^C N ^qjf GB ^)C R)- ^`^ `^ er W N IACA Q U d z • i i..i ,, ww.+ C opy RM E57G29 RESEARCH MEMORANDUM CLASSIFICATION MIANGE Tol^YL1^" By authority 0-6 J/^^, C ^'lh^^ Changed byAL- - - -----Date =^Z= ALTITUDE STARTING TESTS OF A 1000-POUND-THRUST SOLID-PROPELLANT ROCKET By John L. Sloop, R. James Rollbuhler, and Eugene M. Krawczonek Lewis Flight Propulsion Laboratory Cleveland, Ohio is NISE 'T T ala^ °t to be ^T i^ ez' f Pet_ W 0= ' ,y/ced, ^_.,; tte&Ao en t I e Withou bitiot^ tT 1 11AGp- CLA IED ..NT This material contains information affecting Lionel Defense of the United Statea within the meaning of the espiona ge laws, Title 18, U.S . C., Secs. ,l93 794, the transmission or revelation of which In any manner to unauthorized person to prohibited w. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS W A S H I N G TO '^nys I b AU Y»1faG:^+F^^^Gt^
Transcript
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RESEARCH MEMORANDUMCLASSIFICATION MIANGE
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Changed byAL- - - -----Date =^Z=ALTITUDE STARTING TESTS OF A 1000-POUND-THRUST
SOLID-PROPELLANT ROCKET
By John L. Sloop, R. James Rollbuhler, and Eugene M. Krawczonek
Lewis Flight Propulsion LaboratoryCleveland, Ohio
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This material contains information affecting Lionel Defense of the United Statea within the meaningof the espionage laws, Title 18, U.S . C., Secs. ,l93 794, the transmission or revelation of which In anymanner to unauthorized person to prohibited w.
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FOR AERONAUTICSW A S H I N G TO '^nys
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NATIONAL ADVISORY COMMITTEE FORO^NA`ICS
RESEARCH MEMORANDUM
ALTITUDE STARTING TESTS OF A 1000-POUND-THRUST SOLID-PROPELLANT ROCKET
By John L. Sloop, R. James Rollbuhler, and Eugene M. Krawczonek
SUMMARY
Four solid-propellant rocket engines of nominal 1000-pound-thrustwere tested for starting characteristics at pressure altitudes rangingfrom 112,500 to 123,000 feet and at a temperature of -75 0 F. All enginesignited and operated successfully. Average chamber pressures ranged from1060 to 1190 pounds per square inch absolute with action times from 1.51to 1.64 seconds and ignition delays from 0.070 to approximately 0.088second. The chamber pressures and action times were near the specifications,but the ignition delay was almost twice the specified value of 0.040 second.
INTRODUCTION
For long-range missiles, the separation of the nose cone from thepropulsion part of the missile is desirable in order to improve the aero-dynamic characteristics of the nose cone upon re-entering the atmosphere.This separation can be accomplished by the use of small solid-propellantrockets fired so as to decelerate the propulsion unit. Such engines,called "retro rockets", should start and operate throughout an altituderange of sea level to 100,000 feet and over a temperature range of -650to 165 0 F (ref. 1).
At the request of the Western Development Division of the Air Researchand Development Command, U.S. Air Force, the NACA Lewis laboratory conducteda brief series of starting and operating experiments with proposed retro-rocket engines at pressure altitudes from 112,000 to 123,000 feet and ata temperature of -75 0 F. Rocket operational parameters were measured and,in addition, temperature measurements were made downstream of the nozzleto aid in predicting the flame temperatures of surfaces in the vicinityof the rocket jet.
The work reported herein is similar to the tests reported in reference2 except that a different rocket engine was used,ai" ,\^^,sv,/I+/
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APPARATUS
The apparatus consisted of a solid-propellant rocket, an altitudetank that provides altitude pressure only, and a cooling system to coolthe rocket to the desired temperature. The altitude tank and coolingsystem are similar to those reported in reference 2 and are shown infigures 1 and 2.
