The life analysis used to compare copper tubes and milled copper channels for rocket engine cooling passages is described. Copper tubes were chosen as a possible replacement for the existing milled copper channel configuration because (1) they offer increased surface area for additional enthalpy extraction; (2) they have ideal pressure vessel characteristics; and (3) the shape of the tube is believed to allow free expansion, thus accommodating the strain resulting from thermal expansion. The analysis was a two-dimensional elastic-plastic comparison, using a finite element method, to illustrate that, under the same thermal and pressure loading, the compliant shape of the tube increases the life of the chamber. The analysis indicates that for a hot-gas-side-wall temperature of 100 F the critical strain decreases from 1.25 percent in the channel to 0.94 percent in the tube. Since the life of rocket thrust chambers is most often limited by cyclic strain or strain range, this decrease corresponds to an expected tube life which is nearly twice the channel life

Jankovsky, R. S.

Kazaroff, J. M.

English

NASA Technical Reports Server

NASA Technical Memorandum 103613 
A Life Comparison of Tube 
and Channel Cooling Passages 
for Thrust Chambers 
R.S . Jankovsky and J.M. Kazaroff 
Lewis Research Center 
Cleveland, Ohio 
Prepared for the 
1990 JANNAF Propulsion Meeting 
Anaheim, California, October 3-5, 1990 
NI\5/\ 
11-/ 3- '1D 
e S?;JIj 
https://ntrs.nasa.gov/search.jsp?R=19910001746 2020-03-19T20:00:56+00:00Z
A LIFE COMPARISON OF TUBE AND CHANNEL COOLING PASSAGES FOR THIWST CHAMBERS 
R. S. J ankovsky and J. M. Kazaroff 
National Aeronautics and Space Administration 
Lewis Research Center 
Cleveland, Ohio 44135 
ABSTRAC!' 
This report describes the life analysis used to compare copper tubes and milled copper channels 
for rocket engine cooling passages . Copper tubes were chosen as a possible replacement for the 
existing milled copper channel configuration because (1) they offer increased surface area for addi-
tional enthalpy extraction; (2) they have ideal pressure vessel characteristics; and (3) the shape 
of the tube is believed to allow free expansion, thus accommodating the strain resulting from thermal 
expansion. The analysis was a two- dimensional elastic-plastic comparison, using a finite element 
method, to illustrate that, under the same thermal and pressure loading, the compliant shape of the 
tube increases the life of the chamber. The analysis indicates that for a hot-gas-side-wall tempera-
ture of 1000 of the critical strain decreases from 1.25 percent in the channel to 0.94 percent in the 
tube. Since the life of rocket thrust chambers is most often limited by cyclic strain or strain 
range, this decrease corresponds to an expected tube life which is nearly twice the channel life . 
I NTRODUC!' I ON 
The future cryogenic Space Transfer Vehicle Engine is required to have a long life and be reli-
able. This can be achieved with a thrust chamber that is compliant. Compliance can be inherent in 
tubular coolant passage construction, because this contour allows relatively free expansion. Free 
expansion accommodates and distributes the strains preventing local accumulation, and therefore 
extends the life of the chamber. This initial investigation is a 'simplified, relative comparison of 
the expected lives of mi lIed cooling channels and tubular cooling passages. The analysis compares 
only plug- nozzl e chaInbers that have been used at NASA Lewis Research Center for thrust dlaIDber study 
and development (Fig. ) , Ref . )). The plug- nozzle chamber has been tested and studied since the 
early 1970 ' s and the milled channel configuration has been very well characterized. This investiga-
t i on compares two pI ug- nozz I e chamber con f i gura t ions: (l) a mi I led OFHC copper channe I type 
(Fig. 2); and (2) an OFHC copper tube bW1dle type (Fig. 3) . The comparison is done at the same pres-
sure loading (i .e. , the same chamber pressure, and same coolant pressure) and the same thermal load-
ing so that the pure compliance, or geometric benefit can be determined. 
FIGURE 1. - PLUG-NOZZLE CHAMBER BE I NG TESTED AT NASA LEW I S 
RESEARCH CENTER. 
