Expendable infrared source and method therefor

An IR source and method for generating IR radiation whereby a propellant is urned in a first chamber to produce a product gas which is exited through a critical exit, is accelerated to a supersonic velocity by expanding into a second chamber, is passed through a standing shock wave in the chamber to reduce the gas velocity to a subsonic level, is exited through radial orifices into a larger third chamber where the gas is mixed to obtain a substantial uniformity in temperature and specie, is accelerated by expansion to a chosen subsonic velocity, and is exhausted to the atmosphere.

BACKGROUND OF THE INVENTION 
The present invention pertains generally to targets and in particular to a 
source of IR radiation for targets. 
Expendable sources of infrared radiation are required for purposes such as 
target practicing and testing. The IR radiation, in about all 
circumstances, must meet precise requirements as to wave length and 
intensity of radiation and to the spacial distribution of the radiation. 
If the IR source is attached to a tow target, severe restrictions on the 
thrust level must also be met. 
There are four types of expendable IR sources and each has serious 
limitations. Flares have been used as expendable IR sources. They are 
pyrotechnic devices which burn at atmospheric pressure and produce a 
high-intensity, point source of radiation. These devices are generally 
adequate at tail-on aspect angles, but do not adequately simulate the 
scale factors of rocket or turbojet exhaust plumes at side-on aspect 
angles. 
Infrared radiation has also been obtained from seeding torch flames or the 
exhaust of gas generators, either solid or liquid fueled. The advantages 
of this type are high source levels with little propulsive thrust and less 
of a point-source character than flares have. The disadvantages are the 
black-body nature of the radiation, smokiness and inadequate scale factors 
at side-on aspect angles. An example of this type of IR source is found in 
U.S. Pat. No. 3,946,555, issued on Mar. 30, 1976 to Philip J. Goede in 
which inorganic materials, capable of emitting after ejection a continuum 
infrared radiation in the wave length range of about 1 to 14 microns, are 
dispersed in a composite propellant. 
The exhaust from conventional rocket propellants after expanding through 
typical converging-diverging supersonic nozzles does produce IR emissions. 
Unfortunately, to achieve the required plume length and IR-source 
strength, unacceptably high thrust levels are produced which would disturb 
the aerodynamics of the tow target. Also, the required mass-flow rate and 
desired duration may be inconsistent with the payload capability of the 
tow target. Encouraging afterburning to increase IR output would reduce 
mass-flow-rate requirements and thrust level, but afterburning in itself 
is an undesirable characteristic for many applications because it alters 
the spacial distribution of radiation in the plume, which is important for 
side-on aspect angles. 
Another expendable infrared radiation source is a turbojet powered target 
drone which has carbon particles injected into its exhaust plume to 
augment IR output. The disadvantages are black-body spectra as opposed to 
line spectra of the emitted radiation and the smokiness of the exhaust. 
Also, this source of infrared radiation is not applicable to tow targets 
which are preferred over jet-powered target drones because of their 
reduced costs. 
SUMMARY OF THE INVENTION 
It is, therefore, an object of this invention to provide an expendable and 
inexpensive IR radiation source which has, for each given gas-generation 
rate, a minimal reactive thrust and a maximum plume length and integrity. 
Another object of the present invention is to produce IR radiation with a 
predictable IR intensity, i.e., afterburning with the surrounding air is 
minimized. 
A further object of the present invention is to provide line-spectra IR 
radiation as opposed to black-body radiation, i.e., IR emissions from 
gaseous products of combustion as opposed to emissions from solid 
particulate matter in the exhaust stream. 
These and other objects are achieved by combusting a propellant in a 
chamber to produce a hot, high-pressure product gas, exiting the product 
gas through a critical exit so that the combustion process is controlled 
by rocket design principles, shocking down the product gas to a subsonic 
velocity which reduces the stagnation pressure and thus the capability of 
the product to propel, restoring uniform flow conditions, accelerating the 
product gas, and exiting the gas to the atmosphere.

DETAILED DESCRIPTION OF THE INVENTION 
Referring to FIG. 1, IR source 10 is shown to comprise a combustion chamber 
11 containing an energetic composition 19 and having a critical exit 15, a 
shock chamber 13 fixedly connected to and in communication with the 
combustion chamber 11 through the critical exit 15, and a plenum chamber 
21 fixedly connected to and in communication with the shock chamber 13 
through diffuser orifices 17. The plenum chamber 21 exhausts to the 
atmosphere through exhaust means 23. 
The shock chamber 13 is centrally located in the plenum chamber 21. The end 
opposite to exit 15 is preferably blunt because of simplicity of 
construction and near that end is a plurality of radial orifices 17. The 
L/D for the chamber is preferably 1:1 to 5:1 and preferably 2:1 to 5:1. 
The number and size of the orifices are selected to maintain a standing 
shock wave and exhaust the product gases subsonically, preferably at Mach 
0.1 to 0.5. 
