volume ! main propulsion - nasalately', he has taken on a broader and even n_ore important role...
TRANSCRIPT
II
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VOLUME !
MAiN PROPULSION
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PROCEEDINGS
SPACE TRANSPORTATION SYSTEM
PIiOPULSION TECHNOLOGY CONFERENCE
APRIL 6-7, 1971
M_%RSHALL SPACE FLIGHT CENTER
VOLUI_LE I
SESSION I .k,LA.!NPROPULSION J. THOMSON, CI-IAIRMAN
April ZS, 1971
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6
PROPULSION TECHNOLOGY CONFERENCE
PAPERS
VOLUM/E I MAIN PROP_)LSION
I. "Final Results of the XLR-IZ9 Program"
2. "Two-Phase Flow LH g Pump Inducers"
3. "Saturated LH 2 Turbopump Operation"
4. "Characteri_tics of Feed System Instabilities':
5. "Combustion Oscillations Damping Devices"
6. "Minimum Pressure Loss in High Velocity Flow
Duct Systems"
7. "Engine Onboard Checkout System Study"
8. "Low Frequency Analysis of Rocket Engines
Using Compressible Propellants"
9. "Titaniun: Pump impeller Fabrication and Testing"
I0. "Advanced Thrust Chamber (Non-tubular)"
VOLUME II AUXILIARY PROPULSION
1. "Hydrogen/Oxygen ACPS Englne" s"
2. "Hydrogen/Oxygen ACPS Engines"
3. "High Pressure Reverse Flow ACPS Engine"
4. "Space Shuttle ACPS Shutoff Valve"
5. "Space Shuttle ACPS Shutoff Valve"
6. "Ignition Devices for ACPS"
7. "Spark and Auto Ignition Devices for ACPS"
8. "Catalytic Ignition/Thruster Investigation"
9. "Auxiliary Propulsion Subsystem Investigations"
i0. "injector Performance, Heat Flux and Film Cooling
in Oz/H g Engines"
1 1. "Auxiliary Propulsion Subsystem Definition Study"
12. "Auxiliary Propulsion Subsystem Definition Study"
13. "Acoustic Cavity Use for Control of Combustion
Ins tability"
14. "Noncircular Injector Orifices and Advanced
Fabrication Techniques"
P&W
Rockctdyne
MSFC
Martin M_-_r letta
P&W
S. R. I.
Martin Marietta
Aerojet
Rocketdyne
Rocke[dyne
Rocketdyne
Aerojet
Bell
Marquardt
Rocketdyne
Aerojet
Rocketdyne
TRW
MSFC
MSC
TRW
MDA C
Rocketdyne
Rocketdyne
PROPULSION TECHNOLOGY CONFERENCE
PAPERS
VOLUME III AUXILIARY POWER UNIT AND AlP.BREATHING PROPULSION
I. "Auxiliary Power Unit Design Studies"
2. "Auxiliary Power Unit Design Studies"
3. "Hz Fuel Sy,,_tem Investigation"
4. "Booster and Orbiter Engine Studies"
VOLUME iV CRYOGENS
I. "Orbital Cryogenic Acquisition and Transfer"
2. "Zero Gravity Incipient Boiling Heat Transfer"
3. "Zero Gravity Propellant Transfer"
4. "Recent Developments in High Performance Insulation
Purge Systerns for Shuttle Application"
5. "Effect of Env?: ...................................
6. "PPO Foam Internal Insu'atian"
7. "Internal Insulation Systems for LH 2 Tanks" and
'_Gas Layer and Reinlorced Foam"
8. "Shuttle Cryogen Technology Program"
Airesearch
Rocketdyne
G.E.
P&W
General Dynamics
Univ. of Michigan
LeRC
MDAC
.... __ _[
General Dynamics
MDA C
M<r tin Marietta
MSFC
TABLE OF CONTENTS
VOLUME I
1. Welcome to the Conference - Jerry Thomson
2. Welcome to MSFC - Dr. Weidner
3. Opening Remarks - De1 Tischler
4. Working Group Keport - Jerry Thomson
IvLAIN PROPULSION
I. "Final Results of the XLR-IZ9 Program"
_. R. Atherton
2. "Two-Phase Flow LH2 Pump Inducers"
J. A. King
3. "Saturated LH Z Turbopump Operation':
H. P. Stinson
4. ':Characteristics of Feed System Instabil_tles
R. D. Vaage
5. "Combustion Oscillations Damping Devices"
G. Garrison
P&W
Rocketdyne
MSFC
Martin Marietta
P&W
Page No.
II
13
!5
17-27
31-73
75-119
121-145
147-171
173-197
,
7.
* 9.
"Minimum Pressure Loss in High Velocity
Flow Duct Systems"
C. R. Gerlach
"Engine Onboard Checkout System Study"
R. W. VandeKoppel
"Low Frequency Analy¢is of Rocket
Engines Using Compressible Propellants"
J. M. McBride
"Titanium Pump In_peller Fabrication
and Testing"
J. E. Wolfe
Southwest Research
Institute
Mar tin Mar ietta
Aerojet
Rocketdyne
199-ZZ9
ZZ9-ZJ5
g57-gS8
289-234
9
1
'-'10.
':" 11.
"Advanced Thrust Chamber (Non-tubular}"
D. L. FultmRocketdyne
"High Pressure Hydrogen Effects on Materials" MSFC
W. McPherson
Included in publication but not presented during conference.
325-361
363-378
10
Jerry Tho_lson, ,_r'onfer=nceCh{_irman
W el,- om e
Good Morning Gentlemez:.
I would like to officially open the Space Shutde P='opulsion Technology
Conference and get on with our agenda and the business.
My name i_ Jerry Thomson of Marshall Space Flight Center; and i am
the Chairman of the Propulsion Technology Working Group. I will
serve as the Co_fference Chairman for the next two days.
There are some administra,lve matters and gene:,'al information that I
would like to present to you _bout the Conference, but first Iwould
l',ke to introduce the Dil-ecto?: of Science and Engineering here at the
Marshall Space Flight Center; a man very much at hnme in *.he field of
propulsion. I'm sure me, st of you already know him. tie is Dr. }terman
K. Weidner -- Dr. Weidner, p._ease.
11
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W_tI:in the :)ast dozen years or so, this country has had a total of about
600 succes;ful space flights to her credit. And propulsion devices did their
thing in each and every one of them.
I mention tt_ese fact_ merely topoint out that our professio_'_al specialty
field - Space Propulsion - has come of age al_d has grown to be a
dependable tool in our storehouse of engineering skills. All in all, a
very impressive and proud record.
But this is no time to rest on our laurels. As of late, a new objective
in the field of ,ipace transportation is making its even greater demands
known, it is calling on the hest talents of all of us to do better still.
We are heavily banking on the results of your combined efforts and skiUs
to fuxnish u._ with the firm technological basis so that we may commit
ourselves, with good confidence, to new appli:ations and concepts of
advanced propulsive maci_inery.
It is the purpose then of taxis Space Shuttle Technology Conference to
lvt us take stock of where we stand today in our efforts and to sense how
much still r_m;;ins to be done.
Ann I like to wish this conference mucil success in accomplishing this
_,b.ie ctive.
I do not _a,,t to close here this morning, witt-out saying "Thank you" to
all of you who !,ave come, to Jerry Thomson who acts as the Chairman
/or the whole affair, a:_d to the many session Chairmep. and Speakers
,at.o have worked s.= di;Agently to make this conference possible.
Le,' me conclude, ti_en, by recognizing one mar. who is with us here
tl:i:, morning. This man is Del Tischle r, from NASA's office of
Advanced Re.,;earcl_and Technology. He has served for many years
directing NASA's Technology efforts in the field of Chemical Propulsion.
His leadership has contributed significantly to our past successes and
to the state of our art today.
Lately', he has taken on a broader and even n_ore important role in NASA,
xvhi_h is to plan and to pull together the totality of NASA's Shuttle
Teci_,o!o_y Program. A critically needed and a very difficult task.
I am sure that Del would want to make a Jew .-,.,marks of his own and
_iace the objective of this conference into an even sharper focus.
Thank you,
14
P
Del Tischler, Director, Shuttle Technologies Office, NASA Fieadquarters
Thank you Dr. Weidner.
Gobci Morning, Gentle,hen.
This conference on propulsion and its related area of cry-.genic p"opellant
handling is the third of a series of conferences sponsored by OAI4. T on ther°
technology programs which support the development of the co,_cept of a
fully r.eus._ble Space Launch Vehicle.
it has been suggested that this vehicle be called a Space Plane rather than
a Shuttle. Space Pla;:e may be a dignified desigaation. No matter what
you call it, its' ability to leave the confinenaent of the "I:7,arth's ;..tmosphere
is tota-ly dependent on the performance of rocket equipment. And to do so
ii_ two stages (or less, for that matter) whih: carrying the thermal
protection for reentry, the lifting :iurface,i for aerodynanfic flight, and
tht_ v_hv,.ls for landing as well as a crew compartment for human
intelligence aboard, de_:-_nds the utz_.ost in perfc_-mance of the hydrogen-
oxygen-fueled rocket engines planned for this vehicle.