FIGURE 2. - CHANNEL PLUG-NOZZLE CHAMBER. 
FIGURE 3. - TUBULAR PLUG-NOZZLE CHAMBER. 
ANALYTICAL PROCEDURE 
The analytical procedure consisted of a thermal analysis, a strain analysis, and a life correla-
tion. This three part analysis was performed on a section through the throat of the thrust chamber. 
This area was selected for the analysis because the area of highest heat flux generally produces 
the highest thermal strains. The thermal analysis was performed using SINDA '85, a general purpose 
heat transfer /fl u id program that simulates transient or steady-state energy storage and flow in a 
system modeled as a " thermol circuit." The steady state temperature profiles were determined for 
both the channel and tube structures. The channel model used 185 elements, and the t ube model used 
180 elements (Figs. 4(a) and (b)). The coolant temperature in both models was set at -360 OF to 
account for the warming of the hydrogen before it gets to the throat. The heat flux for the channel 
model was set at 33 Btu/in.2/sec (Ref. 1). The maximum heat flux for the tube model was 33 Btul 
in .2/sec and decreased according to the profile reported in Ref. 2. This compensates for the thick-
ening boundary layer about the periphery of the tube. The coolant-side film coefficient was varied 
in both mode ls until a maximum hot-gas- side wall temperature of 1000 OF was obtained . (Ref. 1 shows 
that 1000 OF was the hot -gas-side wall temperature for past testing.) 
The strain analysis was performed using a two-dimensional elastic-plastic MARC analysis. MARC 
is a general purpose finite element analysis program. Both strain models contained the same number 
of elements as their thermal models. The models consisted of 8-node, three-dimensional, isoparame-
tric elements (MARC element 7). The strain analysis used the temperatures from the SINDA '85 
analysis along with a chamber pressure of 600 psi and a coolant pressure of 1200 psi. (These are 
standard operating conditions for plug-nozzle tests.) (Ref. 1) The constraints were the same for 
both configurations. The cool outer structure was fixed and the nodes along the planes of symmetry 
assigned zero displacement perpendicular to these planes, but permitted to move in the radial direc-
tion (Figs. 4(a) and (b)). The material properties such as coefficient of thermal expansion 
(Fig. 5(a), Ref. 3), yield strength, and modulus of elasticity were input as a function of tempera-
ture (Figs. 5(b) and (c), Ref. 4). The strain analysis was done using two simplifying assumptions 
and is considered valid when making only relative comparisons. First, no creep analysis was per-
formed. This absence of creep is considered a valid assumption for the analysis of the plug-nozzle 
chambers because of the short cycle duration (3.5 sec) (Ref. 1). Second, the first cycle was taken 
to be representative of all cycles to follow . The crit ical strains used for the life analysis were 
determined from taking into account the magnitude of the strains (from the finite element analysis) 
along with other effects such as "ratcheting " (from test experience) that are known to shorten life 
(Ref. 5). These critical strains at steady state conditions were used to determine life instead of 
the more complicated, tranSiently analyzed strain ranges used in other predictions. 
The life correlation was done using experimental low cycle fatigue data (Fig. 6, Ref. 6). This 
portion of the analysis was not meant to give an absolute value of life, rather it uses the low cyc le 
fatig ue data as a measure of the material's sensitivity to strain. Knowing the materials sensitivity 
to strain, and the relative decrease in strain, a life improvement prediction can be made (Refs . 7 
and 8). 
2 
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(a) 185 ELEMENT CHANNEL . (b) 180 ELEMENT TUBE. 
. 020 
.016 
. 012 
.008 
.00q 
o 
FIGURE q. - MODELS WITH CONSTRA INTS . 
400 1200 
( a ) COEFFICIENT OF THERMAL EX PANSION VERSUS TEMPER-
ATURE. 
FIGURE 5. - OFHC COPPER MA TER I AL PROPER TI ES. 
3 
60 
SO 
'" 
'" 40 
UJ 
'-" z: 
<: 
'" z: 
<: 
'" l-V> 
10 
OL---______ L-________ -L ________ ~ 
o 500 1000 
(b) STRENGTH VERSUS TEMPERATURE (F). 
