A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds. The maximum thrust was 7200 pounds and occurred at 4.8 seconds after launching. No thrust irregularities attributable to effects of the flight longitudinal acceleration were observed. Certain small thrust irregularities occurred in the flight test which appear to correspond to irregularities observed in static tests conducted elsewhere. A hypothesis regarding the origin of these small irregularities is presented

Bond, Aleck C.

Thibodaux, Joseph G., Jr.

NASA Technical Reports Server

https://ntrs.nasa.gov/search.jsp?R=20090026460 2019-08-30T07:22:42+00:00Z
NACA RM SLgL13a 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
for the 
OrdnaJnce Department, Department of the Army 
FLIGHT TEST OF THE AEROJET ~XS-6000 T-27 
JATO ROCKET MOTOR 
By Aleck C. Bond and Joseph G. Thibodaux, Jr. 
SUMMARY 
A flight test of the Aero jet Engineering Corporation's 75-6000 
T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft 
Research Station at Wallops Island, Va, to determine the flight- 
performance characteristics of the motor. The flight test imposed an 
absolute longitudinal acceleration of 9.8g upon the rocket motor 
at 2.8 seconds after launching. The total impulse developed by the 
motor was 43,400 pound-seconds, and the thrusting time was 7.58 secoads. 
The maximum thrust was 7200 pounds and occurred at 4.8 seconds after 
launching. No thrust irregularities attributable to effects of the 
flight longitudinal acceleration were observed. Certain small t m s t  
irregularities occurred in the flight test which appear to correspond 
to irregularities observed in static tests conducted elsewhere. A 
hy-pothesis regarding the origin of these small irregularities is presented. 
INTRODUCTION 
The Aerojet m-6000 T-27 Jato rocket motor was developed by the 
Aerojet Engineering Corp. under the supervision of the Air Materiel 
Command and the Office of Chief of Ordnance as a solid-propellant, 
medium-duration, low-acceleration booster for NACA pilotless-aircraft 
research models. As part of the evaluation of this rocket motor it 
was desirable to determine its performance in flight as well as in 
static thrust-stand tests. Accordingly, a flight test of this motor 
has been conducted at the Langley Pilotless Aircraft Research Station 
at Wallops Island, Va. The results of this test are given in the 
present paper in correlation with results of static tests by the 
manufacturer and the Aberdeen Proving Ground. 
The static tests of the motor performed by the manufacturer azld 
the Aberdeen Proving Ground indicated that the motor functioned 
satisfacCorily during static firing. Such static tests yield internal 
ballistic data and time histories of pressure and thrust; however, 
these tests are inadequate for evaluation of the flight performance 
of the motor, as they do not indicate the effects of acceleration 
in free flight on the various components of the motor. 
The components most likely to be affected by acceleration in a 
solid-propellant rocket motor are the powder trap and the propellant. 
Failure of either of these components may lead ultimately to case 
failure due to high internal pressures resulting from blocking of the 
nozzle or increased propellant burning area. 
SYMBOLS 
a longitudinal acceleration with respect to the ground, 
feet per second per second 
% bag coefficient ( s ) 
D drag, pounds 
g acceleration due to gravity, 32.17 feet per second per second 
t time from instant of ignition, seconds 
T thrust, pounds 
W weight, pounds 
7 flight-path angle, degrees 
S cross-sectional area, square feet 
9 dynamic pressure, pounds per square foot 
P air denaity, slugs per cubic foot 
B velocity, feet per second 
DESCRIPTION OF THE ROCKET MOTCJR 
The Aerojet 71I;5-6000 T-27 Jato rocket motor as supplied by the 
manufacturer is shown in figure l(a) . A complete breakdown of all the 
component parts af the motor is shown in figure l(b) . The motor 
utilizes two single-perforated cylindrical propellant grains. The 
propellant is composed of an aslmonium perchlorate oxidizer bound in 
a nratrix of thermosetting resin fuel. The ends of the grain are 
restricted b~r a thermosetting resin backed with tar paper. The two 
grains are supported in the case by means of three trap plate6 held 
together by a rod passing through the central perforation of the grain. 
Inertia, viscous, and pressure forces are transmitted to the rear trap 
assembly, which is supported on steel legs welded to the nozzle entrance 
cone, and through the center trap assembly and tie rod to the boss 
in the front of the chamber. The loaded weight of the motor was 
527.75 pounds. The empty weight of the motor was 317.25 pounds. A 
more detailed description of the motor can be found in reference 1. 
DESCRIPTION OF THE MODEL 
The external configuration of the test model equipped with four 
stabilizing fins and nose is shown in figure 2. The fins were flat 
plates having wedge leading edges of 280 total angle measured in the 
direction perpendicular to the leading edge and square trailing edges. 
The nose was a cone 18 inches long and 1 3 3  inches in diameter. Within 
the nose was mounted the telemeter together with sufficient lead 
ballast to bring the center of gravity to the positions shown in 