The location of the shock wave is between the combustion chamber exit 15 
and the plane of the most upstream radial orifices 17. Since the purpose 
of the shock is to reduce the energy and pressure of the product gas and 
to increase the temperature of the product gas, the preferred position of 
the shock wave is at or near the point of maximum velocity. 
Infrared source 10 operates through the burning of an energetic composition 
19 which produces mostly gases with few solid products in the combustion 
chamber 11. The product gas enters the shock chamber 13 through a critical 
exit 15 which is preferably convergently contoured, e.g., a nozzle. A 
critical exit is one in which the velocity of a fluid passing through has 
a Mach number of 1. 
Upon entering the shock chamber the product gas expands, increasing the gas 
velocity to supersonic levels. The product gases pass down the chamber to 
a standing shock wave created by the restricted outlet, i.e., the diffuser 
orifices 17. The shock chamber, thus, acts as a supersonic diffuser. 
Passage of the product through the standing shock wave is a nonisentropic 
process which results in a large entropy gain, a large loss in stagnation 
pressure and a conservation of stagnation temperature. The static 
temperature and pressure are increased by the decrease in velocity to 
close to their maximum values, i.e., the value at zero velocity. Since the 
entropy gain is proportional to the amount of decrease in the gas 
velocity, it is preferred that the velocity of the product gas reaches at 
least Mach 2 and most preferably at least Mach 3 before passing through 
the standing shock wave. The net result of the shocking/diffusing process 
is the establishment of high-temperature/low-pressure conditions in the 
plenum chamber 21. The high temperature provides the proper IR radiation 
and the low pressure insures little or no propulsion. 
The terms: static temperature and pressure and stagnation temperature and 
pressure are used in their accepted meanings. Stagnation pressure and 
temperature are measured at zero velocity. They represent the maximum 
temperature and pressure of a fluid stream. Static pressure and 
temperature are measured near the boundary layer of a fluid stream and 
represent the temperature and pressure of a stream in which a portion of 
the energy of the system is in the velocity of the stream. 
Upon exiting the shock chamber, the product gas passes through a plenum 
chamber 21. This chamber serves as a mixing chamber where the product gas 
is mixed to obtain a uniform temperature and specie distribution before 
being exhausted to the atmosphere through exhaust 23. The free volume of 
the plenum chamber 21 is sufficient to insure an adequate residence time 
(0.1 to 4 msec) to achieve adequate mixing. The L/D of the plenum chamber 
is from 2:1 to 5:1 and preferably 2.5:1 to 4:1. Plenum chambers larger 
than these dimensions would be effective, but they would not provide any 
benefit for the increased cost, would become too large for many target 
applications and would suffer substantial energy losses to the walls. 
An additional purpose of exhaust 23 is to give a final adjustment to the 
velocity of the product gases, either increasing or decreasing the 
velocity in order to closely match the target speed. If the velocity of 
the product gas exhausting to the atmosphere is greater than the target, 
then thrust to the target would exist, thereby decreasing the stability of 
the target. Also the integrity of the exhaust plume would be too quickly 
destroyed by the shear forces between the relatively stationary atmosphere 
and the fast flowing exhaust. If the velocity of the exhausting gases is 
less than the target, the destruction of the plume integrity would also be 
hastened by the shear forces generated. Consequently exhaust 23 is 
preferably sized to exhaust the product gases at about the velocity of the 
target. The preferred configuration of exhaust 23 is one or more nozzles. 
The single nozzle would produce a longer and narrower plume than would a 
multiple nozzle exhaust. The nozzle shape is preferred because of its 
relatively simple construction and the relatively low energy losses 
through the nozzle. 
FIG. 2 shows a cross section of the IR source 10 in detail that was used in 
the following examples. The outer shell of IR source 10 is manufactured in 
two metal sections 29 and 31 which are joined by a union 33. The 
combustion chamber 11 is placed inside shell 29 and the O-rings 35 
establish a seal between the two. The combustion chamber 11 comprises a 
nozzle 25, a shell 37, an insulator 39 and a spacer 41. A propellant 19 
with an inhibitor liner 43 is placed inside the combustion chamber. 
Between the two sections 29 and 31, an insulating shim 45 and a back-up 
ring 47 are placed. The shock chamber 13 comprises a shell 49, diffuser 
orifices 17 and an anti-erosion disc 51. The plenum chamber 21, defined by 
shell 31, has an insulator 53 and an exhaust nozzle 27. 
The shell 37 defining the combustion chamber 11 is fabricated from a 
reasonably high-strength material, such as steel. The surface of shell 37 
is protected from the heat and corrosive effects of the reaction by the 
insulator 39 made from a suitable material, e.g., phenolic or polyepoxy 
plastic. At the exhaust port a converging nozzle 25 is fabricated from an 
inert, high-temperature material, such as graphite. It is sized to produce 
critical flow into the shock chamber 13. Between the propellant 19 and the 
shell 37 along with the insulator 39 is a spacer 41 made from silicon 
rubber or a similar inert resilient material. An insulating shim 45 made 
from phenolic or another suitable material is placed between sections 29 
and 31 in order to reduce heat transfer from the combustion chamber 11. To 
give added strength to the insulating shim 45, a steel or similar material 
back-up ring 47 is added to the shim. 