At the beginning of the Saturn/Apdlo prograrn such engines were
inconceivable. The technological advances since thep have created a
pregnant situation in which birth cf these engines is a matter of nourishment
and tin, e. P:'oposais to d_:velop this equipment are due h,_re in appre:.imately
two weeks. While the main engines are technoiogicaUy in gcnd si_ape ,heir
integration it, to the vehicle, which embodies propellant ve._:sel, cargo carr.;cr,
reentry body, and aircraft, still constitutes a significant prcbler:.. The
propellvnt feed lines in the booster stage, for example, contain :_ ":oltune
equivalent to the THOR Missile. Conditioning this bulk of propellants is a
.............. v ......... _ for r-Ou:_, in which or: a La Grangian coordinate
scale the pilo_s and the engines travel alternately in opposite directions, must be
countered. Consider that the booster shrinks about a foot when propellants are
put aboard and you begin to appreciate that the intra-w__hic_.,lar mot'_ons are not
of negligible magnitude.
Furthermore these vehicle stages need mechanisms for directional control,
and in the orbiter, for translation between orbits. The problems of the
At_-xiliary Propulsion Systems aad the orbit mal_euvering system, bofi_ of
which are fed high-pressure gaseous propellants from a iow pressure liquid
propellant ta,zk source, cannot be regarded as proportional to the thrust of
their cornbustors. Our experience v¢ith the generation of high pressure gas
in flight and the use ef gaseous combustion systen_s is extren_ely lind.lied.
And so there is a lot to learn, define, and verify about such 3yste:ns in the
next two years.
15
Finally both stages need to return to the launch/landing site. The orbiter
gets there tile long way, and can make it wif.hout airbreathing engines wii:h
sacrifice in maneuver and go-around capability at the landi."g site. The
booster, however, will have to bore its _,ay through 350 miles of atmospilere
to get back tobase. This job requires good airbreathing turbojet engines
of large size.
Until these systems work, the Space Plane, or Shuttle, will go no where.
Thus we can conclude that this conference will cover the most vital as
wel[as challenging areas of work necessary to the Space Plane's successfu]
operation.
To get on with it, may I promptly turn the business of runni_g this s]:c_."
over to Jerry Thomson of MSFC, who has bcei_ ci_iefly responsible for
org._ nizing it.
_°
16
Working Group Reoort
Jerry Thc, lnson:
As you nx::y have noted [rom the agenda thi_ Co.'./erence is divided iqLo
five sessio:16. "I'}_L}" are;
I. Main Propul,_ion #. Thon_bcn, Chairman
LI. Auxiliary Propul._ion E;. Jacobs. Chairman
III. Airbreathing Propulsion W. Stewar:, Chairman
IV. Auxiliary f:)_wcr Unit I). Beremand, Chairman
V. Cryogens C. Wo_,d, C!:;llA'm;*;;
The last _ession, Cry(_gens is not a respon,,,ib;l_ty of the Prop,d>,_c:,:
Wo:king Group; but fall.,, _mder the Operatiov_, W,,.vking Gr,.mp, chaired
by .Mr. Sam Beduingfield of KSC Mr. C. Wood heads the Cr_ogeu_
sub-group.
The last Propulsion Conference _pon,ored by t_:e Propul._;i_)n 'I'm hx(()|.,gy
Working Group was held at L, cwis Research Ceuter iz_ ,_ul'/. l_J70. A
greet deal of the mat_,ria[ preseqted at that ,',_[vrct_ce c_)','_:r_,d "wh._ w_.
were going todo;" our conference t,)day and tom,rr_,w will pr(,_(,,il ,n,aL
on "what has been done." I believe that those oi you 'a,))_, w,.r,, pre._cnt ,,l
that last Conference will bc ;lblt. to note lhal l-ll,4l)y .51Rniflt,ll_! ,_c_;),',cll1('))!',
b.ave bee,: made si,:ce la.-at .'t_y.
Before we gut on with the technical pre,,,entati,m I t_ou|d |ikt, _,) (_,vcr ._ ft._
charts to show the structure ,)f lhe W,)rt_,i))t; G_.,)t,p.
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S I_:SSION I
MAIN DROPUI.SION
J. THOMSON, CIIAII-,'MAN
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29
3 9 5 7"_'
"Fi_;AL RESULTS OF XLR-)29 Pf<OGRA.M _:
R.R. ATHERTON
PRATT & WHITNEY AIRCRAFT
TECHNICAL MA!qAGER
CAPT. R. PROBST
USAF, ROCKET PROPULSION LABORATORY
P1LISCE, I)EqG I'AC, i/ BLA.NX NOT "'' ;'_
31
FINAL llESULTS OF TIlE XLR12_2 I)ROGR;\,M
Pratt $. Whitney Aircra/tFlorida Research and Dcvolopment Center
Vvest Palm Beach, Florida
R. II. -Ytherton, Project Manager
Contract No. F0.1611-68-C-0002
United States Ai_" Fo,'ee- Rocket Propulsion Laborator:,"
PRE_'ZL" ,,.•_ l j •
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33
MIII
ABSTIb_,CT
This paper reviews the progress made on tlae XLR129-P-1 Reusable Rock-
et Engine Program conducted by Pratt _ hlaitn.ey Aircraft under Air Force RPL
sponsorship at the Florida Research and I)evelol)ment Center. The engine was
designed to provide 250,000 pounds of thrust in vacuum with a high area ratio
nozzle. Th,- size v:s selected to provide mozmin;ful data that could be scaled
over the ral_ge of lu0,00,') to 1,000,000 pounds thrust. The design of:-all eom-
t:onents and the engine system was completed. A full-so:de engineering mockup
of the demonstrator engine was also completed. "l'es'_s have been conducted on
the preburrter, the main case, and the fuel "c.ucbopump, which are
reviewed and discussed. The assembly of txvo po,verhcads consisting of these
components anti tests on these powerheads are re,dewed. The component re-
usabilicy experience in this program is also reviewed. Component programs in
specific areas with demonstrated reusability are: (1) wear life of critical mov-
ing components in valves, (2) fati_,,ue life in turl)ol,ump bearings, (3) creep life
of turbine blades, and t4_ low cycle fatiffue life of exhaust nozzle coolant pass-
ages. The design procedures and criteria t',at have been used for the.
XLI{129-P-1 reusable rocket engine are based on those that have been used
sttccessMlly for the design of commercial and military jet engines with proved
darabili_' and reliabili*y.
pgFEEDLN"
35
I I
t
SECTION
II
C C_N'I" I.:371"S
ILLI rSTI1 A TIONg .............................
XLR129- P-1 DI.:3ION_'I'IIATC)II t.:NGINI.: ............
PAGE
iv
1
,,,'_" "TOT , 1,'"_,
37
Iz.L U&'I'P A "[:,3NS
FIGURE PAGE ,
1 Demonstrator Engine Arrangement 2
Z Schematic Diagram of Fl._w 3
3 Full-Scale Engine Mockup 3
4 Preburne:" Test Rig 4
5 Preburner Mixture Ratio Profile 6
6 XLRlZ9-P-1 Deme_n-.trator Engine Program 7
7 Major Compment Plug-In to Main Case 9
8 Schematic Diagram of Internal Flow 9
9 Case During Testing 10
l0 Internal View of Transition Case 10
11 Tested Fuel Turbopump Assembly l 1
12 Test Points for Efficiency vs Inlet aI:d Flow 14
13 'fypical Flow Excursions 14
14 Turbopump Test Data Plot 15
15 Hot Gas System Mc_mted on Test Stand 16
16 Hot Gas System Rig Assembly Schematic 18
17 E-8 Test Stand 20
18 Sati:,factory HoopSeal Configuration !3
19 Main Chamber Oxidizer Valve Seal Rig 24
20 Typical Leakage Curve va Shutofl Cycles 2-5
P!_I;cI[Y_i>,'G l'.',Gi." L'!..',:,2K .',701' iYlL!gb
39
i:IC U[(l,_
;'.l
22
Z3
Z4
_5
26
ILLUSTRX TIONS
(ton't)
Roller Beari,_g Test Rig
Roller Bea.ing After Test
Tubes !_razed to Stiffener Blocks
Nozzle Tube Test Rig Details
Nozzle Tube Test Data Points
._tress vs Material Creep
PAGE
27
27
z9
Z9
30
31
I'RI_;CI_I.,\'G pA{_B _EA.UK NOT 7'T: ....
