Table XIII. Injection Molding Conditions Total cycle time Injection time Rear cylinder Front cylinder Mold temperature Injection pressure (to fill cavity) Back pressure
40 sec 10 sec 500-600” F 500-600” F 100-150” F 6000-10,000 psi 500 psi
D790-63. Heat deflection was measured according to ASTM D648-56. Acknowledgment
The author thanks the following for contributing to this work: G. C. Boyd, D. W. Broderick, L. R. Jacobson, C. R. Olson, E. P. Plueddemann, H. L. Vincent, and D. E. Weyer. literature Cited
The mixing and molding operations were carried out using a 1.5-oz Newberry injection molder with a reciprocating screw. The chopped glass and thermoplastic powder or pellets were dry blended and fed into the injection molding machine. I n cases where samples were to be evaluated by extrusion blending, pelletizing, and remolding, the compound was extruded through the injection molder and chopped into %-in. long pellets. In the majority of cases, the test specimens were molded from a dry blend. The injection molder settings are given in Table XIII. Tensile strengths were tested according to ASTM D63864T. Flexural strengths were tested according to ASTM
Engelhardt, J. T., Krantz, F. G., Philips, T. E., Preston, J. A., Wood, R. P., “The Influence of Reinforcements on Strengths and Performance of Fiberglass Reinforced Thermoplastics,” 22nd Annual Technical Conferences, Society of Plastics Industries, Reinforced Plastics Division, February 1967. Plueddemann, E. P., “Temperature Dependence o f Coupling of Thermoplastics to Silane Treated Glass,” 21st Annual Technical Conference, Society of Plastics Industries, Reinforced Plastics Division, February 1966. RECEIVED for review February 27, 1970 ACCEPTED August 21, 1970
Ablative Composites Containing in Situ Reaction-Formed Refractory Metal Carbides Fred 6. Speyer TRW Equipment Laboratories, Materials Technology, Cleveland, Ohio 441 17 A unique concept of in situ carbiding in ablative composites was prognosticated and, by combining a novel organometallic synthesis with a composite fabrication process, extreme environmental material properties were realized. Refractory metal carbides were formed within the fiber bundles of graphite cloth and throughout the fiber bonding matrix of laminate-molded, pyrolyzed, and carburized composites. Refractory metal alkoxide and glykoxide solutions were prepared and used to saturate graphite cloth. The saturated cloth, after heat treating to provide polymetalloxanes within the fiber bundles, was impregnated with phenolic resin and laminate-molded into composites. Carburized composites were also fabricated using refractory metal hydride dispersed phenolic resin matrices. Pyrolyzation, reimpregnation (densification), and carburization of the composites were performed. Resultant (graphite-refractory metal carbide-carbon) composites were machined into test specimens, plasma flame treated, and evaluated for high temperature, oxidation-erosion resistance. Metal carbiding of both the reinforcing graphite fibers, and the resin matrix by in situ reactivity within ablative composites provided improved high temperature oxidative-erosion resistance when measured b y plasma flame tests. Also, rocket nozzle specimens have been fabricated using in situ carbiding procedures and have been exposed to high pressure liquid propellant environments. Postfiring analyses have indicated good erosion resistance.
