The starting transients of solid propellant engines were investigated to develop design principles for predicting transients in high performance igniter and engine configurations. Research and diagnosis were conducted on the processes affecting ignition transients, such as heat flux, igniter, flame spreading over the surface, and nonsteady combustion gas dynamics. Abstracts of published reports related to this research are presented. Topics discussed include: development of an analytical model, analytical prediction of the entire ignition transient, transient behavior of pressure and regression rate during reignition after shut-down, and problems associated with restarting hybrid rocket engines

Caveny, L. H.

Summerfield, M.

English

NASA Technical Reports Server

(NASA-CR-14C800) INVESTIGATION OF THE N75-12060
STARTING TRANSIENTS OF HIGH PERFORMANCE
iSOLID-PROPELLANT MOTORS Final Report, 1
;Oct. 1967 - 30 Sep. 1974 (Princeton Unclas
dUniv.) 12 p HC $3.25 _ . CSCL 21H G3/20 02798
CLOSE OUT REPORT
for
INVESTIGATION OF THE STARTING
TRANSIENTS OF HIGH PERFORMANCE
SOLID-PROPELLANT MOTORS
Conducted Under NASA Grant NGR 31-001-109
From 1 October 1967 to 30 September 1974
Co-Authors:
Leonard H. Caveny
Senior Member of Professional Staff
Martin Summerfield
Professor of Aerospace Propulsion
Submitted by:
Principal Investigator "
October 1974
Department of Aerospace and Mechanical Sciences
Princeton University
Princeton, New Jersey
https://ntrs.nasa.gov/search.jsp?R=19750003988 2020-03-23T03:20:48+00:00Z
INDEX
Page
Title Page
DESCRIPTION OF RESEARCH PROGRAM 1
SUMMARY OF RESEARCH RESULTS 2
ABSTRACTS OF:
"Thrust Transient Prediction and Control of Solid
Rocket Engines". 4
"Starting Thrust Transients of Solid Rocket Engines". 5
"Combustion Anomalies in Stop-Restart Firing of
Hybrid Rocket Engines". 6
"Restart Transients of Hybrid Rocket Engines". 7
"The Starting Transient of Solid Propellant Rocket
Motors with High Internal Gas Velocities". (AMS) 8
"Thrust Transients of Large Solid Rocket Motors". 9
"The Starting Transient of Solid Propellant Rocket
Motors with High Internal Gas Velocities". (AIAA) 10
-1-
DESCRIPTION OF RESEARCH PROGRAM
The initial studies under the grant (from 1967 to 1969)
produced significant progress in the investigation of the
fundamental processes that occur during the starting transient
period of solid propellant motors with low loading densities
and small length-to-diameter ratios. The program of research
culminated in an investigation of the processes, which occur
during the starting transients of high performance solid pro-
pellant motors. This is the practical class of motors which
have high loading densities, large length-to-diameter ratios,
and low port-to-throat area ratios.
The purposes of this program was to obtain better under-
standing of the starting transients of solid propellant motors
and to develop design principles for the prediction of the
transients in high performance igniter and motor configura-
tions. Concurrently, basic research and diagnosis were con-
ducted on the processes affecting ignition transients such as
heat flux to the propellant, propellant response to the igniter
stimulus, flame spreading over the propellant surface and
nonsteady combustion gas dynamics of the entire motor.
The program consisted of the following parts:
1) Development of theoretical models for the starting
transient of solid propellant motors.
2) Design and manufacture of instrumented rocket motors
equipped with transparent windows, heat flux gauges,
pressure transudcers, etc.
-2-
3) Diagnostic experiments to determine such items as the
flame spread rate, ignition delay times, flow patterns,
erosive burning rates and heat flux distribution.
4) Revision of theoretical models based on results of
diagnostic experiments.
5) Establishment of prediction and design rules.
6) Rocket motor firings to compare the predicted
pressure transients with the actual pressure transients.
7) Quantitative analysis and interpretation of the firing
records of service rocket engines to the extent that
they were available, for the purpose of testing the
applicability of the design rules derived from the
program.
8) As a separate topic, investigation of the restart
transients of hybrid rocket motors.
-3-
SUMMARY OF RESEARCH RESULTS
During the seven year period that includes 1967 and 1974
the results of our research on solid propellant rocket motor
ignition and propulsion related topics were described in
reports and technical papers. These publications form a set
of self-consistent sources on the broad topic of solid rocket
motor ignition. This section contains abstracts of the pub-
lications.
The publications summarized in this document are the
technical reports and papers which have been archived by the