41
................. _ ....... _ -. .. _ . ":" - ' =_
[I. XLI{_Y9-P-! I)I.IM(_NS'IH.Xi_,I_ t N_;IN[7
The adv-:nce,_ D_'w'lopnat nt l_ro<ram (AI)P) for Pr:_tt _'. V,hil:_cv Air-, rzllt's
XI.I{I:2!}-P-1 licus:d)le I{wcl¢.et Engine was started ,m ¢; No'v_::nb-r Ibm,7 unchq-
contl'aet from I.he Air I.'orce l{ockct Propulsion L:lu_z'rlt,,rv. I!:e _vt'l-.qlI c0)-
jective of this arot_ram was to demonstrate the },e,'f Jrm:_nue :tn,] meeh:,nic;tl
integrity era 250K o.xygen/hydrogen reusable r,_cke_ en_line. This,!*.m,me:trat,w
engine operate s at hi.Kh chamber pressure and uses the sta,_ed combustion en-
gine cycle. The des:ign and demonstration characteristics f_,r the XI,RI2q-P-1demonstrator are shown in tal,le I.
N,,m i n,_ "rh ru s t
Minimum Delivered
Specific impulse
Ef[iciency
Th rottling Range
Ore r_dl Mixtu tx
I'¢:lt it, Ikmgc
Rated Chamber
Pressure
gngi neWcight0,Vith 75:1 Nozzie)
Expansion Ratio
Durability
Single Conti nuou s
Iiun Duration
Engme .gin rt s
Thrust Vector
Control
Cent rol Capahi' i t3"
Propell'mt
Conditions
Envi ramnental
Conditions
Engine Vehicle
250,000-1b vaem_m thr.,st with area rati,, qf 166:1.
244,000-1b vaCUtllll t}lrq.lst with ar,.a ratzo ,_t 75:1
209, 0_p0-1b s_a h.,vel thl'u:,t witla area ratio of 35:1
96_ of theoretical shifting Is at nominal thrust;
94% of theoretical shihing !s duric,,- throttling
Continuous from 100 to 20_ of n,'m_thal thrt_st
over tim mixt.:re ratio range
Engine operatron" frotll 5.0:1 to 7.0:1
2740 psia
3520 lb (with flight-type actuators and enginecommand unit)
3380 lb (less llight-type actuators and engilie
command unit)
Two-position booster-type nozzle with area ratios
of 35:1 and 75:1
10hours time between overhauls, lb0 reuses,
300 starts, 300 thermzfl cycles, 10,090 valve cy'eles
Capability from i0 seconds to 600 seconds
.Multiple restart at sea levei or altitude
Amplitude: :2 deg;l'_,>_'n • '_(I ¢I .... 1 ......
Acceleration: 30 rad.'see"
:3_ accuracy in thm_st and mixtu"e ratio at. nomimd
tiH'ust. Excursions from extreme to e_.'treme ip thrust
a';d mixtuce ratio within 5 s.:conds.
LO2:16 ft NI-'bil from 1 atmosphere _milingtemperatt/re to 180°R
1.112:60 :t NPSH from I atmosphere boiling
temperature t_ 45°R
Sea level to vacuum conditions
Combined acceleration: log axial with 2 g transverse,
6.5 g ,nxi_ with 3 g transver:;e, 3 g axial with
8 t,: transverse
1"he engine ",qtl receive n,:_ ext,:rna.1 power, with t;",'.:
exception of _ormrd electr,cal power a_,d 1,.5_)0-i,s_ahelium, from the vehicle
%
p._..,., > •_ _" "G P V3": "_,OT F I'-_'''_'7)43 pp.l:*.1 .,)L-N ' -
\
The general comp.ment arr,_nizcment o1"the dcn:or_strr, tor cng_r:t: i:-_zho_,ir. fi_,mro I.._,_ [lo'_,, s,-'h_m it/c is ,!:?-,vi_ In {iwdre .". In the ;_:;,_fe,.l c )nlou_tioP
cycle, the prebur_er proddces hot fuel-rich gases ,_nat lriv_, tqe turbines of thetur,;%um!,._, and ahese gases are burned in P'e mare hu:'ner cha:nber wit,_; thebalar._e of the oxygen to pr_duce thrt;st. A twn-I_,sitim_ translatinl, nozzle isused to decrease l.he overall stowed engin,, h,n_h. A fvll-scMe mockup of tl',eDemonstrator Engine was.made as shown in fi_ure _, which _hows the two-position nozzle .:xtcnded _,For economy, the two-posilion nozzb_ used in thismockup was only a partial ler.gth nozzle, not the full l-r:knh nozzle. )
",V-. _ ._,.:._ -; _., '
_-_ I_ "_,e_:_.2 _ :r;_ .._
_"¢ k" _ : ......%" t ' ....(, _._: ,:_C';r,_X ' i
. 1 \ .t
FIGURE I. DEMONSTRATOR ENGINE AIJ_AXGEMENT
44
" k... "y'f,
FIGL_[E _, }'L'LL-.qCALE ENG[._,F MOCKt'P
1. Component Test Results
a. I'reburner Test Rig
A test rig consisting of thepreburner,)xidizer valve, the iniector :rodigmiter, the housing and cooled liners, a_d a main burner simulator was assem-bled •_s shown in figl_re 4 and tested. Objectives of these tests were to o;,al:_.ate temperature profile, comb:tstion stability, combustion efficiency, simu-lated e_gine starting and durability of the hardware. Twenty-two tests wereconducted between August 1969 and April 195"0, and a sun, mar' of the test re-sults is presentcd in table II. These tests showed that thc temperr, ture profileon thc last series of tcsts is suitable for continued testin_ with the fuel turbo-pump in the "powerhead" rig. Dyna:nie combustion stability was also (lemon-strated. Combustion effieieney was demonstrated to be quite high and appro;,.eh-ing 100r.: in most cases. Engine starts were simulated and hardware durabilityfor the basic components was shown to be satisfaetn,.y. Figt, re 5 is a plot ofthe preburner mgxture ratio profile that was attained in a run at 75": thrust andat an equivalent engine mixture ratio of 7.0.
FIGURE4. PREBURNFR TEST RIG
46
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FIGI.'RE 5. PREBURNER .M1LXTI.RE RATIO PROFILE
48
III I
The XLR129-P-1 demonstrator engine program schedule is shown in
fignire _,. This 54-month program was divided i;,to five basic phases. The
first p|;ase, which was completed early in 1968, generated test and analytical
data to complete the technology n_cessary to design the engir, e and components.
During the second phase, which ended in the latter porti.on of 19G9, all the
components were dcsig'ned. The tJ_ird [ahame,which was the fabricatloi_ a_d
testing of corr, ponents was started, but was not completed becm:se all test work
on the cm_tract was terminated. Some of the most significant compm,ents were
fabricated and tested; these are de_;cribed in the following sections. The fourth
phase was to t:ave b_en the integration of the components into the demonstrator
engine and the test£ng of the demonstrator engine. The fifth phase was to have
been the definition of a flight engine that could have resulted from a subsequent
Engineering Development Program. These phases were not started.
L_._._ (Supporting TesLt
L[1N_v1967 19E8
FIGUIi I,: 6.
Design "i
Component l_'_elopmenl l
[ EngineTe,tj
E.cine Oefinition i]
I I I ,,I, I
1969 1970 1971 1972
CALEND_,R YEARS
XLItI29-P-1 DEMONSTI{ATOII ENGINF: PlIOGI_.AM
49
I II I
_ =,,
b. Main Case
Procurement of raw nmterial for" the outer m:|in ea_;e and l:d)ri('ation of
the case werf. started early in 19(;9. "I'he eoplanar intersecting sphere (lef_t!..qt
was selected over the canted-components version after desi,ql studies and model
testing ha, lbeen ce.nducted. 'Fais coplanar version was selecte,.l because it
offered the best e,mfi_,-ueation considering the inner duct design, cool!,_g,
thrust loacI hg, ndlmg, and assem_qy and manufactt, ring. l.'it,,-u_'e 7 show:; how the
major components plug into the main case. !.'igx,,'e s shows an i_ternal l],)wsehen::_tic of t!_e main t':ts_:, with the i':,el turb_q:t,l_i: _.llcI the _xidizer tttt'l_')l_t.ln_, p
turb,::c simulator inqtalled, l}m intersecting sptaere design was selected to
minimize we,ght _',nd m:Lximize confidence in the predicted strllerLttr:ll l_l:ll'gins.
Following fabrication, the case was successfully proof pre._sure/thrust tested
under conditions cxeeed,.'ng normal test levels. "lne case was tested under two
internal pressu;'e and thrust ratios; namely {1) Simulated preburner testing
pressure-thrusL ratio; lOOq- proof test load equal to ,t5')0 psig and (;:), t)O0 pounds,
and (2) simulated engine desigm pressure-thrust ratio; leOn. proof lest load equal
to .t500 pstg pad 375,000 pounds. I.'igure 9 shows the case during testing, w'th
stress coat to indicate areas of high stress. Figure 10 shows an internal view
of the transition case with the internal eenterlcody installed.