T h e ultimate objective of the work reported here was the development of an ablative composite which could be fabricated into rocket propulsion specimen capable of withstanding highly oxidative liquid propellant environments a t high temperatures and pressures. Composite reinforcement for aerospace use in such applications as rocket thrust chambers where silica and carbon were shown by a T R W Report (1968) to be inadequate in withstanding the combination of high temperatures, high gas pressure, and highly oxidizing environments can perhaps be best augmented through the use
of refractory metal carbides in that certain of them have very high melting points (Speyer, 1969133, and were shown by Laverty (1967) to have good resistance to oxidation. Laverty (1967) demonstrated that the threshold of oxidation for the carbide materials is considerably higher than for graphite materials. Therefore, the use of selected materials in an oxidizing environment is desirable, since even when most carbides become oxidized, a nonvolatile product is formed which can provide a protective barrier layer in the form of a solid oxide or viscous molten layer. A TRW Report (1968) shows monolithic refractory Ind. Eng. Chem. Prod. Res. Develop., Vol. 10, No. 1, 1971
99
metal carbides used in rocket propulsion systems were limited because of their extreme brittleness and sensitivity t o thermal shock. Thus, reinforcement of carbides with fibers is desirable. Also, there is a limitation in size in which carbides can be fabricated by arc melting, powder metal methods, or other fabrication techniques. The method of producing a pyrolyzed, carburized composite offers the achievement of a fiber reinforced carbide structure wherein the reinforcement is a refractory metal, carbide-coated fiber such as carbon or graphite. The coating might represent perhaps 10 to 20% of the fiber cross section and the carbon char matrix of the composite might be converted almost entirely to a carbide structure. The result would be a reinforced carbide having essentially the heat and corrosion resistance of monolithic carbides but not subject to thermal shock or size limitations. Discussion
I n Situ Carbiding of Ablative Composites. Present modes of carbiding (carburizing) ablative composites result primarily in surface coating. A means of in-depth carburizing or penetration into the fiber bundles of the cloth is needed to give optimum property requirements. A method was conceived and developed for the synthesis of refractory metal organic complexes and their incorporation into the fiber bundles of graphite cloth prior to resin impregnation and laminate molding. Microscopic examination of a carburized tantalum-containing composite revealed tantalum carbide interspersed within the graphite fiber bundles. A preliminary examination of a refractory metal halide was made to determine its possible impregnation into graphite cloth fiber bundles from alcohol solutions (Bradley and Carter, 1961). The resultant alkoxide formation gave metalloorganic solutions which could readily saturate graphite cloth fiber and be held in bond upon evaporation of the carrying solvent. Excellent fiber wetting occurred, perhaps because of the presence of HC1 in the solutions. This treatment greatly improved the handling quality of the otherwise fragile graphite cloth. Even better properties, such as solution stability and less friable solids, were obtained from metal halide-glycol combinations to give
metal glykoxide solutions. By using an excess of the hydroxy-containing components one can prepare solutions containing concentrations of alkoxides or glykoxides predetermined t o give a particular refractory metal to carbon ratio thus regulating the degree of carbide formation when submitted to carburization temperatures. I t is also possible to prepare organometallic solutions containing various combinations of refractory metal complexes. Metalloxane-treated graphite cloth can be impregnated with a thermosetting resin (phenol aldehyde) which has been precombined with a metalloxane to increase the metal content of the composite. This provides a bonding matrix which can itself be later carburized to a metal carbide containing linkage between metal carbide-coated graphite cloth fibers. Refractory Metalloorganics. The polyvalent refractory metals are generally characterized by the formation of stable durable bonds with electronegative elements, resistance to carbon bonding, and colorlessness. I t has been postulated by Blumanthal (1967) that partially hydrolyzed alkoxide polymers are formed by oxygen atom bridging. The strongly electronegative oxygen atom induces considerable polarity in the M-0 bond. The properties of the M-0-C system are also affected by the electronic behavior of the alkyl group. Chloromethoxides of hafnium, niobium, tantalum, and zirconium were prepared and used t o impregnate graphite cloth. The resultant chlorometalloxane-saturated fabric, when impregnated with phenolic prepolymer solutions containing free hydroxyl groups, reacts to give an organometallic phenolate and free HC1. The presence of the chlorine in the chlorometalloxane and the released HC1 are deleterious to the subsequent laminate molding process. The highly acidic outgassing is extremely corrosive to mold dies and equipment and also precatalyzes the phenolic resin bonding matrix resulting in poor laminates. Synthesis of the refractory metal glykoxide solutions provided alcohol soluble organometallics containing fewer chlorine atoms than the alkoxide counterparts. Also, the resultant metalloxanes were less friable, and solution stability was greater. An analysis of the tantalum and zirconium metalloxane derived from the glykoxide solutions resulted in the following postulated molecular configurations:
C1 I
I
ZrCl,+ 2C2H4(OH)2-O-Zr-O-C~H4-OH + 3HCI 1
I
1
HzC 0
\ / C H2
TaCls + 3C2H4(OH)2+
HOH
100
Ind. Eng. Chern. Prod. Res. Develop., Vol. 10, No. 1, 1971
+ 5HC1
Figure 1. Multipurpose composites flow chart
STE3 1
STEP '2
STEP
STEJ
YED
=3
=?