figure 2 in order to achieve sufficient aerodynamic stability. The 
gross weight of the model was 708 pounds. The propellant plus igniter 
weight was 210.5 pounds, so that the empty weight of the model was 
497.5 pounds. 
TEST APPARATUS AND INSTR-TION 
The Aero jet 7~-6000 T-27 Jato rocket -motor flight model was launched 
from a near-zero-length, crutch-type launcher at an angle of 42.50 to 
the horizontal. The launcher consisted of a crutch-type wooden A-frame 
mounted on a T-shaped base. When mounted, the model was supported in 
front by the A-frame crutch and in the rear by means of a support under 
the nozzle. A photograph of the model on the launcher is shown in 
figure 3. 
The model was instrumented with a standard NACA two-channel 
telemeter transmitting measurements of longitudinal acceleration from 
two accelerometers. One accelerometer covered the range from 20g 
acceleration to lg deceleration, and the other accelerometer covered 
the range from lg acceleration to 15g deceleration. 
In addition to telemetered data, a time history of the flight 
velocity was obtained using a 100-watt continuous-wave Doppler radar 
unit. T b e  histories of range, azimuth, and elevation were obtained 
by means of a modified SCR-584 radar theodolite. Prevailing atmospheric 
conditions were obtained from a standard radiosonde released at the 
time of launching. A detailed description of the functioning of the 
CW Doppler radar and radio telemeter may be found in reference 2. 
The launching of the rocket was photographed with two fixed line- 
of-sight Army K-24 type rapid recycling aerial cameras and a high-speed 
35-millimeter Mitchell motion-picture camera 'in order to observe the 
operation of the launcher. Both the launching and the flight of the 
rocket were photographed with a tracking 16 -millbet er Cine -Kodak 
motion-picture camera in order to provide qualitative information on 
the stability of the flight and possible correlation of the appearance 
of the rocket jet with any anomlies that might occur in the rocket 
performance. 
RESULTS AND DISCUSSION 
Velocity 
The time history of the velocity obtained from the CW Doppler radar 
is presented in figure 4. The initial part of the curve was not 
obtained because of imbllity to track with radar equipment at close 
range. Zero time was established as the t h e  when ignition began. The 
model reached a maximum velocity of 1525 feet per second, corresponding 
to a Mach number of 1.37. 
Acceleration 
The time history of the acceleration as obtained from telemetered 
data is presented in figure 5. The accelerations shown are absolute 
accelerations and include the component of gravity resulting from model 
attitude. The record indicates a short delay in the initial portion 
of the curve which corresponds to the ignition delay of the motor and 
then a rapid change in acceleration is observed which corresponds to 
motor thrust build-up. At 1.1 and 4.4 seconds sudden increases in 
acceleration take place, indicating an erratic increase in thrust. 
CONFIDENTIAL 
A naaximum acceleration of 9.8g is reached 2.8 seconds after launching. 
Burnout occurs at 7.58 seconds. A maximum deceleration of 7.36 is 
reached at 7.7 seconds. A complete time history of the acceleration 
was obtained for the total flight oovering 17.3 seconds. 
Drag Coefficient 
Assuming that the angle of attack of the model is zero, the drag 
of the model during coasting flight may be obtained from the equation 
of motion 
= -w(i + sin 7) 
a 
where - + sin y is the absolute.longitudina1 acceleration as indicated Q 
directly by the accelerometer in g units. The assumption was made that 
the burnout weight of the model was equal to the initial weight of the 
model minus the propellant and igniter weight. 
The time history of the altitude as obtained from the modified 
SCR-584 radar is shown in figure 6. The variation of air density and 
speed of sound with altitude as obtained from the radiosonde data is 
shown on figure 7. These data were used for conversion of drag into 
drag coefficient based on the maximum model cross-sectional area 
of 0.948 square foot. Drag coefficient plotted as a function of Mach 
number is shown in figure 8. Data were inadequate for obtaining drag 
coefficients below a Mach number of 0.67; hence, the curve was 
extrapolated for lower Mach numbers . 
Thrus t 
The thrust may be determined from the drag and the absolute 
longitudinal acceleration by the following equation: 
where W is the instantaneous weight of the model- in pounds. The 
propellant was assumed to be expended at a constant rate for a first 
approximation of thrust. A second approximation was computed, assuming 
a change of propellant weight proportional to thrust. It was assumed 
that at any given Mach number the model drag coefficient during the 
thrusting period was equal to the drag coefficient during the coasting 
period corrected for the pressure drag over the nozzle exit. Since no 
specific base-pressure data were available for this particular model, 
the data presented in reference 3 were used for an approximate correction 
to the drag. Thrust was not corrected for altitude since this correction 
was considered negligible for the altitudes attained by the model. 