50
N
I
I
'h---_"'" _ '. "':..-I_,_ -_-_'\.;J_--_'_'- v _,_
FIGURF_ 7. MAJOR CO._IPON'ENT PLUG-IN 2"O MAIN CASE-
• /__ _
_,,,,,._'.'.'.'.'.'._-,,..,i;°., _ -'" : ........ f _._ _ v-• _ ¢,_,,,Y'- ",-, ._._i_-
FIGURE 8. SCHEMATIC DIAGRAM OF LNTERNAL FL_,W
51
_'_ " ® I !-_c"_i_!_
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• ' ",,,5,'_'_';_ "'_"/t" ;_ * '"
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I,'IGI.'RE 9. CASI,." Dt_'tllNG "I'E,STING
FIGURE 10.
iJ
INTERNAL VIF.W OF TI_..\NS:TION CASE
.SZ
2. l_MetTurbo|rompTest B.esults
The fuel h_rbopump is a single shaft unit wit5 two back-to-back centrifugalpump stages driven by a t_'o-stage, pressure-compou.'_ded turbine, The turbo-pump shaft is located radially by two roller bearings and ,_.xially by a double-actir, g thrust balance piston. The fuel turtx_pump delive:-s liquid hydrogen at aflow rate of 91.3 '.b/see at a pressure of 5654 psia at the design point (mixtureratio of 5), The tw_-st._e turbine delivers approximately .t9,900 horsepowert,) the pump and operates at a minimum inlet temperature of 1980°R and a maxi-mum hflet temperature of 2292°R at i00% t/unast. One o..."the fuel turbopumpassemblies that was tested is shown in figure 11.
FIGURE II, TESTED FUEL TURBOPUMP ASSEMBLY
53
Threefuel turbopumprigs weretestedfor a total of 488.7secondsrundurationandtheresultsof thesetestsare summarizedin tablelII. Objectivesof thesetestswereto determinethepumpperformance,thethr'.st balance,•andthesuc;:ionperformance. Flowexcursions_ere cor_uctedfrom 50%to100%of pumpspeed. Thepumpperformancewasbetter thanpredicted, thethrust balarcewassatisfactory,andthesuctionperformancewasalsobetterthanpredicted. Figure12showsnumeroustestpointsfor overall pumpeffi-ciencycomparedto tnepredictedvalues. The test efficiencies were higherthan the predicted efficiencies. Figure 13 presents two typical flow excursions,which show haw the pump discharge pressure varied as the flow varied duringtwo tests. Figure 14 is a plot of the fuel turbopump test data obtained at theconditions shown, wl,ieh shows how closely the test values approach the pre-dicted values. ,.
54
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FIGURE 12.
10q
80_
60_
40
/
20-
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XLR12<J-I'-I F.ef Pure Rig 33147.1/_
I
Q 24.000 rpe_
32,500 rT_:40.000 rpn_
_ 2,500 rpm47,000 rpn_
G 4_.000 rpr'*
, I i0.0_ 0.12 0.1 0.20 0.;-4 0.28
INLET UNIT FLOW, Q/N
TEST POINTS FOR EFFICIENCY VS INLET AND FLOW
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FLOW - gc,m (Thousands)
FIGURE 13. TY'I='ICAL FLOW EXCURSIONS
56
I000
l'I(;t I{E I-. "r[ I{[_()PI*MP I'I-;ST !)\ ['A PI,()T
_7
3. Powerhead Test Results
The hot gas system, was tested first during May and June of 1970 consist-ing of file preburner injector, preburner oxidizer valve, oxidizer purap simul:aor,fuel pump simulator and a nozzle back pressure plate mounted on the transitioncase. Figure 15 shows this hot gas syst,.'m mounted on the test stand. SLx com-bustion tests were conducted on this hot gas system which demonstrated thefeasibility of all the hardware with hot gas that simulated engine conditions. Thetot:3 test time was 42, B seconds. Table IV presents a suilnmary el the test dataobtained.
Following :he successful tests, the fuel turbopump was then substitutedfor th.: fuel pump simulator, and this powerhead or hot gas system rig assem--bly is shown in fig'ure 16. One of the priocipal cbjcctives of the second power-head was to qualify the fuel turbopump turbine with hot ga_ for engine operation.Rig operation up to 75,_ of ratc, d prebur_,_r conditions was attained. Table Vsun,.marizcs the second series of pcwer},,:ad tests. The hardware was in goodcon,litlon fuilowLng this series of t(,sts, these tests demonstrate'.] the feasibil-ity o." the integrated components, and parti:'ularly, the combination of thetransition case a_.d the fuel turbopump turlAne under hot _;as conditions. Fig-_'e 17 shov:_ a llow schematic of the E-$ ttat stand '.hat was used to test the
hot gas system or powerhead.
wIGURE 15. IIOT C_.$ S_._STEM ,*,iOUN'rED ON "rEST STA_ND
58
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59
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FIGURE IS. HOT GAS S}_T_'.M I_IG ASS£MBLY SCHL%_ATIC
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22.000 pl
195 I_ia
Flol, me_erk_
9600 psia
3375 wcf
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6600 psiJ
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9OO ;ll6600 psia
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Flowmeter
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Control Valve
Oxidizer Pump
Turbine Simulator
F[GURE 17. I_-8 "" ESI" STAND
PR _i]C_h_;GPAGI_ NLh\'K NOT FH_._f_
63
t. Component I_eusability
One cf tile specific life requirements of th( XLP, 129-P-1 reusable roel.a,ten_inc is a 10-hour Time Between Overhauls {'I'I:O). Lon_ engine lift' in eitherturboje! ,'ngines or rocket engines begdns in the initial d esi_;n of tile compon-ent.s and engine and is achieved through endurauce rdiability and durability
testing. At Pratt & Whitney Aircraft rclial,i_ity rind dependability are influ-ential considerations from the start of design until the eng6ne is retired fromsevvice. There are certain specific comlmnent life factors that have reeelvcde!ose attention on the XLR129-P-1 reusable rocket engine. These component!ire factors are: wear, fatigue, low cycle fati_._ae and creep.
a, Wear
Wear of parts in rubbing contact becomes a dominant factor in the designof suc'h parts as turbopumv., sha_ft-sea!s and shutoff valve seals. Pratt &_A2;itnt, y Aircralt has conducted extensive valve s,.,als tests under Air ForceContract to improve the life of these seals. For exam!.'le, the main chamberoxidizer valve, whieh is a butter;ly valve, incorporates a shutoff seal for theoxidizer flow to the mai,: burner injector. To accommodate this shutofffeatm'e a ca;:ted shaft wi_ an integral disk was selected so that an uninter-rupted disk :_ealing surface _muld be pt'ovided. Ineorporxtior of the shutoffseal in this valve eliminates the need for a scp:;rate shutoff wive between themain t'ha|niJer oxidizer valve :rod the main hurner injector. A test rig wasmade for the purpose of environmental e.durance tests on various designs "_ftl,e main chamber oxidizer valves, The tests were conducted by submergingthe v,"tlve in liquid nitrogen or liquid argon and pressurizing with I_trogen toin_ernal pressure of 50 to 6000 psig.
The valves were cycled at these conditions, al_ valve seal leakages weremeasured periodically. For the latter tests, liquid argon was seieeted for _-'::cryogenic bath to ensure that all. of the nitrogen leakage vaD_rized at the va_,external ambiert press;we. This allowed satisfactory steady-state leakagemeasaremenz accuracy for all data points. The objeettve of the leakage testbwas to obtain a seal design that would have !ess thzm 10 sees leakage. AWpical endurance test would consist of 10,000 shutoff cycles and 500 press_Jrecycles. A satisfactory hoop seal was developed with a eonfibmration as shownin figure 18. Tins hoop sea! was silver plated a_.d bad a 0.0".0 in. tight fit onthe disk. Figure 19 shows a D'picai disassembly of the main chambe,," oxidizervalve seal rig. FiDlre Z0 is a typical leakage cur_e versut_ shutoff cycles with,he hoop shutoff sea__s. As a result of these tests a satisfactory seal wasdeveloped which showed that wear could be controlled to attain a 10-hour ii.fe,consisting of 10,000 shutoff eyries and 500 pressure cycles. This type ofseal will be used on the main chamber oxidizer valve.