=5
8-EP =E
STEP
=7
STED
=a
S:Ea 59
Figure 2 . Process evaluation flow diagram R-120. Phenolic resin (Coast M f g . 8 Supply Co.) M. M e t a l (Tu, Zr, Hf, etc.) MX. M e t o l halide (TaCli, etc.)
UZF
:r 'IX
Composite Fabrication. Process development of ablative composites was carried out keeping in mind the end requirement of fabrication into rocket propulsion hardware capable of withstanding highly oxidative liquid propellant environments at high temperatures and pressures. Figure 1 represents the various combinations of cloth reinforcements, resins, and metalloorganics which can be multiprocessed into composites possessing a conglomerate of possible properties being sought. The initial intent of the program was to effect an improvement over unmodified graphite composites against high temperature and high oxidation environment erosion. Figure 2 outlines a series of steps whereby this might be accomplished. Steps 4, 5, and 6 represent a densification procedure not necessary for preliminary comparative evaluations. Step 2 encompasses the possibility of saturating the cloth with a metal halide, a metal powder, or a resin, any one of which, alone or in combination, would be carried in solution or dispersion by a solvent. The addition of very fine particle size metal and metal hydride powders to the cloth-saturating polymer solution and to the phenolic preimpregnated solution provides the basic ingredient for later heat conversion to the metal carbide. Besides dispersion of these fine powders into the resin solutions for application to the laminate plies, they
can be dusted onto the graphite cloth either before or after saturation or preimpregnation. Metal powder addition offers the optimum in weight-to-volume ratio thus minimizing porosity and density distribution problems in the pyrolyzed composite. Experimental
Synthesis of Refractory Metal-Organics. The chloromethoxides of hafnium, niobium, tantalum, and zirconium were prepared from the metal chlorides. The metal chloride powders were slowly added to methyl alcohol in a glass reaction vessel fitted with an air driven stirrer, thermometer, condenser, and a ground glass joint and stoppered addition funnel. Because of the heat of solution and the methanol volatility, an external cooling system was used to maintain the solution temperature below 25" C. The refractory metal chloroglykoxides were prepared in the same apparatus. Typical compositional recipes are shown in Table I. Compositions 4101-77A and i 9 A are very stable at room temperature, but 4101-78A, even under refrigeration, will polymerize to an insoluble gel within a few weeks. Fabrication of Polymetalloxane Composites. The incorporation of the polymetalloxanes into ablative composites was generally accomplished using Hitco G1550 Ind. Eng. Chern. Prod. Res. Develop., Vol. 10, No. 1, 1971
101
parison of materials. A nitrogen plasma with a secondary oxygen gas injectant was chosen to simulate the oxidizing environment of a liquid propellant. This environment produced by the plasma easily duplicated the gas temperatures seen in a liquid propellant system. The gaseous oxygen injected into the plasma stream created oxidizing species and simulated the shearing action achieved. The gas velocities in the combustion chamber (for which the composites were developed) are well below Mach 1.0. Pressure effects and the increased reactivity associated with it are not duplicated by the plasma. After reviewing various test specimen configurations cs. test data, one configuration was chosen by Speyer (Sept. 1968) as a standard test specimen for screening candidate ablative composites using a plasma flame. The standard specimens, 0.25 in. square x approximately 3 in. long. were cut from molded laminates. These laminates were made from virgin pyrolyzed, carburized, or graphitized composite plies. The specimens were held during testing so that the plasma flame was parallel to the long direction