, the combustion

chamber is divided into two regions: the

region upstream of the combustion plane

where the mean velocity is assumed to be

zero and a region downstream of the

combustion plane where the velocity is

assumed to be non-zero and constant.

Using the separation of variables

technique, a general solution governing

the velocity potential function (obtained

by combining the continuity and momentum

equations and an isentroplc

pressure/denslty relationship) can be

obtained. The injector face boundary

condition is applied to calculate the

chamber admittance (the perturbation in

velocity divided by the perturbation in

pressure) upstream of the combustion

plane. Using the nozzle admittance as a

boundary condition, the chamber

admittance downstream of the combustion

plane is calculated. Continuity is then

applied at the combustion plane to relate

the burning admittance-to the upstream

and downstream chamber admittances. The

pressure interaction index, n, and the

sensitive time lag, t, for neutral

stability can then be calculated using an

expression derived by Crocco (Ref. 3).

The 3D Distributed Combustion Model

(DIST3D) (Ref. 2) contains a more

detailed model of the chamber acoustics

that can include the effects of stability

aids (absorbers and baffles) and a more

realistic treatment of the combustion

zone. The combustion zone is modeled as

being distributed over a significant

portion of the combustor. The combustion

zone parameters (chamber acoustic

velocity, chamber gas velocity, etc.) are

modeled in a plecewlse-llnear fashion.

The features for stability aids are not

used in the current analyses.

The Combustion Response Prediction (CRP)

model (Ref. 2) calculates the open-loop

response of the burning rate to a

specified acoustic oscillation in the

combustion chamber. Using analytical

solutions for the pressure and velocity

fluctuations in a closed cylinder, the

pressure and velocity histories for a

specific instability mode can be

calculated. With the pressure and

velocity fields specified, the

vaporization history for a droplet can be

calculated using Ranz-Marshall (Ref. 4)

correlations for the heat and mass

transfer coefficients and assuming the

transfer processes are quasl-steady

state. Using an energy balance, the

temporal variations of the drop diameter,

temperature (assumed uniform), and the

vaporization rate can be calculated. The

continuous injection of propellant is

simulated as arrays of single drops that

are injected from various radial,

circumferential and temporal locations.

An In-phase and out-of-phase combustion

response can then be calculated.

The Lumped Parameter Injection Response

Model (INJ) (Ref, 2) is similar to that

described in Reference 5. The injection

element admittance is characterized by

its resistance, capacitance and inertance

terms. It is assumed in INJ that the

wavelength of an element oscillation is

much larger than the elements physical

dimensions.

In the Lewis Injection Response Model

(LEINJ) (Ref. 2), flow and pressure

oscillations in the tubes of a coaxial

injector element are determined by

solving the one dimensional mass and

momentum conservation equations as a

function of axial position. Pressure

losses in the orifice are modeled using a

discharge coefficient. Oscillations in

the manifolds are described by the three

dimensional wave equation for finite Math

number flow solved by using the

separation of variables technique.

SMITH-REARDON refers to the widely used

N-tau correlation developed in the 1960's

(Ref. 6). AEROJET 20% Vaporized is an

analytical procedure developed to

calculate the sensitive tlme-lag (Ref.

2).
AEROJET and DROPMIX are two of the drop

size correlations currently available in

ROCCID for coaxial injectors.

_ICCIP Description

HICCIP treats flow oscillations, produced

by chamber pressure oscillations, in both

the fuel and oxidizer elements using the

same approach as outlined for LEINJ.

The atomization process was followed on a

quasi-steady basis using the relations of

Mayer for coaxial injection as described

in Reference I. Instantaneous local

velocities were determined at various

times in an oscillation cycle. The

atomization rate and drop size produced

in the combustion chamber were calculated

for conditions corresponding to these

velocities. An effective atomization

plane was established at the location

where 50% of the mass of the jet has been

atomized. The vaporization process was

assumed to begin at this effective

atomization plane.