64
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SA'I'ISFACTORY HOOP SEAL CON FIGUI-bkTION
65
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FIGt'IIg 19. hiA_ {'HA,XlI3EIi OXII)IZEII VALVE SEAL HIG
66
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0.0015'0
i I' I
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ordeal at 50 psid
O Rig at Liqu;d Argo:l Temperature
(_ Rig at Ambient Temperature
AIL Rig at Liquid Argon Temperature
No Indicated Leakage
Z_ Rill at Ambient Temperature
No Indicated Leakage& |
4OOO 6OOO 80O0 10.000
S_'IUTOF F CYCLES
FIGURE 20. TYPICAL LEAKAGE CURVE VS SIIUTOFF CYCLES
: i
67
b. Fatigue
Fatigue, in ,_hich a part is subjected to a cyclic stzess below the elasticlimit for millions of cycles, becomes a dominam, factor in the design of certaincomponen,*.s such as propellant cooled bearings opcr_,ting under high load whereeach time the ball or roller passes a point on a bearing race a stress i¢ im-posed on the race and removed after the bearing or roller Las passed. At highrotational speed, with a large number of bails or rollers in the bearing, manymillions of cycles can be accumulated in a short period of time. A rollerbearing durability test program was conducted at Pratt & Whitney Aircraft, and55 x 96.5 ran, roller bearings were tested and e_aluated for use in the XLR129-P-1 fuel turbopump. These fuel turbopump roller bearings operated at a max-imum DN of about 2. 65 million mmx rpm. The beat'ings operatg in a liquidhydrogen environment that is provided by the propellant being pumped in thefuel turbopump. A test rig was made that has the capability of testing twobearings simultaneously at speeds up to 62,000 rpm with radial loads up to
2400 lb. 1.'i_ re 21 shows a cutaway of the roller bearh_g test rig that wasusecl. During Phase II of the atlvm_eed ,h,velopm.2nt pro._ram, a total of .':5hours of test time at 48,000 rpm was _eeumulatcd. These tests were con-ducted to evaluate the effects of roller length to diameter ratio, roller end toside rail clearance, internal clearance m_r.I roller crowning on roller t,ndwear and bearing durability. During all these te:;ts, a 1700-1b radial load wasapplied to the load bearing rcsulthlg in approximately 1-t-t5-1b radial load onthe reaction bearing. Five bearing configmrations were dev,Aoped that passedthe 10 hour goal test duration at the design operating conditions. "rher(.fore,it can be concluded that fatigue on the race of the bearing can oe eontrolle,.'through proper design to attain the desired 10 hour life. Figxtre 22 shows aroller bearing alter 15.3 hours of test. The most promising bearing config-uration tested used stainless steel (AMS 5630) inner race and rollers: anouter race guided Armalon cage; a steel alloy (AMS 6265) outer race, singlecrown, L/D = 1.0 rollers with 0. 020 in. roller end to timer race, side railclearance &aid 0.00.t3 in. rmgative diametrical internal clearance. Bearingsof tkis configuration will be used in the XLIl129 fuel turbopump.
68
GN In
L- - • "-_,_
. Turbine Ball Bearing--J
c Radial Load Test Bearings
Fuel Out _" _- Fuel In [_J
FIGURE 21. ROLLER BEARLNG TEST RIG
X
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FIGURE 22. ROLLER Bt::ARII_IG AI_ TER TEST
59
c. Low-Cycle Fat __."ue
L.,w-cycl_ fatigue is ,also a major factor in the design of rocket enginec,',,:_t_',nents. Every time the en_ne sta_'ts and reaches full power, largeth.-.v,,_ai ¢.r_dients are produced on the components. The component areas ex-postxt to the high temperatureq heat rapidly and expand rapidly. Those areasthat are not e.x'posed to the high temperatures do _ot heat as rapidly and,therefore, ,,x'pand more slowly. These large temperature gradients result inhi;;h stresses. The thermal process reverses d_ring the shutdown. Some ofthe comi>oncnts that heat_x_ rapidly during engine operation now cool rapidly,and many ot the components that heated slowly now cool gradually. This re-suits in stresses dropping to low levels. Some may even reverse producingeompressiv,: stresses. Therefore, with each ignition, acceleration to power,shutd3wn and cooldown, the en_ne is subjected to large stress variationsconsisting of both high tensile and compress:re stresses. Low-cycle fatigueis characterized by loads imposed that strain the materiM beyond its elasticlimit into the plastic region. This type of operation occurs, for example inthe nozzle coolant tubes that have a large thermal gradient through the wallthat cy:les on and off each time the engine is started and stopped.
The primary nozzle of the XLRI29-P-I engin_ consists _f two tubularhi'at exch,'mgers. The material selected for the coolant tubes was Inconel 625(A.',IS ._59(.)). This material can be iormed into tapered tubes and is easilyb,'aze, i. The primary nozzle coolant tubes were sized by ,analytical consider-ations of temperatures ,'u,d the fatigue cycle life requirements. The tubes hadto be dcsik_ned for a life of :_,00 operating cycles and a 10-hour total operatingtime. The 30,)ocycle requirement is actually more severe. Thi_ had to bedefined from low-cycle fatigue tests, which were conducted by Pratt & WhitneyAircraft. Inconel 625 tubing with nominal 0.322 in. outside diameter andnominal 0. 012 in. wall thickness was obtained, ,rind 12 sample tubes were fab-ricated. Twelve Inconcl 62G 1-in. diameter bars were also fabricated, and thetubes were brazed to tbe stiffener blocks as showr_ in figure 23. The testequipment consisted of a copper induction heater coil, liquid nitrogen supplytank, plexigl:_s atmos_,here container, helium supply bottles, radiation pyrom-eter, timer, cycle cou_}:er, thermocouples, potentior.eter and pressure gage.Fil.nare 2z, shows the details of the test rig with a s,)ecimen supported in theliquid nitrogen trough and the method of induction t eating. The tweiv.: testspecimens were pressurized in_ernally with helium gas to the hoop stress re-quirements. The thermal strain was induced by placing the irduction coil ap-proximately 0.25 in. about the tube crown of the specimen and placing thestiffening block in liquid nitrogen. This method was the most representativeof gctual and engine condltions that the laboratory could accomplish. Thecrown of the tube was heated, and the temperature was determined by radiationpyrometers emissivity. Two thermocouples on each edge of the heated tubewere used as a check. An optical pyrometer was used to verify the tempera-ture readings of the radiation pyrometer at the 1880°R test points. Therefore,the temperatures recorded are fairly accurate. Each thermal cycle was con-trolled by an automatic timer _,v.ndcounter. The thermal cycle consisted of60 seconds total time of which approximately 30 seconds were required to attaintemperature, 15 seconds were at the test temperature, and 15 seconds tocool down to the liquid nitrogen temperature (159°R), The te.qt data pointsare shown in figure 25. The number of cycles reclulred for rupture of the tubesamples at the stress level and temperature indicated are shown in paren-thesis below each series of data points. From these data it can be concludedthat at the engine design pomv of 29,300 psi tube stress level, the aG0-cyclelife can be att_Aned if the maximum tube wall temperatures are maintain_.dbelow 1300"F.
?0 ©
FIGURE 23. TUBES BRAZED TO STIFFENER BLOCKS
/ , TO INDUCTIrON
,.) i f HEATER
)
C_RAMIC
}N ,'%.*L ' '_OR"\
[ \ ¢ .-- INDUCTION COID
c • _CI_['N TUBE -_ \ _ ¥
| STIFFEN l 1 | , I LAFFLE
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_ _-- DKAIN _ "
FIGURE 24. NOZZLE TUBE TEST RIG DETAILS
71
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FIGURE 25. NOZZLE TUBE TEST DATA POINTS
?Z
Figure 26 shows a relationship between the stress and how long it takes this
material to creep 1,% at 1800'F. Simil9r data are available at other tempera-
tures. Through proper design, and by maintaining the stress levels to attain
a one-percent creep life, it is possible to attain e_remely long life on the
turbine blades and vanes in jet engines. This technology is directly applicable
to the att_nment of long life on the turbines _d vanes of the reusable shuttle
rocket engine.
d. Creep
Creep of components operating in a high temperature environment will
occur, as it does in turbojet engines, particularly in tm'bine blades an2 vanes.
Therefore, experience in turbojet blade and vane al|oy development wtlI be of
direct benefit _o the reusable shutOe rocket engine. For example, on advanced
military jet engines, it has been demonstrated that it is feasible to attain a
3000-hour total life for cooled tu.bine blades with an average combustor Lem-
perature as high as 2400°F. One of the important developments that has con-
tributed to this long Iife is the us_ of a direction,_lly solidified high temperature
alloy for the tu_bi_ie bladp_. Tb!s particular alloy offers maximum strength
and excellent fatigue and creep resistance at metal temperatures ap to 20000F.
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FIGURE 26.
1800 "F
100 1000 10,000
TIME - hr
STRESS V$ MATERIAL CREEP
73
, _ __71 - 295 7 8
"_WO-PHASE FLOW LH z PUMP INDUCERS"
J. A. KING
ROCKETDYNE
TECHNICAL MANAGER
C. D. MILLER
MARSHALL SPACE FLIGHT CENTER
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"LOW FREQUENCY ANALYSLS OF _OCKET ENGINES
USING COMPRI'ISSIBLE I_ROPELL_NTS"
J. M. McBRIDE
AEROJET
PRI_i.T.'iN(I PA'3i2 i,:,A.'(K NOT P"..._.'":_,,,._,
TECIINICAL XS"_NAG ER
R,_ ,_a ]._,l 1MOI','D
MARSII.ALL SPACI: FLIGtlT CI£-NTER
!