of the specimen and impinged upon the laminate end grain. The plasma flame environment consisted of 7 5 ft i / hr of nitrogen plasma with 75 ft:'.:hr of oxygen injected into the flame a t ambient flow pressure. The plasma gun nozzle was held 1 in. from the test specimen, and the test duration was 15 sec. The temperature of the specimen was measured by optical pyrometer and recorded as the maximum heat obtained during the 15-sec test. Both the depth and weight erosion rates were determined. Composite Evaluation. Metal carbide-containing composites which exhibited adequate structural integrity were submitted to the standardized oxidizing plasma test previously described. Table I1 presents test data on a variety of composites. Specimens 4101-7E and -lOB were fabricated using the same batch preparation of chlorotantalum methoxide solution for impregnating the graphite cloth. The phenolic resin binder matrix contained no metallic content. The depth erosion of the -10B was considerably less than -7E even though the TaC content was smaller. This could be owing to a density differential (not known for -10B) or to the temperature developed by the plasma flame. Since the temperature was over 500°F greater for the -7E, the depth erosion would be expected to be greater. The weight erosion of the -7E would also be expected to be greater except that the -10B was reimpregnated
AND G A S SUPPLY
HEATER CABLES
Figure 38. Schematic diagram of impingement test apparatus
The physical integrity of the composite was greatly improved by a phenolic resin vacuum reimpregnation and repyrolysis. This also increases the amount of carbonaceous char required in large amounts to satisfy both the requirements for forming the metal carbides and for removing oxygen (from metal oxides) as carbon oxides. Carburization (refractory metal carbide formation) was carried out in a furnace at 3000" to 3500°F under 1 to 10 microns vacuum. Materials Evaluation Parameters. I n selecting the test conditions and configuration under which candidate composites might be evaluated, we kept in mind that actual hardware would be fabricated into ablative rocket thrust chambers. The liquid propellant, a standard in the aerospace industry for low thrust rocket engines, would be nitrogen tetroxide oxidizer with propellant Aerozine 50, an equal mixture of hydrazine and unsymmetrical dimethyl hydrazine. The target material requirements for the ablative composites were resistance to highly oxidizing high temperature gases a t about 5200°F and to combustion pressures in excess of several hundred psia for several hundred seconds. The plasma jet, Figures 3A and B, was selected as the most useful screening method for the preliminary com-
Table II. Plasma Flame Tests on Composite Specimens
Specimen no. 4101-
7E 10B 12C 55
55B 72A 18A CARB-IT E X 700
Comp
Oh
b y wt'
SP 9'. g/cc
G-1550
C
R-120
M e t a l carbide
Metol
40 68 46 21 22 29 66
33
...
27 19 27
Ta Ta Ta 66 Zr Zr
1.51
...
1.69 2.22 1.83 1.73 1.22
...
1.41
d
27
...
13
... 13
...
...
22
20 12
...
... 78 51
...
...
...
Plasma flone tests'
In situ reactivity', n e t o l distribution
In Isec
G/sec
Kone None Ta ZrH. ZrH, ZrHI
...
4080 3500 4200 4120 4180 4400 3800
0.0143 0.0060 0.0109 0.0073 0.0024 0.0060 0.0084
0.0282 0.0296 0.0342 0.094,3 0.01 84 0.1176 0.0504
...
3400
0.0090
0.0244
In binder
TaM TaM None Kone None ZrMIZrH,
...
...
Erosion
Max temp ' F
In fiber
"-1550 = Hitco graphite cloth, R-120 = Coast Mfg. & Supply phenolic resin. ' T a M = tantalum nietalloxane. Ta = tantaluni powder. 'Test specimen L, x !A x 3 in. bars. Plasma flame 75 f t ' / h r N, + 75 f t ' / h r 0, impinging on laminate plies 1 in. from nozzle for 15 sec. Rate of energy absorption using a !< x ,:I x 3 in. water cooled calorimeter was 1.673 Btu's:sec, Speyer (1969a. p 14). '6857 includes both graphite cloth and pyrolytic carbon. ' CARB-I-TEX = Carborundum's graphite fiber 'graphite binder laminate.
Ind. Eng. Chem.
Prod.