appropriate libraries and agencies. Some types of publica-
tions (i.e., preprints of papers later published in journals,
progress reports which were superseded by final reports,
administrative summary reports which merely summarize publi-
cations listed in this document, and informal presentation
summaries) have not been included if the results reported in
them are also contained in a more comprehensive archive pub-
lication. It should be noted that the interplay among the
publications is great since there has been a commonality of
propellants, fuels, rocket motors, data reduction techniques,
etc. throughout our investigations.
-4-
"THRUST TRANSIENT PREDICTION AND CONTROL OF SOLID ROCKET
ENGINES"
W. J. Most, B. W. MacDonald, P. L. Stang and M. Summerfield
WSCI Paper 68-33, The Combustion Institute, October 1968.
The research reported in this paper is directed toward
the development of an analytical model for predicting the
thrust-time curve during the entire ignition transient of a
solid propellant motor. This model characterizes the local
ignition event by a constant critical surface ignition tem-
perature and by including a surface heat release term to ac-
count for exothermic decomposition below the critical tempera-
ture. Flame spreading is described by coupling this ignition
model with an empirical description of the gas phase heat con-
vection to the propellant grain. The propellant burning rate
after ignition is achieved is taken as the steady state burn-
ing rate. The model is completed by writing the dynamic ener-
gy and continuity equations. This model is compared to others
appearing in the literature.
The theoretical predictions of the model are compared
to experimental test firings of a two-dimensional rocket
motor and of motors with more realistic grain configurations.
An extensive range of igniter system and engine design param-
eters are covered. Both aluminized and unaluminized propel-
lants are considered and shown to be similar. The abrupt
shift from a vigorous, well-ignited rocket motor to an extend-
ed hangfire or misfire can be caused by small changes in igni-
ter design parameters. This sensitivity is predicted theoret-
ically and demonstrated experimentally.
This paper concludes that a physically rational theory
has been developed that can predict the entire motor ignition
transient for motors with head-end pyrogen igniters. This is
verified experimentally.
Based on work performed under NASA Grant NGR-31-001-109
supervised by Langley Research Center.
Accession No. A69-18365*# - Available from AIAA.
-5-
"STARTING THRUST TRANSIENTS OF SOLID ROCKET ENGINES"
W. J. Most and M. Summerfield
Aerospace and Mechanical Sciences Report No. 873, July 1969,
Princeton University, Princeton, N.J.
In the past, the design engineer has been forced to rely on
empirical and statistical knowledge of previous firings in
order to design the ignition system of a solid rocket motor or
to predict the starting delay time of a given rocket-igniter
system. In order to predict the entire ignition transient
analytically, the processes of local ignition, heat transfer
and subsequent flame propagation and the gas dynamics of the
combustion chamber must all be described quantitatively. Each
of these elements has been the focus of extensive research in
itself. The purpose of this paper is to present the results
of these research efforts, in our laboratory and elsewhere, in
relation to the objective of predicting analytically the entire
ignition transient.
A particular analytical model, developed for the class of
engines with large port-to-throat area ratios and head-end
mounted pyrogen igniters, is presented. This model character-
izes the local ignition event by ascribing to the propellant
a critical surface temperature for ignition and by including
a surface heat release term to account for exothermic decom-
position while the surface is still below the critical tempera-
ture. Flame spreading is described by coupling this ignition
model with a general description of the heat transfer from
the gas phase to the unignited propellant grain. Any. propel-
lant burning rate law, steady or nonsteady, can be used once
ignition has been achieved. The model is completed by the
dynamic energy and continuity equations for the motor free
volume. This particular model is compared to others which have
appeared in the literature, with special attention paid to
those reports in which comparisons between theoretical pre-
dictions and experimental test firings are offered.
The limitations of the various models are examined, es-
pecially with regard to those particular assumptions which
cannot be justified experimentally. The applicability of the
various models for the prediction of marginal (hangfire)