The resulting thrust-time curve, as obtained from computing 
instantaneous values of thrust at various times, is presented in figure 9. 
The preignition temperature of the propellant was 75O F. An ignition 
delay of 0.09 second is indicated by the record. A maximum thrust 
of 7200 pounds was reached at 4.8 seconds and the total burning time 
was 7.58 seconds. The total impulse of the rocket is 43,400 pcnmd- 
seconds, giving a specific Fmpulse value of 207 pound-seconds per pound 
of propellant. The thrust-tFme curve and specific-hpulse value are 
in good agreement with data obtained in static firings by the mnufacturer. 
'Selected curves from static tests conducted by the manufacturer are 
shown in figures 10(a) and 10(b) . Results from static test by the 
Aberdeen Proving Ground are presented in figure ll. 
It will be noted in figures 9, 10(a), and 10(b) that the thrust 
changes erratically in the time region 0.7 to 1.8 seoonds and again in 
the region from 4.0 to 5.2 seconds . These thrust irregularities are 
not chmacteristic of single-perforated cylindrical propellant grains. 
Since they occur in flight as well as in static tests, they cannot be 
attributed to flight effects. 
A possible explanation of the origin of these thrust fluctuations 
is the following. The propellant grains are supported by being clamped 
between the three trap assemblies shown as parts 16, 17, and 18 in 
figure l(b). The clamping action is provided by tension within the 
two tie rods connecting the trap assemblies, and possibly to some extent, 
dependent upon assembly clearances, by tension in the chamber producing 
pressure of the nozzle assembly against the rear trap assembly and 
pressure of the boss within the front of the chamber against the front 
trap assembly. Thermal expansion of the tie rods and chamber m y  
release the grains. If a grain becomes displaced an asymmetric flow 
condition exists resulting in increased propellant erosion or other 
anomalies which change the thrust. Presuming the cbmber not to be 
contributing to the clamping because of loose assembly and thermal 
expansion, the tie rod clamping the rear grain may be expected 
to expand more rapidly than the forward tie rod because of being 
subjected to a greater flow of combustion gases, thus permitting 
displacement of the rear grain first and giving rise to the first thrust 
fluctuation. Subsequently, the forward tie rod expands sufficiently to 
permit displacement of the forward grain giving rise to the second 
observed thrust fluctuation. 
Camem Studies of Launchings 
Sequence photographs of the launching and initial part of the flight 
with the corresponding time are shown in figure 12. A study of 
enlargements of these photogkaphs did not reveal any irregularities 
or indications that propellant failure occurred. 
The flight test of the Aerojet m-6000 T-27 Jato rocket-motor 
model indicated the following: 
1. The maximum absolute longitudinal acceleration imposed upon the 
rocket in the test was 9.8g and occurred 2.8 seconds after launching. 
2. The maximum thrust developed was 7200 pounds, and occurred 
4.8 seconds after launching. 
3. The total -Impulse was 43,400 pound-seconds, and the thrusting 
time was 7.58 seconds. 
4. The total impulse is in good agreement and the variation of 
thrust in general agreement with static tests conducted by the 
manufacturer and the Aberdeen Proving Ground. 
5. No thrust anomalies attributable to effects of flight 
longitudinal acceleration were observed. 
6. Certain -11 thrust i r regular i t ies  occurred i n  the f l i g h t  
t e s t  which appear t o  correspond t o  i r regular i t ies  observed i n  some 
s t a t i c  t e s t s  and may be at tr ibutable t o  the txap design. 
Laagley Aeronautical Laboratory 1 
National Advisory Committee f o r  Aeronautics 
Ungley A i r  Force Base, Va. 
C. 
Aleck C . Bond 
Aeronautical. Research Scientis t  
oseph G. Thibodaux, Jr. 
Aeronautical Research Scient is t  
Approved : 
Robert R e  Gilruth 
Chief of Pflot less  Aircraft  Research Division 
NACA RM SLgL13a 
REFERENCES 
. P $ f/. .-- 1-, I. Anon: Jato Ma,nual&~IA/~l> The Johns Hopkins Univ . , Appl. Phy . 
Lab . *-&9:-'- 
2. Pitkin, Marvin, Gardner, William N., and Curfmam,  Howard J., Jr.: 
Results of Preliminary Flight Investigation of Aerodynamic 
Characteristics of the NACA Two-Stage Supersonic Research Model 
RM-1 Stabilized in Roll at Transonic and Supersonic Velocities. 
NACA RM ~ 6 ~ 2 3 ,  1947. 
3. Anon: Survey of Bumblebee Activities. Bumblebee Rep. No. 70, 
The Johns Hopkins Univ . , Appl . Phys . Lab., Nov . 1947. 
NACA FW SLgL13a 
(a) Complete uni t  as supplied by manufacturer. 
Figure 1. - Aero je t o  rocket motor. 
NACA RM SLgL13a 
1. Capscrews and washers for  nozzle assembly. 
2. Nozzle assembly with ign i t e r  leads, 
3. Gasket. 
4. Motor body. 
5. Ring assembly. 
6. Gasket for  r ing assembly. 
7. Retaining sleeve. 
8. Gasket for  retaining sleeve. 
9. Motor body thread protector. 
10, Closing cap gasket. 
11. Closing cap assembly, 
12. Propellant. 
13. Propellant. 
14. Igni te r  assembly. 
15. Nut for  rear  t r ap  assembly. 
16. Rear t rap  assembly. 
17. Center t rap  assembly with t ie rods. 
18. Front t r ap  assembly. 
19. Nut for front  t r a p  assembly. 
(b) Component parts .  
NACA RM SLgL13a 
P
 