The mean drop size of the spray was

specified as the mass average of the

drops produced at various locations in

the combustor. Drop size groups are

calculated using the mean drop size and a

log normal distribution. Normally, Ii

drop size groups are used to calculate

steady state performance and 5 drop size

groups are used for instability

calculations. The quasl-steady

vaporization theory of Reference 7 was

modified to include finite thermal

conductivity by calculating the heat

transfer between shells within each drop.

The influence of internal droplet

circulation was approximated by

multiplying the physical thermal

conductivity by a factor of 2.7.

Injection of drops was considered to

occur at 30 discrete intervals during a

cycle of oscillation. The average and

oscillating burning rate during a cycle

are determined at 30 axial locations.

The drops were injected at a radial

location corresponding to the median area

and at a single angular location.

Gas phase oscillations in the chamber

were calculated assuming distributed

combustion and by using a wave equation

solved by separation of variables.

Between axial locations in the combustor,

the oscillation profiles were calculated

using this wave equation. At the various

axial stations, the average and

oscillating velocity components were

adjusted to account for the average and

oscillating burning rates. An iterative

solution technique is necessary to obtain

the same burning rates from the

vaporization calculations as those used

in the wave dynamic calculations. The

wave dynamic calculations begin at the

nozzle and proceed to the injector with

an assumed average burning rate response.

If the velocity at the injector face was

not zero, a new average burning rate

response was assumed. The calculations

were then repeated with different

frequencies (real and imaginary) until

the average burning rate response (the in

and out-of-phase components) agreed with

the vaporization rate calculations. The

local speed of sound used in the wave

dynamics calculations varied as a

function of axial position based on the

amount of burning that had occurred by

that position. Oscillations at the

nozzle end of the chamber, were assumed

to follow those for an ideal, shortdistributed nozzle.

Experimental Data

The experimental data chosen to be used

for model comparisons were the data

generated in Reference 8. This data set

was chosen for several reasons. First,

there is significant documentation of the

hardware specifications and operating

conditions for meaningful comparisons to

be made. Second, these data are not

included in the sets of test cases used

in the development of ROCCID. Finally,

the absence of stability aids (absorbers

or baffles) in the combustor permits

comparisons to be made with HICCIP.

The hardware used in Reference 8 are 20KIbf-thrust Liquid Oxygen (LOX)/H 2 engines

operated at a nominal chamber pressure of

3
300 psi. Six coaxial injector

configurations were tested. The thrust

per element varied from 20 to 1000 Ibf.

The chamber consisted of a 10.78 inch

diameter heat sink cylindrical section

bolted to a convergent-dlvergent heatsink nozzle. Stability characteristics

were mainly determined using hydrogen

temperature ramping (several of the

injectors that were still stable at the

lowest obtainable H2 injection

temperature were bombed (pulsed using an

explosive charge). The stability of the

combustors was characterized by a H 2

injection transition temperature. The

transition temperature was defined as the

H 2 injection temperature where unstable

combustion was first encountered.

The results of this test program are

summarized in Table II. Three of the six

injectors (992, 397, and 201 elements)

were driven unstable during temperature

ramping. The I00, 35, and 20 element

injectors did not have a discernible

transition temperature in the range of

hydrogen temperatures tested. No further

analysis was attempted for these

injectors because they do not provide a

very useful test for the codes. The 992,

397, and 201 element injectors all

encountered instabilities during

temperature ramping. Reference 8

provides strip chart traces of a typical

instability for each of these three

injectors (Fig. I). These traces

indicate the acoustic mode, frequency,

and peak-to-peak limit amplitude of the

instabilities encountered. The 992

element injector is of particular

interest in that the temperature ramping

did not induce the "usual" first

tangential (IT) mode but excited the

second tangential (2T) mode (Fig. i).

This is of interest because in a recent

test program with methane, temperature

ramping failed to induce the IT mode but

consistently excited a higher order mode.

The 992 element injector was chosen as

the focal point for these analyses.