Res. Develop., Vol. 10, No. 1, 1971
103
Figure 4. Cross-section photomicrographs of composite 41017 graphite fibers interspersed with tantalum chloromethoxide matrix between fiber bundles
was 120°F higher. Figure 6 shows a photomicrographic cross section of Composite 4101-12C1 where Ta powder was added between the preimpregnated plies of graphite cloth. The composite was laminate molded, pyrolyzed, and carburized. The bright areas within the solid gray areas are TaC. Actually there were several void areas not shown in these photos so even though there was about 34% TaC in the composite, the specific gravity was only 1.40 g per cc. Composite 4101-55 was prepared by dispersing ZrH, into R-120 phenolic resin solution, preimpregnating into graphite cloth, B-staging in an oven a t 200" F, and then laminate molding to give a 0.027-in. per ply composite. A good flame erosion comparison can he made between specimen -55 (noncarhurized) and -55B (carburized). The latter, as expected, even at considerably lower density gave 33% less depth erosion and 195% less weight erosion. Composite 4101-72A contained ZrC both within the fiber bundles and throughout the resin hinder. At a 1000" F higher temperature than the nonmetal carbide containing CARB-I-TEX 700 there was 60% less depth erosion. This composite was also plasma flame tested for 30 sec, and at 4300" F gave onlv 0.0035 in./$ Y
Specimen Fabrication n
.- -.-.
-.w.,-.
L..II.IIY,=
micrograph of composite 4101..,...Ubd and carburized to give graphqposite. Bright areas are ToC ch was not carbonized-thus during pyrolysis weight loss occurred. ,pecimen with CARB-I-TEX 700, :d similar test temperatures, we weight of TaC accounted for its I erosion. When we compare the phenolic composites prepared by he graphite / TaC /phenolic composidified graphite phenolic composite :pth erosion rate and 170% greater the tantalum-organic complex does idle interstices. The gray areas are ; as the hinder between graphite vertically about midway through ght areas are graphite fibers which .meter in the photograph. Between IS can he seen light gray areas lark areas which are voids left by of the organometallic during tomicrograph a t 500x magnification n e -10B after reimpregnation with oast Manufacturing &Supply Co.). indicative of TaC. Because of the ation temperature there was proh,sion of the tantalum to its carbide. contains the same percentage of 'site, hut it is dispersed throughout ?r than within the fiber bundles. gher density and a greater graphiteshould contribute to better erosion I even though the test temperature 1. Res. Develop., Vol. 10, No. 1, 1971
jl
nasea1 upon m e results from iaooracory evamacions, a few compositions were selected for fabrication int,o rocket thrust chamber test specimens with subsequent propellant test firings and examination of the results. The basic design of the thrust chamber test units was selected to adapt to existing injector and test facilities a t Aeronutronics Division of Philco-Ford Corp. where the firing tests were carried out. The basic test engine configuration consisted of an injector, chamber module, chamber test component, and a water-cooled throat module shown schematically in Figure 7. The chamber test component was designed to include a 4-in. long material specimen. The chamber test component design is shown in Figure 8. The design incorporated a molded silica phenolic insulating sleeve between the chamber test specimen and the cylindrically shaped steel shell. The relatively simple design was selected to permit the thrust chamber assembly to he easily installed.into the reusable steel shell. The heavy wall of the shell permitted thermocouple packing glands to he threaded into the wall at desired locations. The processing data for the,five in situ carhided chamber units is given in Table 111. Fabricated chamber units
Figu
12c tanh
I l E T ~ LH C l t S l N G
Figure 7. Ablative test chamber
Y
U
U
U
Figure 8. Ablative chamber assembly print
.
2 . 'IX 26CC YO.i!\G C31WOUMD 2 . CGRBIDEC C0YPCS:TE
Table Ill. Rocket Nozzle Chambers Process Analysis Chamber unit no.