situations is examined.
This paper concludes that the analytical models now avail-
able can provide useful predictions of the entire ignition
transient, at least for the class of motors for which they
have been developed, and this has been verified experimentally.
Based on work performed under Grant NGR 31-001-109 sponsored
by the Office of Advanced Research and Technology, NASA,
supervised by Langley Research Center.
Accession No. N72--75433 - Available from NTIS.
-6-
"COMBUSTION ANOMALIES IN STOP-RESTART FIRING OF HYBRID
ROCKET ENGINES"
M. A. Saraniero, L. H. Caveny and M. Summerfield
Aerospace and Mechanical Science Report No. 945, Sept. 1970,
Princeton University, Princeton, N.J.
An experimental investigation of the restart process of
an oxygen-Plexiglas hybrid rocket demonstrated that preheating
the fuel (as a consequence of a previous ignition and a tem-
porary extinguishment) significantly increases the rate of
chamber pressurization and produces regression rate overshoots
during reignition. Higher rates of chamber pressurization
measured during the restart transient imply faster instantan-
eous regression rates of the fuel during restart. Transient
periods following the initial ignition and the restart after
a two-second shut-down were observed for four experimental
tests which were conducted at two oxidizer flow rates, 0.078
ibm/sec and 0.039 lbm/sec, and two chamber pressures, 80 psig
and 320 psig. Thermocouples made from 0.001 inch diameter
chromelalumel wire were embedded in the fuel to record sub-
surface temperature histories.
A mathematical model of the thermal processes in the fuel
and the events that occur during an experimental firing was
developed to calculate the regression rate transients. Quan-
titative agreement with experimental results was obtained by
making nominal corrections to the calculated convective and
radiative heating rates. A parametric study investigated the
influence of the fuel's thermal characteristics (thermal con-
ductivity, surface temperature, and heat of gasification) and
the shut-down duration on the transient responses. The
calculated results showed that instantaneous regression rate
overshoots were as high as 29% when high energy fuels are
reignited.
The results of this study indicate that the transient
behaviors of pressure and regression rate during reignition
after shut-down (the duration of which has been varied over
a range of 0.9 to 4.0 sec) can be anticipated and predicted
once the regression characteristics of individual fuel/oxidizer
combinations have been established. Thus, extensive experi-
mental firings of full scale rocket motors will not be re-
quired to adequately predict pressure-time histories and fuel
consumption following reignition if proper attention is given
to understanding the thermal distributions during shut-down
periods.
Based on work performed under Contract NGR--31-001-109 issued
by Office of Advanced Research and Technology, NASA.
Accession No. N72-26960 -- Available from NTIS.
-7-
"RESTART TRANSIENTS OF HYBRID ROCKET ENGINES"
M. A. Saraniero, L. H. Caveny and M. Summerfield
Journal of Spacecraft and Rockets, Vol. 10, No. 3, March
1973, pp 215-217.
The problems associated with restarting hybrid rocket
motors (i.e., rocket motors wherein a liquid or gaseous
oxidizer is injected into the port of a solid fuel grain with
subsequent mixing and combustion of the oxidizer and fuel)
following a brief period of extinguishment were investigated
experimentally. In the extreme, reignition and re-establish-
ment of the oxidizer flow of a briefly extinguished hybrid
motor which is heated in depth and undergoing subsurface
reactions will produce catastrophic increases in burning
rate. If a precise Pch vs t program is required, results
demonstrate that the designer must carefully program the
external ignition stimulus and oxidizer flow rate during the
restart ignition period by taking into consideration at least
three factors: (1) pressure prior to extinguishment, (2)
duration of shutdown interval, and (3) desired pressure
following restart. The measured pressure vs time traces
explicitly show the extent to which restart ignition differs
from the first ignition. Increasing mQ greatly reduces
the t9 on reignition. The time to achieve 900/o of the oper-
ating pressure t9 is reduced by as much as 88%/, and the time
to achieve ignition is reduced by as much as 45%. Also the
measured pressure time traces reveal that there is no appre-
ciable increase in burning rate after the thermal profile is
fully established. This indicates that the specific energy
(approximately 350 cal/g) required to gasify PMM is much
larger than the specific energy increase attributed to
indepth absorption of thermal radiation (30 cal/g) during