ROCCID Results

The absence of stability aids and the use

of input parameters generated by the

existing logic in ROCCID (i.e. no

"tweaking") provide a very stringent test

of the models in ROCCID. The ability of

ROCCID to provide stability margin by the

deslgn of stability aids (its strongest

feature) is not demonstrated. There was

one major drawback with choosing to

analyze the 20K LOX/H 2 data with ROCCID.

The INJ model in ROCCID is hardwlred to

analyze coaxial elements in which the

orifice is located at the top of the LOX

tube upstream of a substantial diffuser

section (Fig. 2). This configuration

represents the so called "modern"

injection element. Unfortunately, the

elements in the 20K LOX/H 2 engine had

orifices at the exit of the elements with

no diffuser section. To remedy this

situation, a very short (i0 -s inch),

fictitious diffuser section was added

downstream of the orifice. The width of

this diffuser section was then adjusted

until reaso_nable agreement was obtained

with experimental pressure drop data.

Adjusting this fictitious diffuser

section width did not significantly

affect the predicted drop size. The

location of the orifice in an element is

not a problem when using the LEINJ

module.

Tables III and IV show the performance

and stability characteristics for the

matrix of ROCCID models run in the study.

The arrangement of the table is based

upon the increased detail or mechanistic

approach of the models utilized. It

should be noted that with the interactive

framework of ROCCID, once the geometry

and operating conditions for an engine

are input, different analysis modules can

be selected by interactlvely setting the

appropriate switches. With the exception

of CRP, the computer runtlmes associated

with these modules are sufficiently short

that a substantial matrix of models can

be run in an afternoon.

The experimental performance data for the

20K LOX/H a engine is somewhat erratic.

An examination of Table II shows that C*

efflciencles in excess of 100% were

reported. Unfortunately, the C* reported

for test 0002 is probably erroneous due
to incorrect hydrogen weight flows

resulting from mixing tube storage.

Additional performance data reported in

Reference 8 indicates that the 992

element injector is characterized by a C*

efficiency of approximately 99.7% during

stable operation. During temperature

ramping, the C* efficiency of the 992

element injector is approximately 96%.

The C* efficlencles calculated by the

performance modules of ROCCID are in

reasonable agreement with those observed

experimentally. This is despite the fact

that the drop sizes calculated by two of

the available correlations can disagree

by more than a factor of three. This is

not an uncommon occurrence for the drop

size correlations in ROCCID. Both

correlations generally predict a larger

drop size and lower performance for the

temperature ramped test case 0002.

However, when DROPMIX is used in

conjunction with LEINJ this trend is

reversed. This is similar to the trend

observed with HICCIP.

There was a wide range of calculated

stability results obtained as indicated

in Table IV. For all the models where an

instability is indicated (growth rate (A)

greater than 0), the ratio of burning

admittance (YB) to injector admittance

(YJ) is less than one (indicating an

injection coupled instability). The

cases run with LEINJ indicate an even

more pronounced injector influence (even

smaller values of YB/YJ).

For test case 0001, fourteen of the

eighteen different model combinations

correctly predict stable tests. For test

case 0002, only six of the eighteen

combinations predict a 2T instability.

Of those six, the predicted frequency

varies by ±I000 Hz. from the frequency

that was reported in Reference 8.

The influence of calculated drop size on

stability results is demonstrated by

examining the first two model entries in

Tables III and IV. Using the same

chamber acoustics, injection, and

combustion models but changing the drop

size correlation model, the mode of the

calculated instability switched. It is

also interesting that erroneous IR and 3T

modes occurred only when using the

smaller calculated drop size (AEROJET).

The use of LEINJ might be expected to

make the 2T instability more prominent

(due to the resonance at 4100 Hz

indicated in Fig. 3), but this was not

observed. In fact, LEINJ seemed to

increase the occurrence of a calculated

IT instability (Table IV HNLAS, HNLDS,

and HCLA). In general, it might be

expected that as one proceeds down Table

IV (increasing model complexity) better

agreement would be obtained between

experimental and analytical results, but

this is not the case.