B
 
NACA RM SLgL13a 
0 Z 4 6 8 /O 1'2 /# 16 
Tme, t, sec 
Figure 4.- Velocity-time curve, Aerojet Jato ~ K S - ~ O O O  T-27 flight model. 

2 9 
r1 
9 
z 
i? 
\O r 
W 
Zu 
c.5 
\ 
-84 
S 
'h 
$ 3  
2 
/ 
/ 
/ 
/I 
/ T- 
I I I I  
/2 /4 /6 
k e ,  t, sec 
Figure 6. - Altitude-time curve, Aero jet Jato ~KS-6000 T-27 flight model, 
/ 
/ 
/ / 
/ / 
/ 
/, / 
I 
/ / \, 
\ 
\ 
\ 
\ 
\ 
//ZZ 
/o -~  
l 
0 
aJ 
\ 
?///€I 1 1. k  ---- 1- 
6 \\ \ 
S ----- ----- 
2 1/14 
ko ----- 
-------------------- ------ 
2 
F'igure 7. - Variation of density of a i r  and velocity of sound with al t i tude.  
1- 1 1 
I 
De~+~s/f y of cr/r 
M z h  number, M 
Figure 8.d Variation of drag coefficient with Mach number for Jato ~KS-6000 T-27 flight model. 
NACA RM SLgL13a 
8 
7 
6 
5 
4 
K~ 
$3 
S 8 2 
I 
0 
Time, f , sec 
(a) Test 6 4 ~ ~ 4 6 .  
Figure 10.- Thrust-tlme curves obtained from static tests by the manufacturer. 
/ 3 4 5 6 7 8 9 
h e  ,t, sec 
(b) Test 6 4 ~ ~ - 6 7 .  
Figure 10. - Concluded. 
z 
ei 
* 
z 
8 
\D G 
W 
P 
0 / 2 3 ' 4  5 6 7 a 9 
Tme , f , sec 
Figure 11.- Thrust-time curve obtained from s t a t i c  t e s t s  by the Aberdeen Proving Ground. Round 2, 
f i r e d  January 22, 1948. 
NACA RM SLgL13a 
t = 0.38 sec 
t = 0.25 sec 
t = 0.05 sec 
Figure 12.- Sequence photographs 
of take-off of Aerodet 
71~j-6000 T-27 Jato rocket- 
motor flight model. 
t = 0.00 sec 
- 
NACA RM SLgL13a 
t = 1.04 sec 
t = 0.71 sec 
t = 0.58 sec 
Finwe 12. - Concluded . 


Flight Test of the Aerojet 7KS-6000 T-27 Jato Rocket Motor

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server