Rocket Engine
The solid-propellant rocket engines (shown in fig. 3) were suppliedby the Phillips Petroleum Company and were designated Douglas Model DM-18retro-rocket engines. The engine is 22 inches in over-all length with aninternal-external burning grain of ammonium nitrate supported by syntheticrubber pads and ignited by a squib energized by 24 volts direct current.
The specified performance temperatures at -65 0 , 600 , and 1650 F aregiven in reference 1. An extrapolation of these values to -75 0 F is asfollows:
The measurements made were: (1) two combustion-chamber pressures,(2) altitude-tank pressure, (3) exhaust-flame temperature, (4) firing-pulse time, and (5) coolant-bath temperature. The first four measurementswere recorded on two multichannel oscillographs. Pressures were measuredby strain-gage transducers. Exhaust-flame temperatures were measuredwith 18 Chromel-Alumel thermocouples, positioned as shown in figure 1.Two thermocouple rakes, each containing, 9 thermocouples, were located inapproximately the same position relative to the exhaust nozzle (fig. 4);the thermocouples of one rake were shielded by a simple shield, as shownby the insert of figure 1. Firing initiation is shown by the start ofa 60-cycle trace on the oscillograph record.
PROCEDURE
Prior to the test runs, the rockets were stored in a controlled tem-perature refrigerator at -75 0 F for a minimum of 12 hours. After refrig-eration they were removed, assembled in the capsule (fig. 1), and mounted
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on the altitude tank. The coolant pump was started, and coolant fluid wascirculated through the system. Until the coolant bath surrounding therocket engine reached -750 F, all of the coolant was directed through theheat exchanger. When the desired bath temperature was reached, the mixingvalve was controlled manually to keep the bath temperature at -75 0 F.The coolant fluid was circulated at the constant temperature for a timeperiod equal to, or greater than, the time that had elapsed from the re-moval of the rockets from the refrigerator until the coolant was circu-lated around the engine at -75 0 F. These time intervals and the otheroperating conditions are shown in table I.
The rocket performance terminology used in presenting the data isdefined in the appendix.
RESULTS AND DISCUSSION
The four rockets ignited and operated successfully. Oscillographrecords of chamber pressures are shown in figure 5. The runs are charac-terized by a rapid pressure rise at the start, followed by a slump andthen by a slow or stepwise build-up to the maximum pressure. This ismore clearly illustrated in figure 6, where all pressure data are plottedon the same scale.
The chamber-pressure records A and B did not agree in either run 2or run 4. In each case, the transducer connected to oscillograph Bfailed to return to its pre-run zero level; the transducers may have beenaffected by the rocket-chamber temperature. This difference in zero levelsamounted to 120 pounds per square inch for run 2 and 150 pounds per squareinch for run 4. The pressures were corrected by assuming that the truezero drifted continuously and linearly from the beginning to the end ofthe run. Consequently, a proportionate part of 120 or 150 pounds persquare inch was added to each indicated pressure throughout the run. Whenthis continuous correction to zero is made, the two records agree moreclosely with each other and with the other runs (e.g., fig. 6).
Data from the pressure records are tabulated in table II. The averagechamber pressures ranged from 1060 to 1190 pounds per square inch absolute,which is near the expected value of 1050 pounds per square inch absolutefor -75 0 F. Action times ranged from 1.51 to 1.64 seconds, or slightlylonger than the specified time of 1.42 seconds. These values are betterthan those required in the specifications. The ignition delay, however,ranged from 0.070 to approximately 0.088 second, which is almost twicethe value given in the specifications. The ignition-delay value was ap-proximate because the firing initiation failed to record for run two, andfor one of the records on run 4.