_ICCIP Results

A review of the reports from the NASA

LeRC LOX/H 2 stability programs of the

1960's shows that modes other than the IT

have occurred but not as frequently as

IT. Indeed, if as indicated below that

LOX tube coupling plays a significant

role in the occurrence of these other

modes, then the repeated use of hardware

with similar injection element dimensions

may account for the occurrence of these

modes.

The performance module in HICCIP

calculated C* efflclencles of 97% and 99%

using drop sizes (D30) of 172 microns and

90 microns for test cases 0001 and 0002

respectively. Currently, there are no

correlations for the constants in the

relations for drop size or atomization

rate in HICCIP. These constants are the

two remaining "knobs" that are tweaked

that limit the pre-test applicability of

the code.

Figure la shows the amplitude spectral

density traces for test 0002 with the 992

element injector. At a hydrogen

transition temperature of 78°R, a second

tangential instability was excited and

grew to a 150 psi peak-to-peak amplitude

at a frequency of approximately 5420 Hz.

There was also some activity at

approximately 3200 Hz with an amplitude

of 28 psi peak-to-peak. Figure 3 shows

the normalized response (normalized by
the ratio of chamber pressure to

flowrate) of the LOX injection element

for the first and second tangential modes

for test case 0002. The LOX tube has a

resonance at approximately 4100 Hz. The

LOX side response is near a minimum at

the IT frequency for the chamber (3435 Hz

assuming thoroughly mixed gas

composition). Although the resonance for

the LOX system is well below the 2T

frequency for the chamber (5697 Hz),

during ramping, the injection of the cold

hydrogen can substantially reduce the

local acoustic velocity at the injector

face (Ref. 9). This can allow the LOX

post to tune with the chamber and

initiate an instability that shifts in

frequency as it grows in amplitude. This

is apparently what happened in the case

of the 40KLOX/methane IT instabilities.

The sharp spike in Figure 3, occurring at

approximately 2700 Hz, is due to a

resonance in the LOX manifold. These

resonances generally have such a small

bandwidth that they have difficulty

initiating and sustaining an instability.

The fuel side response for the stable

test 0001 and the unstable reduced

hydrogen injection temperature test 0002

are shown in Figure 4. Figure 4 clearly

shows the significant increase in fuel

side response as the hydrogen injection

temperature is reduced (test 0002). The

fuel side response also shows a

significant change as a function of

frequency for the reduced temperature

case. The maximum response occurs near

the 2T chamber frequency (5697 Hz).

While the response plots for the

injection elements may provide some

insight into why, and if a particular

instability mode occurs, they are not an

absolute indicator. The graphs only show

the magnitude of the injection response

and not the phase relationship with other

processes. HICCIP attempts to account for

the phasing and magnitude relationships

of the dominant physical processes

occurring during an instability.

exhibiting a negative damping rate are

unstable. Figures 5 and 6 indicate the

ambient hydrogen temperature case (test

0001) would be stable until an

approximately 35 psi half-amplltude (half

the peak-to-peak value) organized

disturbance would be introduced into the

combustor. From the HICCIP results, it

would be expected that this case would be

spontaneously stable, which conforms with

experimental observations.

Unfortunately, the 992 element injector

was never bombed. A higher amplitude

disturbance might have induced an

instability. Figures 7 and 8 are the

HICCIP results for the IT and 2T modes

respectively for test case 0002 which was

ramped unstable. The IT results indicate

an instability at this hydrogen injection

temperature that grows to a limit cycle

amplitude (ie. where the damping rate

curve crosses from negative to positive)

of approximately 30 psi (60 psi peak-topeak) at approximately 3100 Hz (the point

on the frequency curve corresponding to

the zero of the damping rate curve).

This overestimates the magnitude of the

IT activity observed by a factor of 2

(Fig. I). However, the frequency

calculated for the IT is in good

agreement with the data. Since the

growth rates for the 2T oscillation are

substantially larger than those for the

IT oscillation, it is possible the test

was terminated before the IT oscillation

reached its limit cycle amplitude. The

results for the 2T mode exhibit an

instability that grows to 80 psi (160 psi

peak-to-peak) at approximately 5580 Hz.