N?1.29584
LOW FREQUENCY ANALYSIS OF
ROCKET ENGINES USING CO,'_PRESSIBLE
PROPELLANTS
k .
uy
J. M. McBride
AEROJET LIQUID ROCKET COMPANY
Z59
_+.= . ., ..
................. I "P II •
ABSTRACT
A low frequency stability model has been developed to represent the
case where one or more of the propellants used in a rocket engine is a com-
pressible fluid. Results for two oxygen - hydrogen rocket combustors are
discussed. Also included is a discussion of the predictu4 effects of mixture
ratio and propellant temperature.
INTRODUCTION
The motivation for developing a model which treats the particular case
of compressible reactants was an inability to predict the trend3 noted in
stability of a gaseous hydrogen - oxygen rocket engine using the conventional
low frequency models such as the type discussed in Reference I, which were
specifically developed for liquid propellant systems. Experimental data such
............................... n_ .... =_L, 6 the fael injector pres-
sure drop reduced the eta_ili=y of _aseous propellant rockets. The trend
toward reduced stability margin with Imcreased injector stiffness is Just the
opposlt_ _o predicted trends of curren: models.
Previous treatments of =he problem of low frequency Instab_llty gener-
ally postulate a model based on the conservation of mass and assumed that all
signilicant energy release is coincident with conversloe of llqui_ r_actants
to gaseous products. It was also assumed _hat th_ dense liquid propellants
occupied...........a n_g_*g_ble volume on ent¢rlng the _._uer,-_..... u_aking it possible to
ignore their e_fect on cha=ber pressure the instant they enter the chamber.
It was not until the reactants were gasified and reacted that their presence
was felt, and these two events were assumed to occur slmult_neous!y.
The analysis presented here takes a somewhat different approach in that
the conset_atlcn of energy law is invoked as a mean_ of determining the sta-
bility of a system. _y so doing, it is possible to separate the e_ect
gaseous reactants have on chanbe: pressure the inste::t they enter the chamber
and their effect _hen they are converted to products.
1 I_t.Ct.I_L\'G ._\I(;L.BL _.\_,- NUT FU_,,..
261
___---"_" ....I -I_ '- -."=:: ...... -_---_e-r- ........ , -_ --
TECHNICAL DISCUSSION
GENERAL -- Consideration of the low frequer:cy feed system coupled sta-
ability charact_rlstics of a rocket engine using _aseous propellants intro-
duces some analytical complexity beyond that of an equivalent liquid system.
However, of more importance are the questions concerning the stability of such
a system as compared to a liquid propellant system. The questions are prompted
by the fact that the gaseous system generally requires larger injector orifices,
which improves acoustic coupling between the feed system and the combustion
chamber. The injection velocities are higher with consideration being given
to velocities a_proaching the speed of sound of the injected gases. It is
also interesting t¢ note that the tuned gas cavities, such as resonators and
quarter-wave tubes, have been used to stabilize an engine. If such were the
case with a gaseous feed system, this would imply that it would he desirable
to maintain good coupling between the feed §ystem _nd ti_e combustion chamber--
an approach Just the opposite to that t_ken for liquid systems.
The basic difficulty in developing a stability model for gas - gas or
gas - lJq,_d combustors is the problem of separating the effect of gases as
they enter the combustion chamber, and their effect w:,cn they ccmbust. The
propellants could react at the instant they enter the chamber or after they
have been in the chamber for suff!c_ent time to r_ix, d_ffuse, or reach a
critical temperature. In order to overcome thls difficulty, it was necessary
to use both the conservation mass ancl ener/'y equa._i,)ns to develop a character-
istic equation for determining the stability of a s[,stem. This approach
allows for an acc)unting of the energy added by the gases as they enter the
chamber separately from the release of chemical e_ergy whun they react. The
gases, as they enter the chamber, are cohslder_d to be isentroplc perfecl
Eases, posse_slng both Internal ar:z flew energy. The enthalpy of the gases
includes only th_ internal and pressure, or _otential, energy of the fluids
an_ not the rhemi(:al _ncrgy associated with each propellant in %oin_ from
reactants to products.
----_ Z6Z
]
The chemical energy released by the reactants is simulated by a heat
source located :.n the chamber. The quantity of energy released per ib ofm
reactants is th_ dlffercnce between the energy levels of the fluids entering
and leaving the chamber. In this way, it is possible to avoid a detailed
description of the intermediate processes.
In addition to the unsteady energy release associated with the per-
turbed flow, a second source of energy is included which is developed using a
linearized Arrenhius rate function to describe this source of energy. The
inclusion of this source turn was prompted primarily based on trends observed
in experimental data from gas - gas rocket testing. A block diagram of the
entire model is :;hown in Figure 3.
UNSTEADY FLOW EFFECTS -- StartJng first with the fluid entering the
chamber, there are three sources of unsteady energy flow into the chamber
associated with each of r_,_......._=_o_._.._ . _a,e first i_ that energy convected
in with the perturbed flow which is given by
E' = _l' e (i)
x I x x
The energy per unit _ass of the fluid, ex' is given by
V"
e -if+ xx 2y
(2)
The perturbed flow can be described by a feed system admittance, Yx' and the
chamber _r(]sure, P' , where Y is defined asC X .... -..._
W !
y m _-x P'
c
(3)
f
Z63
The feed system admittance, Yx' is, in general, complex and a functlon of the
physical prop,_rties of the propellant, the pres3ure drops in feed system, and
the dynamic characteristics of the feed lines upstream of the injector.
In order to make the analysis as genera" as possible and yet not so
complex, as to make interpretation of the results impractical, two types of
feed systems ere considered. The first feed system considered is one which
is terminated upstream pt the injector by an "open line" -- the open line
being representative of a feed line which terminates in a propella_t tank
having a diameter several times larger than the feed llne. Th_ second type
of termination considered is a closed llne, which is indicative of a feed
llne with a sonic flow control nozzle era high pressure drop point upstream
of the injector.
The starting point in both cases is the acoustic wave equation and the
momentum equation. For the case of no mean flow effects and no distributed
line pressure drops, the wave equation is
andthe momentum equation is
Transferring the above equations to the Laplace domain, it can be shown that\
the admittance for a constant area duct is given by
where G is the admittance at L - O.XO X
(4)
5
Z64
h
In order to make the feed line representative of an actual rocket
engine feed system, the system is assumed to be as shown in Figure 4, with
tile restriction in th_ line representing an injector orifice. The assunption
of an "open" (G×o . _,) and a "closed" feed line (Gxo = O) represents the two
possible extreme_ encountered in a rocket engine. Only one of these cases,
tile "closed" feed llne will be discussed.,
CLOSED FEED LINE TERMINATION -- The closed feed llne termination, in
the ecoustic sense, is one where the velocity perturbatioha are zero or very
nearly so. Under this condition, from the definition of admittance
U !
Gox = p'
it can be seen that G = 0 and Z = =, Flow condition3 which approach aOX OX
close/ end or solid wall would be the position Just downstream of a sonic
plane, a high pressure d_.qP orifice, or a normal shock. All of these condi-
tions can exist in a _aseous p opellant feed system.
From Eqoat!o_ (4), it can be seen that, for G = 0.ox
V !
•, i Tan _ L' - - a xy
Pxy g Oxyaxy xy
(5)
In ooin_ frot', rh,* f,._rl14n_ _,, :),o 4.4_^_^. ^.4f¢^_ _._^- _ _--- -- . -"
large pressure drop. Using _he orifice flow equation (Ref 2).
W2xV2x
PIx - P2x = 2 A2x g (i + Kc)
WIxVIx
2 Alx g(6)
?.
Expanding this eq_atlon in terms of its steady-state and perturbation quanti-
tles and incorporatlng Equation 5 gives
D
P2× -I K +i..... c_!__ - _ ix
!
l;2x I+K A2x V2x 2 Alx
g [1 ---f-- .x
-2 {g + W]x }alx 2 Alx (Glx)
(g Vix + Olxa_ixGlx) Alx
(7)
Writing the equation of mo_lon for the orifi,ce gives
Z A u.) d W2x111 -- - " _-- --_1_ = "_2x r3×J a,_x dt (9_ g _ ( g ) _ (8)
which, when linearized and only the perturbed quantities considered, gives
when _s transferred to the. Laplaclan domain
pl nI
3x S _ _2x
W' g W'2x A2x 2x
If it Is assumed in Equation (7) th,-_:over the distance,
(9)
then
W' (i0)2x = W3x
Z66
which gives, from Equations (7) and (9), for S = :._
I _.