Saturated cloth'
Comp"
Zr(ZrC) (2-1550 Organic Ta(TaC) G-l,550 Organic Ta(TaC) Zr(ZrC) G-1550 Organic Ta(TaC) G-1550 Organic Ta(TaC) Zr(ZrCi G-1550 Organic
1-A 2 -A 4-A
2-D1 4-D1
35 32 33 37
37 26 18 8 45 29 36 37 27 18 8 45 29
Laminated molded billet'
42 24 34 39 28 33
21 8 25 46 34 19 47
19 9 23 49
Pyrolyzed, repregged
49 25 26 43 31 26 27 9 33 31 44 23 33 26 12 29 33
Pyrolyzed, carburized, repregged
Pyrolyzed, pack coated, graphitized
(74)
(72) 23
21 3
'3
(45) 30 25
(51) 33
16
Repregged, cured
SP 9'
(62) 17 21 (45) 29 26 ( 3 01
1.78
gicc
1.77 1.69
-(lo) 34 26 (54) r25 21 (32)
(4;ii
~ ( 1 G i
(271
35 17
29
--
1.79 1.72
:11
G-1550 is Hitco graphite cloth; organic includes resin, carbon. oxygen. and hydrogen. All numbers are percentage by weight. '' G-1550 saturated with metal-organic and heat treated. ' Saturated G-1550 impregnated with phenolic resin plus metal hydride or metal-organic. '! High temperature vacuum furnace tube cracked damaging billet. Final analysis after repair is an estimate. AIolded from chopped metal-organic saturated cloth with phenolic 'metal binder. '
~~
~
~~~~
Ind. Eng. Chem. Prod. Res. Develop., Vol. 10, No. 1, 1971
105
were bonded into prefabricated silica-phenolic insulation sleeves which were in turn bonded into an outer steel housing. This bonded assembly fit into the firing test unit per predetermined plan. Figure 9 shows a process flow sheet which designates the sequence of fabrication operations from raw materials t o assembled nozzle chamber. The general scheme of fabrication is shown in Table IV. Figure 10 shows a finished unit. Specimen Firing Tests
Figure 11 is a photograph taken during one of the test firings. The rocket motor plume is engulfed in a nozzle cooling water effluent which is dyed yellow by a corrosion inhibitor additive. The metallic aftsection is the water-cooled nozzle region with cooling water flex hoses attached. Ahead of this section is the 4-in:long specimen holder. Figure 12 shows a diagram of the rocket motor in a schematic o f the total system. The tanks, lines, and valves are shown for the N20d-Aerozeneand cooling water flow systems. Also shown is the pressure and temperature instrumentation. In every test the nominal conditions were: a chamber pressure of 1000 psia, N20d-Aerozeneratio 1.6, and theoretical temperature of 532PF. Table V is a summary of the test program. One of the specimens, A-1, was tested repeatedly. The other three specimens 4A, 2D1, and 4D, were evaluated in only one test firing each. Their high erosion rates precluded further testing. The criterion for stopping any test was a 600" F reading on the insulation. Temperatures were also taken a t three different intermediate depths. Before the final test of 1-A was run, cracked and loosened parts of the insulation were removed. The final test
Figure 9. Process flow sheet In situ carbiding of oblative pyrolyzed composites
106
Ind. Eng. Chem. Prod. Res. Develop., Vol. IO, No. 1, 1971
of I-A was unique. It consisted of five 10-sec pulses alternating with 10-sec shutdown periods. The hot restart pulses were uniform with the exception of the second and fourth pulses when brief pressure surges occurred, possibly owing to chunks of lining being ejected. For each liner specimen the chamber total wall thickness was measured hefore and after the test at four points circumferentially at each of six axial distances (a total of 24 measurements). From these data the erosion profiles were plotted, and maximum and average erosion rates were calculated. Photographs were taken of the tested specimens. Chamber LA, the zirconium-carbide-graphite material, was tested for a total of 221 sec in three separate test starts before taking erosion measurements. Post-test inspection revealed that the ablative chamber inner surface appeared evenly eroded t o about 2.525 in. diameter and was covered with a loose white ash. Prior to retesting of the chamber assembly this white ash was removed with low pressure nitrogen from an air hose, and the parent material was visually analyzed. The 1-A(3) retest was scheduled for 300 sec duration or for a shutdown when the monitored thermocouple reached a value above 2400°F without indicating that a temperature peak-out would occur. The test was terminated a t 109 sec a t which time the thermocouple readings were: 2580°F for thermocouple A; 2545°F for thermocouple B; and 2460°F for thermocouple C. Thermocouple D recorded 13P F. Post-test visual analysis of the chamber's inner surface revealed a very even diametral erosion that tapered inward from the upstream to the downstream end. The upstream end measured 2.550 in. in diameter at the edge with a 0.15-in. depression at the 12 o'clock position. Within
Table IV. Stepwise Processing of ZrC/C Chamber Analysis
Fabrication
Prepare Zr glykoxides (ZrG) Cut graphite cloth (G-1550) Saturate G-1550 with ZrG Heat treat G-1550iZrG Cut plys (washers) Prepreg (R-120 + ZrH,) Heat treat, trim Laminate mold Pyrolyze
Prepreg (R-120 + Zr&) and cure
51% Zr/49% org. 189 Plys/5" x 5"/800 g G-1550 + ZrG +Solvent 2500 g (51% Zr/49% arg)
1213 g (392" G15501425-Zi
499 g (250-R120/249-Zr) 1712 g 1480 gi4.07" h 1316 g (51% Zr/30%
G1550/19% C + 0) 1578 g (49% Zr/25% G1550/26% ore)
Repyrolyze .-...,
Carburize Repreg (R-120)bmachine ID-pyro Pack coat ID and carburize Renree . (R-1201. cure. I
nvro1vze G&hiGze
Repreg (R-120)and Cure Finish machine
-
~~
,
.