Although the frequency predicted for the

2T mode is above the 5420 Hz observed

(the HICCIP results are closer to the

equilibrium value of 5697 Hz), the limit

cycle amplitude results are in good

agreement with experiment (Fig. I).

Figure 8 also indicates that the

frequency at which a small amplitude (5 -

10% of chamber pressure) disturbance

begins to grow is between 4000 - 4200 Hz

(the region of resonance for the LOX

tube).

Figures 5 - 8 show the HICCIP results for

the 992 element injector. The regions
Conclusions

The interactive nature of ROCCID provides

a powerful framework for making

comparisons with existing combustion

instability models.

The qualitative nature of the unanchored

use of the instability modules in ROCCID

is apparent. The models' pre-test

predictive capabilities are questionable.

Additional drop size correlations and/or

guidelines for the range of parameters

the correlations are valid over should be

incorporated in ROCCID.

HICCIP analyses indicate that LOX tube

acoustic resonance probably played a

significant role in the occurrence of the

2T mode of the 992 element injector.

HICCIP can provide some insight into the

mechanisms that may cause an instability

to occur. However, further refinement of

HICCIP is required to make it a useful

pre-test analysis tool.

References

i. Priem, R.J., and Breisacher, K.J.,

"Nonlinear Combustion Instability

Model in Two- to Three-Dimensions,"

NASA TM-102381, 1989.

2. Muss, J.A., and Johnson, C.W., "LOX/

Hydrocarbon Rocket Analytical Design

Methodology Development Validation,

NAS 3-25556, ROCCID User's Manual,

NASA CR-185187, 1990.

3. Crocco, L., and Cheng, S., Theory of

Combustion Instabilit_ in Liquid

Pr__ellant Rocket Motors, Butterworths Scientific Publications,

London, 1956.

4. Ranz, W.E., and Marshall, W.R., Jr.,

"Evaporation from Drops," Chemical

Engineering Progress, Vol. 48,

No. 3, Mar. 1952, pp. 141-146.

5. Harrje, D.T., and Reardon, F.H.,

eds., "Liquid Propellant Rocket

Combustion Instability," NASA

SP-19&, 1972.

6. Smith, A.L., and Reardon, F.H., "The

Sensitive Combustion Time Lag Theory

and Its Application to Liquid Rocket

Combustion Instability Problems,"

(Vol. I Engineer's Manual)

TR-AFRPL-TR-67-314, 1968.

7. Priem, R.J., and Heidmann, M.F.,

"Propellant Vaporization as a Design

Criterion for Rocket Engine Combustion Chambers," NASA TR-R-67, 1960.

8. Salmi, R.J., Wanhainen, J.P., and

Hannum, N.P., "Effect of Thrust Per

Element on Combustion Stability

Characteristics of Hydrogen-Oxygen

Rocket Engines," NASA TN D-4851,

1968.

9. Jensen, R.J., Dodson, H.C., and

Claflin, S.E., "LOX/Hydrocarbon

Instability Investigation," NASA

CR-182249, 1989.

Table I

ROCCID Models

Chambe r Combustion Injector prop SSze TAU

Agoustlcs Response Admittance Co_relation Correlation

HIFI CRP INJ AEROJET AEROJET 20%

DIST3D N-TAU LEINJ DROPMIX SMITH-REARDON
!

m

n

a

G

,q L.

6 •

u

4)

• °

,..,p

-_'_ _ _;8 _ _ _ _ _,_

k_4 k't4 _ k_4

_ee.