the time interval of the test.
Based on work performed under sponsorship by NASA's Office of
Advanced Research and Technology under Grant NGL-31-001-109
and monitored by the Applied Rocket Research Section of the
NASA Langley Research Center.
Accession No. A73-26669# - Available from AIAA.
-8-
"THE STARTING TRANSIENT OF SOLID-PROPELLANT ROCKET MOTORS
WITH HIGH INTERNAL GAS VELOCITIES"
A. Peretz, L. H. Caveny, K. K. Kuo and M. Summerfield
Aerospace and Mechanical Science Report No. 1100, April 1973,
Princeton University, Princeton, N.J.
A comprehensive analytical model which considers time
and space development of the flow field in solid propellant
rocket motors with high volumetric loading density is described.
The gas dynamics in the motor chamber is governed by a set of
hyperbolic partial differential equations, that are coupled
with the ignition and flame spreading events, and with the
axial variation of mass addition. The flame spreading
rate is calculated by successive heating-to-ignition along
the propellant surface. Experimental diagnostic studies have
been performed with a rectangular window motor (50 cm grain
length, 5 cm burning perimeter and 1 cm hydraulic port diam-
eter), using a controllable head-end gaseous igniter. Tests
were conducted with AP composite propellant at port-to-throat
area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pres-
sures from 35 to 70 atm. Calculated pressure transients and
flame spreading rates are in very good agreement with those
measured in the experimental system.
Based on work performed under Research Grant NGL-31-001-109
sponsored by the Office of Advanced Research and Technology,
NASA.
To be cataloged by NTIS.
-9-
"THRUST TRANSIENTS OF LARGE SOLID ROCKET MOTORS"
Leonard H. Caveny, Arie Peretz and Martin Summerfield
Proceedings of the 10th JANNAF Combustion Meeting, Aug. 1973
at Newport, R. I. CPIA Pub. 243, Vol. I, Dec. 1974, pp. 21-44.
An analytical-model which considers time and space
development of the flow field in solid-propellant rocket
motors with high volumetric loading density was used to pre-
dict the ignition characteristics of larger solid rocket
motors (e.g., thrust > 2 x 106 lb and length > 100 ft).
Uncertainties associated with the prediction of peak pres-
sures are attributed largely to incomplete knowledge of
erosive burning. The calculated results demonstrate that
the flow field from large solid propellant motor grains
overpower the uncertainties associated with the detailed
ignition processes. Sensitivity analyses revealed that the
ignition and pressurization times were not appreciably
affected by the expected variations in motor parameters,
if the pyrogen produced sufficiently high convective heating
rates. Accordingly, small (< 0.1% of total motor weight)
head-end mounted pyrogens which promote a rapid onset of
flame spreading are expected to produce rapid and repro-
ducible chamber pressurization.
Based on work performed under NASA Grant NGL 31-001-109
monitored by NASA Langley Research Center and the Jet Pro-
pulsion Laboratory of the California Institute of Technology.
Available only from CPIA.
-10-
"THE STARTING TRANSIENT OF SOLID PROPELLANT ROCKET MOTORS
WITH HIGH INTERNAL GAS VELOCITIES"
A. Peretz, K. K. Kuo, L. H. Caveny and M. Summefield
AIAA Journal, Vol. 11, No. 12, Dec. 1973, pp. 1719-1727.
A comprehensive analytical model which considers time
and space development of the flow field in solid-propellant
rocket motors with high volumetric loading density is described.
The gas.dynamics in the motor chamber is governed by a set of
hyperbolic partial differential equations, that are coupled
with the ignition and flame spreading events, and with the
axial variation of mass addition. The flame spreading rate
is calculated by successive heating-to-ignition along the
propellant surface. Experimental diagnostic studies have
been performed with a rectangular window motor (50 cm grain
length, 5 cm burning perimeter and 1 cm hydraulic port diam-
eter), using a controllable head-end gaseous igniter. Tests
were conducted with AP composite propellant at port-to-throat
area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures
from 35 to 70 atm. Calculated pressure transients and flame
spreading rates are in very good agreement with those meas-
ured in the experimental system.
Based on work performed under NASA Grant NGL-31-001-109
monitored by NASA Langley Research Center and the Jet Propul-
sion Laboratory of the California Institute of-Technology.
The above cited article supersedes AIAA Paper 72-1119.
Available from AIAA as Accession No. A73-14909.


Investigation of the starting transients of high performance solid-propellant motors

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server