The rocket jet temperatures are shown on the oscillograph records(fig. 5). The temperatures for the four runs are presented in figures 7to 10 for both bare and shielded thermocouples. Figure 11 shows a
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comparison of temperatures at the end of each run for the bare and shieldedthermocouples. The peak temperature occurred at the number 3 position, orabout 6 inches downstream of the exit for the bare thermocouples, and atthe number 2 position, or about 4 inches downstream of the exit.for theshielded thermocouples. The temperatures measured with the shieldedthermocouples are higher than those measured with the bare thermocouples.This appears to be either an experimental error, such as misalinementbetween the centerline of the jet and the thermocouple rake, or possiblyjet asymmetry. It is also possible that the unshielded thermocouples werecooled, rather than heated, by radiation. The results refute the idea ofreference 2 that the peak temperatures may be the result of radiationfrom the shock or Mach disc. It is likely that, in the present tests,both types of thermocouples are within the shock zone. Reference 3 showsthe shock or Mach disc very clearly and shows how it increases in diameterand distance from the nozzle exit for altitudes up to 85,300 feet, thelimit of the experiment. The extent of the shock changes for the higheraltitudes simulated in the present investigation is not known. The peaktemperature obtained in reference 2 was inversely proportional to thepressure, whereas, for the experiments reported herein, little change inpeak temperature occurred for a 50-percent change in pressure. It isalso interesting to note that the peak temperatures obtained in theseexperiments and in reference 2 were not greatly different, although thepressure and thrust were greater in the present case.
Lewis Flight Propulsion LaboratoryNational Advisory Committee for Aeronautics
Cleveland, Ohio, July 29, 1957
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5
APPENDIX - TERMINOLOGY
The following definitions used in this report are taken from reference 1:
Ignition time: Time between zero time and 75 percent of maximum pressure
Burning time: Interval between 10 percent of maximum pressure and timewhen pressure begins to drop sharply near end
Action time; Interval between pressure rise of 10 percent of maximumpressure and pressure fall to 10 percent of maximumpressure
Pressure, averagechamber: Area under pressure-time curve between 10-percent
points, divided by action time
Pressure, maximum: Highest pressure developed by rocket engine under anyoperating condition
REFERENCES
1. Anon.: Drawing No. 7607429 - Model DM-18 - Retro Rocket Engine.Santa Monica Div., Douglas Aircraft Co., Inc., July 16, 1956.
2. Sloop, John L., and Krawczonek, Eugene M.: Altitude Starting Testsof a Small Solid-Propellant Rocket. NACA RM E57F21, 1957.
3. Balwanz, W. W., and Ward, E. W.: Interaction Between ElectromagneticWaves and Flames. Pt. I - The Effect of Altitude on Rocket-MotorFlame Geometry. Rep. 4232, Naval Res. Lab., Oct. 30, 1953.
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TABLE I. - OPERATING CONDITIONS
Run 1 2 3 4Rocket 112 104 101 105Igniter I41-34-57 I41-36-57 I41-35-57 I41-33-57Refrigeration time,
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16 CONFIDENTIAL NACA RM E57G29
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(b) Shielded thermocouples.
Figure 7. - Temperatures in vicinity of rocket jet,run 1.
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17
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(b) Shielded thermocouples.
Figure 8. - Temperatures in vicinity of rocket jet,run 2.
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Figure 9. - Temperatures in vicinity of rocket ,het,
run 3.
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Figure 10. - Temperatures in vicinity of rocketjet, run 4.
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Figure 11. - Temperatures in vicinity of rocket het.
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Figure 11. - Concluded. Temperatures in vicinity of rocket ,jet.
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ALTITUDE STARTING TESTS OF A 1000-POUND-THRUST SOLID-PROPELLANT ROCKET
/'Z^John L. Sloop
R. James Rollbuhler
ugene M. Krawczonek
Approved:
Walter T. OlsonChief, Propulsion Chemistry
Division
smr - 7/29/57
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NACA RM E57G29
ALTITUDE STARTING TESTS OF A 1000-POUND-THRUST SOLID-PROPELLANT ROCKET
By John L. Sloop, R. James Rollbuhler, and Eugene M. Krawczonek
ABSTRACT
Four nominal 1000-pound-thrust solid-propellant rocket engines weretested for starting characteristics at altitudes near 115,000 feet andat a temperature of -75 0 F. All ignited successfully and operated atthe expected pressure and for the specified time.
INDEX HEADINGS
Engines - Rocket 3.1.8
Combustion - Effect of Engine Operating Conditions andCombustion-Chamber Geometry 3.5.2