%

,ff

1--

I--

0

00000000 0__

00000000 00000

00000000 000_0

r

000000000

000000000

0000 O0

0000 O0

8
Table III

ROCCID Performance Calculations

MODELS TEST D30 DELTA DELTA C*

P. Po EFF

(_) (PSl) (PSi) (_)

HNIAS 0001 _6,1

0002 61.2

HNIDS 0001 85.6

0002 117,0

HNIAA 0001 26.1

0002 61.2

HNIDA 0001 85.6

0002 _17,o

DNIAS 0001 26.1

0002 61.2

DNIDS 0001 85.6

0002 117.2

DNIAA 0001 26.1

0002 51,2

DNIDA 0001 85.6

0002 I17.0

HCIA 0001 26.1

0002 61.2

HCID 0001 85.6

0002 117.0

DCIA 0001 26,1

0002 61,2

DCID O001 85,5

0002 117.0

A4,1 391,3 99,9

19.9 365.8 97.8

45.3 401.8 98,1

_i.0 385,1 94,7

44.1 391.3 99,9

19.9 565.8 97.8

45.3 401.8 98.1

51,0 $85,1 94,7

44.1 391.3 99.9

19.9 365,8 97,8

45,3 40_,8 98,I

20.9 385,7 94,7

44.1 391.3 99,9

19,9 365,9 97,8

45.3 40!L8 98.1

21.0 385,1 94,7

44_I 39_,3 99,9

19.9 365,8 97,8

45,3 401,8 98,_

21.0 385.1 94,7

44.1 391.3 99.9

19,9 365,8 97,8

45,5 402,0 98,0

21.0 385,1 94,7

HNLAS 0001 26.1 44,% 391,3 99,9

0002 64.0 19.8 282.6 99.9

_U_S 0001 85,6 45,3 401,8 9s.1

0002 58.9 19.2 273,9 99,8

HCLA 0001 Z6,1 44,_ 391,3 99,9

0002 64,0 79,8 _82,7 98,%

HCLD 0001 85,6 45.3 401,8 98,1

0002 51.7 18.1 _23,8 99,8

DCLA 0001 26._ 44,_ 391,3 99,9

0002 _4,0 19,8 282,6 98.1

DCLD 0001 85,6 45,3 40_,8 99.4

000_ 59,0 19.2 Z74,_ 99,8

Node1 designations coded as follows.

First Letter: B-HIFI, D-DIST3D;

Second Letter: C-CRP,N-N-TAU;

Third Letter: I-INJ, L-LEINJ;

Fourth Letter: A-AEROJET, D-DROPMIX;

Fifth Letter: A-A£ROJET 20 Z, S-_4ITH-REARDON.
_4

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W WW,_

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T

150 PSl

(1050 N/M2)

PEAK TO PEAK

1

SECOND-TANGENTIAL MODE

I

0 5420 10 000

FREQUENCY, HZ

(a) 992 ELEMENTS, THRUST PER ELEMENT, 20 LB (89 N); OXIDANTFUEL RATIO, 5.27.

I

90 PSI

(620 M/M2)

PEAK TO PEAK

I I

FIRST-TANGENTIAL MODE

I

0 3400 10 000

FREQUENCY, Hz

(b) 397 ELEMENTS, THRUST PER ELEMENT, 50 LB (222 N); OXIDANTFUEL RATIO, 4.84.

iFIRST-TANGENTIAL

37 PSI 2 _MODE I FIRST-RADIAL MODE

_ (255 N/M ) l]h I_ (QUESTIONABLE)

PEAK TO PEAK

I I 1

0 3490 I0 000

FREQUENCY, Hz

(c) 201 ELEMENTS, THRUST PER ELEMENT, 100 LB (445 N); OXIDANTFUEL RATIO, _.Gg.

FIGURE I. - TYPICAL AMPLITUDE SPECTRAL DENSITY GRAPHS (REPRODUCED FROM NASA TN D-4851).

II
__¢._.,1//*-- _///////_ORIFICE

Y//////_,_ _////////_

GENERALELEMENT

(a) NASA TN D-4851 ELEMENTDESIGN.

DIFFUSER _ - __ _-- ORIFICE SECTI ON_ _1-

I i _- I I

(b) ROCCID ELEMENTINPUT SCHEMATIC.