1163 g (75% ZrC/Z&k G1Ei50) 1134 g (70% ZrC/23% G1550/7% C) 1197 g 1335 E
1228 g (74% ZrC/ZI% G1550/5% C) 1485 g (62% ZrC/l7%
-
G1550/21% org)
assemble and bond
the
Figure 11. Chamber insulation test 0.2 in. from this edge, the diameter measured 2.700 in. on the vertical axis and 2.708 in. on the horizontal axis. At the downstream end, the diameter measured 2.640 in. on the vertical axis and 2.610 in. on the horizontal axis. There was 3 major depression in the center of the liner measuring 2.885 in. in diameter. There was a major undercutting of the liner section for 3 width of 0.3 in. at the downstream edge. The predominant undercuts were 0.2 in. deep a t the 12 o'clock position and 0.12 in. deep at the 4 o'clock position. Although clearly evident, the material chunking was not so pronounced as it had been after the second test and the overall linear surface appeared smoother. Figures 13A, B, and C, show photographs of chamber 1-A after 221 sec of accumulated firings. Figures 13A and B are taken from the downstream side; 13C is taken from the upstream side. ographs after test 5 (584
accumulated seconds of firing) for 2 and 3, and (B) radial positions 1 ai erosion rate is based on the largest ~ I O ~ L U I~Ii i w b u i t .a~ ~ one specific axial and radial position. A deeper erosic could have occurred in some area outside of these point Figure 15 shows the combined erosion profiles for chamba 1A a t the five sequential periods of continuous firing test Table V shows 3 summary of the chamber thermocouple data, the overall average erosion rate for the sequential firing tests, and an average erosion rate a t the midpoint through the chamber. After 584 sec of firing, average overall erosion of this chamber was only 0.33 in. or only about one-third of the wall thickness; this is an erosion rate of only 0.57 milsjsec. A t the 2.0-in. axial midpoint there was 0.40 in. erosion or about 0.68 milsjsec erosion rate. The final 50 sec of pulsed firing caused the erosion
YtNT
.LiNiO
Figure 12. Test cell B ablative liner test
LMHBIR
In,.
.,,Y.-
,.=.... ,
._".
",.
f
vc..
._,.
., ... .