::r- ------
12 L---------~----------~------____ ~ 
- 400 0 500 1000 
TEMPERATURE , of 
(c) MODULUS OF ELASTICITY VERSUS TEMPERATURE . 
FIGURE 5. - CONCLU DE D. 
1 0-1 L---------------~------------------' 
1~ 1~ 1~ 
CYCLES TO FRACTURE 
FIGURE 6. - TYP ICAL LOW-CYCLE FATIGUE OF OFHC COPPER 
ANNEAL ED CONDITION AT £ ~ 0.002 SE C- 1 IN ARGON AT 
1000 of . 
RESULTS 
This analysis of the two chamber coolant passage configurations was undertaken to compare their 
cyclic lives. This effort began with the thermal analysis using SINOA '85. The SINDA '85 analysis 
resulted in steady state temperature profiles that are shown in Figs. 7 and 8. The maximum hot-
gas-side wall t empera ture was set to 1000 of and at this condition the profiles were very similar. 
The hottest elements were located at the center of the channel and at the crown of the tube. 
With the thermal profiles, the material properties (as a function of temperature) and test pres-
sures the strain analysis was performed. This resulted in the strain profiles of the two configura:" 
tions as shown in Figs. 9 and 10. These strains are represented as Von Mises strains which are 
equivalent uniaxial strains. As can be seen from the strain contour (Fig. 9) for the channel struc-
ture, the location of the highest strain is at the center of the rib. This is not the typical fai l-
ure si te for this structure when chambers are actually tested, Fig. 11 shows all actual fai lure that 
has the rupture at the center of the cooling passage. This is not considered to be in error because 
in actual testing there are ratcheting affects (Ref. 9) taking place at the center of the charu1el 
which reduce the life dramatically (Ref. 5). The strain at the center of the channel is just 
sligh tly less than the maximum strain at the rib, but when combined with the racheting effect the 
cri tical strain area is at the center of the channel . 
The strain contour (Fig. 10) for the tube structure shows that the highest strain occurs at the 
crown of the tube. The fai lure is expec t ed to occur at the center line of the tube crown for the 
same reason as the channel above, high strain and racheting. 
Figure 12 is a summary of the life analysis performed. The plot indicates experimental data of 
life as a function of hot-gas- side wall temperature for milled OFHC copper channel chambers (Ref. 1). 
It also shows the calculated life versus hot - gas-side wall temperature for the same configuration. 
There is very good agreement between the experimental data alld the calculated data. The calculated 
data is slightly higher because of the simplifying assumptions made for the analysis and the fact 
that the number of cycles to fai lure was determined using low cycle fatigue specimen data. The low 
cycle fatigue specimen data is the average life at a particular strain range. The low cycle fatigue 
curve is based upon a series of data points which have a degree of scatter, consequently the line 
represents some average. Experimentally the plug-nozzle chambers are considered fai led when the 
first sign of a crack through the wal I occurs. These chambers have 72 challnels, hence 72 fatigue 
specimens. Therefore, when the first crack occurs through the wall the weakest specimen is taken as 
the data point not the average as ill the low cycle fatigue data. Figure 12 also shows the calcula-
ted life of the tube as a fWlction of hot - gas - side wall temperature, showing that the tube cooling 
passage has promise of living approximat e ly 100 percent longer thaJl the channel cooling passage. 
TEMPERATURE. 
OF 
I--------l 145 
1-------1 205 
L-_---;!266 
989 
I 
, 326 
386 
446 
507 
567 
627 
688 
748 
808 
868 
929 
FIGURE 7. - CHANNEL MODEL 
TEMPERATURE CONTOURS. 
------ -------
TEMPERATURE. 
OF 
, 
.---------11156 
I 
! 
~---___;I 215 1 ~V3 
I 1 332 
390 
448 
507 
565 
624 
682 
740 
799 
FIGURE 8. - TUBE MODEL 
TEMPERATURE CONTOURS . 
VON MI SES 
STRAIN. 
IN .lIN . 