FIGURE 2. - SHEAR COAXIAL ELEMENT DESCRIPTION.

i,,,-,-, TEST 002, IT

_" "-'-" TEST 002, 2T

.40 -- I

.30 -- I

_ .20

\/

o \v_]_ /

2000 3000 4000 5000 GO00

FREQUENCY, HZ

FIGURE 3. - 992 ELEMENT INJECTOR OXIDIZER SIDE RESPONSE

CHARACTERISTICS.

12
8

6 ._ / _ TEST 001, 1T

J'* S _ -._-_ TEST 001, 2T

j_ / .... TEST 002, IT

q Z" I¢ ----- TEST 002, 2T

2

o_\,//I I I I

2000 5000 4000 5000 6000

FREQUENCY, Hz

FIGURE 4. - 992 ELEMENT INJECTOR FUEL SIDE RESPONSE

CHARACTERI STICS.

100 -- 3800 --

50

T

t2

o

(..9

-50

-100

-150

3400 --

N

N

z 3000 --

la.J

U.

2600 --

i /

2200 I I \-4" I I I

0 20 40 60 80 100 120

AMPLITUDE, PsIA

FIGURE 5. - TEST 0001 IT HICCIP STABILITY CALCULATIONS.

I

140

13
2OO

100

T

_ o

-100

-200

6000

5500

5000

_5oo

LI.

-- _ _!! _ " FD_UIENNGCY_

\ I k/ /

- Vx/', /

4ooo- J V \k.. / /

3soo ! I I I I I

0 20 40 60 80 100 120

AMPLITUDE, PSIA

FIGURE 6. - TEST 0001 2T HICCIP STABILITY CALCULATIONS.

I

140

'3"

U

LIJ

500 --

400 --

300 --

200--

100 --

0--

-100 --

-200 --

3350

_o_ /.. \,\

315o / \,,,

3°5°I-! _," V

3000 t--_ J

If - FREQUENCY

2950 V -' DAMPING

29001 I J I I I

20 40 60 80 100 t20

AMPLITUDE, PSIA

FIGURE 7. - TEST 0002 IT HICCIP STABILITY CALCULATIONS.

I

140

14
8O0

6OO

T 4O0

200

o

-200

-400

-600

6000 IF--' FREQUENCY

5500

5000

4soo

- 3500 I I I I

20 40 60 80 100 120

_PLITUDE, PSIA

FIGURE 8. - TEST 0002 2T HICCIP STABILITY CALCULATIONS.

I

140

15
National Aeronautics and Report Documentation Page

Space Administration

1. Report No. NASA TM-103166 2. Government Accession No. 3. Recipient's Catalog No.

AIAA-90-2241

4. Title and Subtitle 5, Report Date

A Comparison of Analytical Results for 20 K LOX/Hydrogen Instabilities

7. Author(s)

Mark D. Klem and Kevin J. Breisacher

9, Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

6. Performing Organization Code

8. Performing Organization Report No.

E-5538

10. Work Unit No.

582-01-21

11. Contract or Grant No.

13. Type of Reportand PeriodCovered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE, Orlando,

Florida, July 16-18, 1990.

16. Abstract

Test data from NASA Lewis' "Effect of Thrust Per Element on Combustion Stability Characteristics of

Hydrogen-Oxygen Rocket Engines" test program are used to validate two recently released stability analysis

tools. The first tool is a design methodology called ROCCID (ROCket Combustor Interactive Design). ROCCID

is an interactive design and analysis methodology that uses existing performance and combustion stability analysis

codes. The second tool is HICCIP (High frequency Injection Coupled Combustion Instability Program). HICCIP

is a recently developed combustion stability analysis model. Using a matrix of models, results from analytic

comparisons with 20 K LOX/H2 experimental data are presented.

17. Key Words (Suggested by Author(s))

Combustion instability; Rocket; Liquid propellant engine;

Combustor; Interactive design; Stability models;

Computer program

18. Distribution Statement

Unclassified- Unlimited

Subject Category 20

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of pages

Unclassified Unclassified 16

NASAFORM1626OCT86 *For sale by the National Technical Information Service, Springfield, Virginia 22161

22. Price"

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