._
-
X
c:
i
i
3
w N
N
SJ 5:
t-
d L9
N
i
ct
% N
2m w 2 c
a
$ K >
108
Ind. Eng. Chem. Prod. Res. Develop., Vol. 10, No. 1, 1971
i
C
m
C
C
-
5
B
A
C
Figure 13. ZrC/C chamber 1A after 221 rec of firing (A), (6) downstream side, (C) upstream side
B
A
Figure 14. Chamber liner 1A post-test, 5th test A. View of positions 2 and 3 8. View of positionr 1 and 4
of the remaining ablative chamber. During the second 10 sec of firing, pulse pressure increases were experienced. This was probably owing to ejection of small amounts of ablative chamber material. At the start of the third 10 sec of tiring, the thermocouple seated in the insulation sleeve burned out, indicating that the temperature in the sleeve was about 2500" F. Another pressure change took place during the fourth 10 sec of pulsed firing and probably represents a loss of chamber wall material. Figure 16 shows the post-test photograph after the pulsed firings. The ablative chamber has been washed out and silica
slag from the silica-phenolic insulation sleeve has started to form. Conclusions
Six ablative chamber units were fabricated and test tired. One of these units containing zirconium carbide gave very encouraging results. I t survived tive firings totaling 584 sec a t average chamber pressures over 1000 psia, resulting in an average erosion of 0.33 in. or just one-third of the ablative chamber wall thickness. This amounts to an average erosion rate of only 0.57 mil/ Ind. Eng. Chem. Prod. Res. Develop.,Vol. 10, No. 1, 1971
109
Figure 16. Chamber liner 1A. post-test
sec. An additional 50 sec of pulsed firing (5 pulses of 10 sec on and off in series) did cause erosion ejection of the remaining ablative chamber wall, hut the propellant injector was not optimized for pulsed firings, and this may he a t least the partial reason for the more rapid deterioration during pulsed firings. The temperature-pressure conditions during test firings were in excess of the silica-phenolic chambers survivability owing to molten flow of the silica into the nozzle throat. Although the tantalum carbided chamber had a considerably greater erosion rate (3.2 mils/sec/l14 sec) than the zirconium carbided chamber (0.3 mils/sec/ll2 sec), it should he pointed out that the volume percent of tantalum carbide (3.11) ratio to zirconium carbide (9.45) in the two respective chambers was three times greater for the zirconium carbide containing chamber. Even so the volume percent of metal carbide is relatively small and leaves room for significant increases provided that the physicomechanical integrity of the composite is maintained. Other possible points of vantage in comparing zirconiumto-tantalum carhided composites are lower molecular weight, 103 as against 193, provides a 47% weight advantage, and although TaC melts at ahout 705P F and ZrC a t ahout 6190” F, the TazOsmelts a t 3450” F compared to ZrO, a t 501PF. Since the metal carbides are oxidized during firing, the ZrO? would provide a higher melting harrier layer, and not the least in importance is the fact that zirconium powders and compounds are considerably (ding tantalum products. n. The metal glykoxides ides and upon carburizainder while eliminating iely powdered zirconium ed to air quickly oxidizes.
The present fabrication process is quite lengthy and is accompanied by the release of corrosive and obnoxious fumes. I t has, however, been demonstrated that an in situ reacted zirconium carbide-based composite rocket-thrust chamber can withstand nitrogen tetroxide-Aerozine propellant firings a t 1000 psi for GOO or more sec. Optimization of both material and processing could add significantly to the overall capability of the fabricated chamber. Literature Cited
Blumenthal, W. B., “Zirconium in the Cross-Linking of Polymers,” Rubber World, January 1967. Bradley, D. C., Carter, D. G . , Can. J . Chem., 39, 143443 (1961); 40,15-21 (1962). Laverty, D. P., “Carbides for Rocket Nozzle Inserts,” TRW Inc., Mater. Tech., Intern. Rept. March 28, 1967. Speyer, F. B., “Ablative Composites for High Pressure Liquid Propellant Environments,” AFML-TR-68-243 Part 1, September 1968. Speyer, F. B., ibid., Part 11, Fabrication and Tests, September 1969a. Speyer, F. B., “Structural Pyrolyzed Composites,” T R W Rept. TM-4512, December 196915. TRW Report, “Development of High Performance Materials for High Chamber Pressure Rocket Motor Application,” CR-M389, November 8,1968. RECEIVED for review February 27, 1970 ACCEPTED June 29, 1970 Presented at The Society of the Plastics Industry Meeting, Washington, D. C., February 1970. This work was sponsored by the U.S. Air Force Materials Laboratory, Contract F-33615-67-C-1620.Their direction, encouragement, and sponsorship are gratefully acknowledE,ed. The W ~ P I A F BMonitor was P. J. Pirrung. Ind. Eng. Chem. Prod. Res. Develop., Vol. 10, No. 1, 1971
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