P' K +i3x -I c -- %'Ix
%4-T- = I+K -- V2x 2
3x 22 A2x AIxg [i + _-_ M x ]
_4 WIx ]
alx {g + 2--_ix (glx)}
-- --2
(g Vlx + OlxalxGlx) Alx
io£
g A2x
(ll)
Letting
pt
3x i
W_., Yx(12)
The flow energy convected in with the pezturbed flow becomes, from Equation (I)
='x-I ex Yx P'3x (13)
A second source c,f perturbed energy entering the chamber with the reactants is
that convected in with the mean flow and is given by
E'^ = W e' (14)XZ x x
The perturbed energy per ib of propellant is given by perturbing the energym
eqaution and assuming constant heat capacities
+ p, (\ + v_)2x x
+e' _C (_+T_) +--+ (15)+ p, 2g
X X VX _X X
Z67
and assumirg the gas is isentropic and applying the _mall perturbation restric-
tions gives for the perturbed energy e'x
P' V V'e' --y-I [_ _£+_E___xx _ x_. g
c
(16)
Noting V' is the perturbed velocity at the injector orifice exit, a relation-x
ship for V' _:an be derivedx
w' = v' +Vx x x) A° (17)
t
the. perturbation velocity, Vxy, is given by
A g
Vx, = l___OxAx [Yx Xa 2xx ] P'xy (181.
Substitution of Equation (18) into Equation (16) gives
_x-i V P V A g) P'= ('*x _- ] --
x _'x x °x g A a Px x c
and finally substitution of Equation (19) into (14) gives
¥ -I V P V A g P'
E' - % [-_x _ +--5-x--q x 2x .)]_2 x _x Ax g (Yx a gx c
(19)
(20)
The third an_ most important, in terms of the amount of energy added
by the incoming reactants, is that energy released when the reactants are con-
verted to products. Previous analysis of this type (Ref i) have assumed that
the energy released by the reactants can occur at times oth_.r than when the
268
reactants enter ti_e chamber. This time deldy, which is generally referred to
as the total time lag, is intended to physically represent the time required
to complete the precombustion processes such asr_apo_ization in the case of
liquid reactants, gas phase mixing, heat-up periods, or any of the many pre-
combustion processes associated with reactants whica are not premixed. This
synthes_s is devised to avoid the complexities of mechanistic representation
of processes, _hith are, in general, difficult to d,_termine in a rocket engine.
This assumption is retained in this analysis. This analysis deviates from the
approaches of Keference I in that the perturbed reactants flowing into the
chamber are assumed to comSust independently of each other. This assumption
wa_ mace for two zeasons: the first being that it would seem physically mole
rep-'esentative of what could occur rather than assuming the two perturbed
flows concurrently arrived at th,¢ right time and under the r_ght conditions
to react with each other. _he second reason was that the previous approach
did not ;_how sufficient gale in the system tc sustain a low frequency insta-
bility i_ a number of rocket engine systems where low frequency instability
had been observed.
Using the abuve assumptions, the following representation is developed
for the energy released by each of the reactants:
E' " (; W' e (21)X3 X X C
where G is an amplification factor resulting from the fact that the perturbed
reactant flew reacts in its s:eady-stata counterpart and the net perturbed gas
generation re¢= is =he sum of the perturbed flow plus the amount of its
counterpart with which it reacts, which is assumed to be available from the
steady flow. The degree to which each perturbed reactant is assumed to inter-
act wi_h the steady state reactants is dependent on where the steady state
mixture ratio is relative to the stoichlometrlc m_xture (SR). If the steady
combustion is fuel r_ch, all the perturbed oxidizer is assumed to react at the
269
stolchiome_rlc mixture ratio, SR, and only a portion of the perturbed fuel for
which ox_dlzer is available. It is assumed the amount available is W' =of W_ R
which gives
(sk + i)C =o SR
and
RCf = SE (SR + i)
If the steady state mixture ratio results in oxidizer rich combustion
(SR-+ 1)
R
and
Gf = SR + I
The energy release as_orla_-d w_rh each of the -e_-t_-_s is *_* associated
with the energy release at the Stoichiometrlc m_xture ratio AH c giving
E' = G W' AH (22)x3 x x c
Referring to Equation (12), W' is given byg
P' (tC -- TX)
W'x (t - _x ) = ¥x --p (23)c
where z Is the time delay between injection and combustion of particularX
e!ement of propellant. Substitution of Equation (23) into Equation (22) gives
P' (tE' = G Y AR c - Tx) (24)
X g X C
C
Z70
TEMPERATURE SENSITIVE COmbUSTION -- Resul_ of analytleci work indicate
(Reference 3) there is a dependence of engine stability on mix_uze ratio.
Attempts to show this type of trend using only the representation for the
-_.ergy added by combustion Just discussed were unsuccessful. For this reason,
a second source of energy of combustion is included based on reaction kinetics.
This particular mechanism was selected primarily because it i_ physically
plausible with gaseBus propellants, and the experimental and analy<!cal work
indicated that deviation from stolchlometric mixcure ratio tends tc degrade
the stability of a gaseous rocket engine.
_he approach used in developing an unsteady energy releas_ rate is to
start with an Arrenhlum rate function based or_ an assumed flrst-order reaction
given by
i
C
i..
dc I N
-- = n cj Kdt J =i
_here K i_ given by
(25)
E /RT
K = B e a ;26}
B, the frequency factor from the kinetic theory of gases, has been shown
to be a function of the reantant3 involved and is 8!so proport!onr! te the %._---
perature of the reactants; thus, for a given reaction
A = B T _
where a, in general, vazles between O and i and is commonly as_u_e_ to b- e_l
to 1/2 and B is a constant and a function of the reactat_ts.
ZTl
". /
If it is assumed that the ma=s fraction, YI' of each constituent is
constant within the chamber and noting that the concentration C i is given by
Yl Pc(27)
C 1 = :_WI
then suLs=ftution of Equation (26) _nd (27) into (25) gives
Yl dOc N yj NF TI/2 -Ea/RT c_ o e
MW I dt j=IMW_ c
or for a two s_ecies reaction
do c _c 2 / Ea/RT_=_B T 1-2 e C
dt MWC
(28)
tDefining Qc as _he energy release rate and using small perturbation assump-
tions to expand Equation (28) gi_'es
Q¢ ,,_ 20' (Ea. I T'
C C
(29)
who._ Q Is glver_ by
- _ [nHc]
Using the i¢_ntropic relationships, Equstlon (29) becomes
_a: has been developed thu= far are _malytical expressions for the flow and
internal energies ef Lhe reactants entering the c:_amber and the unsteady energy
Z7Z
t \
released when the raactants are converted to p_oducts, all of which are
expressed as a function of the perturbed chamber pressure. What remains to
be determinea is the unsteady energy being stored in the chamber and the
unsteady energy leaving the chamber via the sonic exhaust nozzle.
t
Considering first the energy leaving the chamber, EN
(31)
From Equation (16)
,v __, ,I
, = _-1 _* _. v veN y i__-- + --
F g(32)
where the (_) is used to designate conditions at the sonic plane of the chamber.
From the isentropic flow relationships and for small Mach numbers in the com-
bustlon chamber it can be shown that
p!
wT t yC
Noting that
_, .i ___ ,I)W_ ffi(p V + V p A t
!
The following relationship can be developed for EN
3¥c-I [ I P'
(33)
(34)
(35>
273
Considering n_.xt the rate of energy stored in the chamber
dt dt c c c
dE .. P'
c d - y-__lh + _ V h c) .._cdt ffid--t(p Vc T c _ c p
c
From the conservation of energy
(36)
(37)
d E
c (38)E I + E' + E' + ' + ' + v _ =ol 02 03 Efl El2 Ef3 q + Qc dt
Expressing Equation (38) in the Laplace domain gives
Z'ol (S) + E'o2 (S) + E'o3 (S) + Efl' (S) + El3' (S) + Qc - E_(S)
= s E (s) (39)c
where Ec(S) is the initial perturbatioa energy in the cha_:ber.
Fro= the above equations, the characteristic polynomial, K G H (S), can be
determined, and _he Nyquist criterion used to determine syetem stability. In
conventional control loop notation, the various components of the system are
shown in Figure 3. It should be noted that the system, as treated in this
analysis is a positive feedback system with a characteristic polynomial ratio
in the form
(s) = z c (s) (40)R I - K G H (S)
Thus, the Nyquist instability criterion is that the gain must be greater than
one and the phase equal to zero degrees.
Z74
RESULTS
COMPARISON WITH EXPERImeNTAL RESULTS -- T_ble I is a !1st of the experi-
mental and predicted results for two rocket engine combustors using oxygen -
hydrogen propellants. One of the combustors which was the pewur source of e
turbopump operates at a mixture ratio in the range of 0.B0 at chamber pressures
in excess of 2000 psia.
The agreement is quite good in terms sf the resulting stability char-
acteristics and the predicted frequency of instability. At each of the fre-
quencies listed the Nyquist criteria of zero phase was satisfied. The fre-
quency at which the gain is greater than 1.0 would be the observed frequency
of instability.
The second grgup of test results were obtained from a oxygen - hydrogen
engine which operated at nominal chamber pressure of 500 psia over the mixture
ratio range of 3.7 to 9.8. The _gceement between ex_erlmental and analytical
results are also quite good in this case.