I 
I ~-----i .000797 
I 
L-----, .001 70 
I 
I : I : .00261 i ~ V 1 00351 L/l ' 
I 
,
.00442 
.00563 
, .00675 
_ _ ....0..:.'--___ 
.0124 - - ~00900 
'--_:=:--.1 __ --.0113 
:---.0120 
.0124 =--L-'-::::::J-:-'0130 
-~ 0135 
FIGURE 9. - CHANNEL MODEL 
STRAIN CONTOURS. 
4 
VON MISES 
STRAIN. 
IN .lIN. 
t i .000754 
n
! .00135 
I .00195 
i /l.00255 
V~·OO31S J i 
I 
I 
.00495-r---....J 
.00375 
.00435 
.00615/ - .00555 
~ -.00674 
~. J.~:::4 .00854 .00914 
FIGURE 10 . - TUBE MO DEL 
STRAIN CONTOURS. 
J 
CHANNEL (EXPER I MENTAL) (REF. 1) 
{::,. CHANNEL (CALCULATED) 
0 TUBULAR (CALCULATED) 
... 1000 
0 
".j 
950 
'" ::::> ~ 
<: 
'" I...J 900 c.. 
>: 
... 
-' 850 
<i 
ox 
'" 800 
'" (J') LIFE , (J') 
<: 750 TO STRUCTURAL COMPLIANCE 
"i' 
0- OF TUBE CONSTRUCT! ON §? 
700 
200 400 600 800 1000 1200 1400 
CYC LES TO FAILURE, Nf 
FIGURE 11. - TYPICAL CHANNEL WALL FAILURE. FIGURE 12. - HOT-GAS-SIDE WALL TEMPERATURE VERSUS 
CYCLES TO FAILURE. 
SUMMARY Of RESULTS 
AIl increase in life of a tubular coo l ant passage configuration, when compared to an equivalent 
channel configuration, of -100 pe rcent wa s found for hot-gas- side wall temperatures ranging from 
800 to 1000 Of. A summary of cri tica l strains along wi th the corresponding stresses and number of 
cycles to failure are provided in Pig . 13 . A comparison of the number of cycles to failure as a 
fWlction of hot - gas - side wall tempe rature for the channel and tube cooling passages is also provided 
i ll Pi g. 13. 
HOT-GAS- STRESS, CR ITICAL CYCLES ro 
SIDE WA LL PSI STRAIN , FAILURE 
TEMPERATURE, IN./IN. 
OF 
MILLED 1000 8650 0.0125 350 
CHANNEL 900 8025 0.0108 500 
PASSAGE 
800 8001 0.0091 740 
TUBE 1000 6969 0.0094 700 
PASSAGE 900 7081 0.0083 900 
800 7184 0.0073 1200 
FIGURE 13. - SUMMARY TABLE OF HOT-GAS-SIDE WALL TEMPERATURE, 
STRESS, STRAIN, AND CYCLES TO FAILURE. 
RECOMMENDATIONS 
Three recommendations are made for tubular configuration life analysis. first, although this 
analysis is valid for relative comparisons of channel and tube structures, a three- dimensional visco-
plastic analysis of the tube structure is required to get a accurate understrulding of what is happen-
ing. Second , alloy tubes should be analyzed to determine if a lloy properties are applicable when 
compliance is the goal. Third, the results of both this analysis and subsequent rulalyses need to be 
validated wi th testing. 
REfERENCES 
1. Quentmeyer, R.j . : Experimental fatigue Life Investigation of Cylindrical Thrust Chambers. 
AlAA Paper 77-893, July 1977 (Also, NASA TM X-73665). 
2. Kacynski, K.J.: A Three-Dimensional Turbulent Heat Transfer Analysis for Advanced Tubular 
Rocket Thrust Chambers . NASA TM-l03293, 1990. 
5 
3. Schmidt, E.M.; Williams, L.R.; and Yow1g, j.D . : Design and Development Engineering Handbook of 
Thermal Expansion Properties of Aerospace Materials at Cryogenic and Elevated Temperatures. 
(R- 6981, Rockwell International; NASA Contract NAS8-19) NASA CR-102331, 1967 . 