The shell 49 of the shock chamber is fabricated from molybdenum or another 
high-temperature material. The closed end of the shock chamber 13 is 
protected by an anti-erosion disc 51 made preferably from graphite. A 
plurality of orifices 17 is placed radially in shell 49 as diffuser 
orifices 17. Orifices 17 open into a larger chamber referred to as a 
plenum chamber 21. The shell 31 which defines the plenum chamber 21 is 
protected by an insulating layer 53 which can be made from a high 
temperature material such as a phenolic or polyepoxy plastic. The gases in 
the plenum chamber 21 exit through nozzle 27 which is preferably made from 
graphite. 
In order to demonstrate the effectiveness of the present invention the 
following examples are given. It is understood that these examples are 
given by way of illustration and are not intended to limit this disclosure 
or the claims to follow in any manner. 
EXAMPLE I 
The IR source of this invention which was tested was 50 cm. in overall 
length and 15 cm. in diameter. The shock chamber had a length of 15 cm., 
an inner diameter of five cm., a sonic (critical) inlet nozzle with a 
cross-sectional area of 0.13 sq. cm., and eight diffuser orifices with a 
total cross-sectional area of one sq. cm. The plenum chamber had an inner 
diameter of nine cm., a length of twenty cm. and an exit nozzle with an 
inner diameter of five cm. at the exit. 
The propellant used in the experiment was the standard AHH cast double-base 
propellant, which comprise 40 weight percent nitroglycerin, 40 weight 
percent nitrocellulose, and 20 weight percent of other ingredients. The 
AHH designation is the one given by the Chemical Propulsion Information 
Agency (CPIA) in their specifications and standards for approved 
propellants. The propellant was inhibited by cellulose acetate and 
asbestos phenolic was used for motor insulation. The propellant generated 
a product gas at a rate of 0.21 kg-sec at 50 atm. and had a flame 
temperature of about 2300.degree. C. and a C.sub.p /C.sub.v of 1.24. About 
400 gm was used in the firing. 
The product gas, after exiting the sonic inlet nozzle, expanded to the 
5-cm. diameter shock-chamber wall and reached a Mach number of about 4.4 
with a static temperature drop of about 500.degree. C. A strong normal 
shock was established in the flow passage of the shock chamber. At 
steady-state operation the normal shock stood off the end plate of the 
shock chamber far enough to have allowed a 0.21 kg/sec gas-flow rate to 
exit the shock chamber radially through the diffuser orifices. At steady 
state operation, the product gas, after the shock wave, had a Mach number 
of 0.4, a static pressure of about two atmospheres and a static 
temperature of about 2100.degree. C. 
Shock chamber and plenum chamber conditions were calculated, as were 
exhaust plume length and IR intensity. Performance of the IR source was 
then determined by comparing the observed plume length and intensity (4-6 
foot length, 400 watts/steradian @ 4.fwdarw.5 microns) with calculated 
values. Exhaust plume characteristics were grossly different from what 
this propellant and gas generation rate would produce if the product gases 
were exhausted through a standard convergent/divergent supersonic rocket 
nozzle. Observations were made with IR radiometers (for intensity) and 
Thermo-Vision for IR plume length. Observed values agreed closely with 
calculated IR source performance. 
EXAMPLE II 
The exit conditions of the exhaust gases of the previous example were 
compared to a rocket firing the same propellant. Table I summarizes the 
results. 
TABLE I 
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Units I.R. Rocket 
______________________________________ 
Static Temp. .degree.C. 2253 1122 
Static Press atm 1.8 1.8 
Stagnation Temp 
.degree.C. 2329 2329 
Stagnation Press 
atm 2.1 47.6 
Mach No. -- 0.5 2.7 
Thrust kg 5.6 13 
Plume Length m 1.8 0.14 
Velocity m/sec 519 2827 
______________________________________ 
The results from Examples I and II confirm that passing through a normal 
standing shock wave is a non-isentropic process which results in a large 
entropy gain and a large loss in stagnation pressure. Stagnation 
temperature is not lost, however, and the net result of the 
shocking/diffusing process is the establishment of 
high-temperature/low-pressure conditions in the plenum chamber. 
Accelerating the product gas through the exhaust nozzle results in a 
high-thermal energy, low-kinetic-energy exhaust stream which delivers a 
much lower propulsive force (less than one half) than that of a normal 
rocket of the same mass-flow rate and exit static pressure. The high exit 
temperature (more than double) and slow mixing rates of the shock-free 
exhaust plume result in a longer (nearly 13.times.) more intense infrared 
plume than that which is produced by a normal rocket at the same mass-flow 
rate. 
Many obvious modifications and embodiments of the specific invention, other 
than those set forth above, will readily come to mind to one skilled in 
the art having the benefit of the teachings presented in the foregoing 
description and the accompanying drawings of the subject invention and 
hence it is to be understood that such modifications are included within 
the scope of the appended claims.