The actual system pressure drops and propellant temperatures were used
as input to the model in making the predictions listed.
In all cases evaluated the time lag between the oxidizer _nJection and
combustion was assuaged to be O.001 seconds. For th_ fuel the time lag was
assumed to be zero.
It was also necessary to assume an effec=ive activation energy. A review
of the literature Indicated a value In the order of 25000°R was reasonable.
This value was used for all cases analyzed.
PARAMETRIC SIUDY -- In evaluating the results from the analysis two
versions of the model were evaluated. This first included all energy terms
with the exception of the energy addition due to a non-unsteady rate of
reaction given by Equation (30). The second version included all terms.
Several parametric studies were made in each case.
275
For the version without 'he rate equation b,c_uded, effects of =z._ture
ratio and propell._rt inlet temperature were evaluat_. Ty;!_al ol the :e,iults
obtained for a rocket engine havicg th_ desi_n c,ara._rl. _i _ , and using the
propellants indicated in Table II are shown I., Figures 5 and 6. 7rc_ these
results, three sigrlflcant trends were noted. The _Ir_t, wh|ch i_ ind_cate_
by the results shown in Figure 4, is that low frequent7 stability _pr_ve_
witi_ a decrease in mixture ratio. Improved stability i_ also _bt:Jinfd by
increaslng propellant temperature while :alntalnlng a constant InJe_.tor pr_s-
sure drop, the Imprmvement being m_st notlceaole at l_)w mixture ratios where
the propellant enthalpy Is _gnx;h:ant re]atlve to thtt added t7 cc,=bustlon.
The parametric studies of tLe same tucker _r_gi::e wlth _he _ate fuDztlon
included were identical to the prev_ous resul_s with the e_cept_on that the
trend with mixture ratio 4as revers:d in that stability dec eased w_h decreas-
ing mixture ratio (Figure B) an_ the effect of 9ro_e]_,nt _.]_t is lot _s pro-
nounced as can be seen from Figure 9. _s with =he ca_e wlth no cgmbus_on, the
effect of propellant temperature _s more pronounc_.d at low mlxtur¢ ra_ios.
During all the above parametrJc studies Injector [.res_ure dre_ wer_ held con-
stant which weuld b_ representative of a design study of a _ixed pressure drop
system.
A parametric _tudy was also made for t_e ca_e w,ete the _ctor design
_as fixed and the propellant islet temperature yes ail_w_d to vary. F?r th_
case the increase _D nron_11anr _-,p_-_,,_- - .... _'_ _ _nc .... '
sure drop. Two conditions w2re considered, one where _he fuel and ogidlzer
time lags were equal, and the other where the _uel ti_e ia 4 was zero. The
results are shown in Figures 9a and b respectiv_!y alon_ with _he ccnatant 'P
cases. Figure 9b is reDres,_ntatlve of the _xperl_ental res_ts dlscuc_ed _re-
viously in that system stabil_ty was decreased with Increa_l_g fuel tnJ_:t_r
pressure drup.
275
277
s
REFERENCES
1, Su_nerfield, M., "A Theory of Unstable Combustion in Liquid Propellant
Rocket Systems," Journal of the ARS, September 1951, pp 108-114.
2o Rohsenow, W M., Choi, }i. Heat Mass and Momentum Transfer,
Prentice-Hall Inc., Englewood Cliffs, New Jersey, 1961, pp 67-70.
_o Culick, F. E., Stability of H:[gh Frequ_ncy Pressure Oscillations in
Gas and Liquid Rocket Combustion Chambers, MIT Report 480, dated
June 1961.
278
NOMENCLATURE
a
A
B
CP
CV
C
e
E
g
G
h
i
k
K
g
L
N
P
q
R
S
T
t
V
Speed of _o,md, in./sec
2Criflce area, in.
Kinetic theory frequency factor
Constant pressure specific heat capacity, in.-Ibf/ibm-°R_.
Cons=ant volume specific heat capacity, in.-ibf/ibm-°R
Concentration
Specific energy, tn.-lbf/lb m
Energy flow rate, in.-ibf/sec
2Gravitational co_stant, lb -ft/lbf-s_c
See Equation 16
Enthaipy, in.-lbf/lb m
Imaginary
Orifice flow constant
Rate constant
Or!fi(:e length, in.
_eed llne length, in.
Integer
Pressure, lbf/in. 2
Rate of heat added due to reaction, in.-Ibf/sec
Mixture ratio
Laplacian operator
Temperature, °R
Time, sec
Velocity, in./sec
Z79
NOMENCLATURE(cont.)
W Massflow rate, ib /secm
Y Feed system admittance, Ibm-ln.21_bf -sec
y Mass fraction
&P Pressure drop, ibf/in. 2
3.0 Density, ib /in.
m
Specific heat
T Time lag, sec
u Rate equation constant
Angular frequency, rad/sec
- Steady state quantity
' Unsteady quantity
* Sonic plane value
Subscripts
a Activation energy
c Chanber
v Constant volume specific heat
p Constant pressure specific heat
o Oxidizer
f Fuel
T Total flow rate
N Nozzle
Z80
/
TABLE I
PREDICTED VERSUS EXPERIMFNT RESULTS
Stable Tests
Predicted
F__reR
935
935
750
780
760
770
73074C
720
Gain
10
i0
0 78
0 86
0 78
0 76
0 79
0 87
0 63
Stable Tests
Predicted
Freq
850800
7407o0
770
760750
760
750730
750760
760
750780
760
770790
Gain
0.57
0.49
0.80
0.7_
0.77
0.81
0.88
0.73
0 o,._,OU
0.86
0.84
0.82
0.87
0.91
0.76
0.950.81
0.74
Turbopump Combustor
Unstable Tests
Predicted
_rr___ Gain
Observed
Freq
1470 1.95 1500
840 1.27 875
810 1.34 770
850 1.40 1250
950 1.02 950
950 1.02 950
940 1.03 750
850 1.24 800
500 psi_ Engine
Unstable Tests
Predicted
[ceeq Gain
Observed
Frec
750 1.0 800
75O 0.98 _00
760 0.98 800
04U l. IJ 740
720 0.9] 800
Z81
!
TABLE II
CHARACTERISTICS OF PARAMETRIC STUD'/ F_GINE
Effect of Mixture Ratio
Injector Oxld, APO
injector Fuel, gPf
Chamb[r Pressure
Throat Area
Oxld Orifice Area
Fuel Orifice Area
Fuel Temp
Oxid Temp
Propellants (Fual/Oxld)
45
70
500
5.6
Variable*
Variabl_
500°k
500°R
tlydrvgen/Ox),gen
Effect of Prop+=llant Temp
Same as Above Except
Mixture Ratio
Fuel Temp
Oxid Temp
5,0 and 1,0
500=R to lO00°R
500°R to 1000=R
*Varied to keep &Po and APfconstant
282
z_Pf
-g---c
0.3
0.]
Oo
UNSTABLE
0• 0 STABLE
0•0"0
Oo
0"2 ' '0.3 0.4
Figure i. Turbopump Combustor
• UNSTABLE
0 STABLE
0.3
O.l
0
0
• I UNSTABLE
0_0 STAgLE
O0
0o%
i "'1I -- i
0. I 0.2 0.3 0.4
Figure 2. 500 psia Combustor
283
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285
r
GAIN
MAX, GAIN
1.0
0.9
0,3
O.7l
0.6-
0,5 -
0.4
/,'1 , I .t, I I |
1.0 2.0 3.0 4.0 5.0 6.0
MIXTURE RATIO
NO TEF_, SENSITIVEBSI.!BUSTION
_c P = CONST
PROPELLANT TEMP. : CO_IST.
F_gure 5. Effect of Mixture Ratio
GAIN
MAX. GAIN
l°i 0.9
0.8
0.7 R = 1.0
0.6
t , t ' I !
500 600 700 800 900 lO00
HYDROGEN TEMP OR
AP
_c = CONST.
NO TEMP. SENSITIVECOMBUST!ON
Figure 6. Effect of HydroBen Temperature
286
:/
1,0
0.9
0.8
0.7
°'°I0.5
0.4 I! ! I g
l.O 2.0 3.0 _.0
MIXTURE RATIO
QC
TEVP. SENSITIVE COMBUSTION
,_,P
" p-'_ = COW,ST
PROPELL._CTTEMP& CONST.
5.0 6,0
Figure 7. Effect of Mixture Ratio
GAIN
MAX GAIN
1.0
0,9
0.8
0.7
0.6
500 600 700 800 900 I000
HYDROGENTEI,:P °R
TEMP. SENSITIVE COMBUSTION
Pc = CONST.
Figure 8. Effect of Hydrogen Temperature
Z87
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N71 -29586
"ADVANCED THRUST CHAMBER (NON-TUBULAR)"
D. L. FULTON
ROC KE TDY NE
TECHNICAL MANAGER
C. D. PENN
USAF. ROCKET PROPULSIO:_ L_%PORAqORY
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