4. Esposito, J.J.; and Zabora, R.F.: Thrust Chamber Life Prediction. Vo lume I: Mechanical and 
Physical Properties of High Performance Rocket Nozzle Materials . (D180-18673-1-Vol-1, Boeing 
Aerospace Co.; NASA Contract NAS3-17838), NASA CR-134806, 1975. 
5. Manson, S.S.; and Halford, C.R.: Treatment of Multiaxial Creep-Fatigue by Strainrange Parti -
tioning. 1976 ASME-MPC Symposium on Creep-Fatigue Interaction, R.M . Curran, ed., American 
Society of Mechanical Engineers , 1976, pp. 299-322. 
6. Conway, J .B.; Stentz, R.H.; and Berling, J .T.: High Temperature , Low Cycle Fatigue of Copper~ 
Base Alloys in Argon. Part 1: Preliminary Results for 12 Alloys at 1000 of (538°C). NASA 
CR- 121259, 1973. 
7. Shoji, ].M.: Advanced Hydrogen/Oxygen Thrust Chamber Design Analysis. (R-9258, Rockwell 
International; NASA Contract NAS3-1G774) , NASA CR-121213, 1973. 
8. Andrus, J .S.; and Bourdeau, R.C.: Thrust Chamber Material Technology Program. P&W Report 
FR-20672, Pratt and Whitney, 1989. 
9. Hannum, N.P.; Kasper, H.J. ; and Pavli, A.J . : Experimental and Theoretical Investigation of 
Fatigue Life in Reusable Rocket Thrust Chambers . AlAA Paper 76-685, July 1976 (Also, 
NASA TM-X- 73413). 
6 
---------
-- ---~ 
NJ\SI\ Report Documentation Page National Aeronautics and 
Space Administration 
1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. 
NASA TM-103613 
4. Title and Subtitle 5. Report Date 
A Life Comparison of Tube and Channel Cooling Passages for Thrust 
Chambers 
6. Performing Organization Code 
7. Author(s) 8. Performing Organization Report No. 
R.S. lankovsky and I .M. Kazaroff E-5724 
10. Work Unit No. 
591-41 -21 
9. Performing Organization Name and Address 
11 . Contract or Grant No. 
National Aeronautics and Space Administration 
Lewis Research Center 
Cleveland, Ohio 44135-3191 13. Type of Report and Period Covered 
12. Sponsoring Agency Name and Address Technical Memorandum 
National Aeronautics and Space Administration 14. Sponsoring Agency Code 
Washington, D.C. 20546-0001 
15. Supplementary Notes 
Prepared for the 1990 lANNAF Propulsion Meeting, Anaheim, California , October 3-5 , 1990. 
16. Abstract 
This report describes the life analysis used to compare copper tubes and milled copper channels for rocket engine 
cooling passages. Copper tubes were chosen as a possible replacement for the existing milled copper channel 
configuration because (1) they offer increased surface area for additional enthalpy extraction; (2) they have ideal 
pressure vessel characteristics; and (3) the shape of the tube is believed to allow free expansion, thus 
accommodating the strain resulting from thermal expansion. The analysis was a two-dimensional elastic-plastic 
comparison , using a finite element method , to illustrate that , under the same thermal and pressure loading the 
compliant shape of the tube increases the life of the chamber. The analysis indicates that for a hot-gas-side-wall 
temperature of 1000 OF the critical strain decreases from 1.25 percent in the channel to 0 .94 percent in the tube. 
Since the life of rocket thrust chambers is most often limited by cyclic strain or strain range, this decrease 
corresponds to an expected tube life which is nearly twice the channel life. 
17. Key Words (Suggested by Author(s)) 18. Distribution Statement 
Copper; Tube; Channel; Life; Rocket chamber Unclassified - Unlimited 
Subject Category 20 
19. Security Classif . (of this report) 20. Security Classif. (of this page) 21. No. of pages \22. Price' 
Unclassified Unclassified 8 A02 
NASA FORM 1626 OCT 86 
'For sale by the National Technical Information Service, Springfield, Virginia 22161 
----- -- - -------- ---- ----- - -~--


A life comparison of tube and channel cooling passages for thrust chambers

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