A nozzle assembly for a rocket type decoy used to protect aircraft from missile attack. The nozzle assembly includes an annular chamber and a mixing chamber which is spaced radially inward from the annular chamber. The chambers are separated by a wall and interconnected by multiple jet ports formed in the wall. The nozzle assembly is positioned adjacent to a conventional rocket combustion chamber such that when energetic compositions are reacted in the chamber, the combustion products flow into the annular chamber, radially inward through the multiple jet ports, and into the mixing chamber. Chaotic mixing of the combustion products occurs in the mixing chamber to improve combustion efficiency and reduce combustion products pass from the mixing chamber through the nozzle exit plane thus producing reduced thrust with enhanced radiation intensity.

BACKGROUND OF THE INVENTION 
The present invention relates to rocket motor assemblies generally, and 
more particularly, to a thrust reducing nozzle assembly for a rocket 
decoy. 
Infrared-radiation decoys have been used to protect friendly aircraft from 
attack by enemy missiles. An effective decoy must match the radiation 
signature and flight parameters of the aircraft it is designed to protect. 
The intensity, wavelength and spacial distribution of radiation generated 
by the decoy must closely match that of the aircraft. Additionally, there 
are limitation on thrust levels. 
Flares are one type of infrared-radiation decoy. Flares are pyrotechnic 
devices which burn at atmospheric pressure and produce a high-intensity 
point of radiation. However, since flares are non-propulsion devices, they 
do not escort the aircraft, but rather free-fall away from it. 
Accordingly, flares provide protection for only a short interval of time. 
Rocket decoys can offer advantages over flares. Rocket type decoys are 
designed to generate thrust as opposed to a zero thrust flare. In this 
manner, the rocket decoy can fly behind and escort an aircraft over an 
extended period of time. However, it has been found that rocket decoys are 
limited by certain factors. These factors include the inability to achieve 
maximum combustion temperature (i.e., exhaust plume intensity) due to 
inadequate mixing of combustion species. This creates problems in 
simulating the exhaust plume of the aircraft to be protected. 
Additionally, the mass flow rate required to achieve the desired exhaust 
plume intensity can produce a thrust level too high for best performance. 
Specifically, an excessive thrust level can cause the decoy to overtake 
the aircraft it is intended to protect. One attempt to reduce the rocket 
thrust was to arrange the exhaust nozzles of the rocket motor assembly 
outboard, away from the flight center line to direct exhaust jets radially 
outward. However, this approach has been found to dissipate the exhaust 
flow, resulting in an undesirable reduction in exhaust plume intensity. 
Thus, there is a need to provide a rocket motor assembly having a nozzle 
that facilitates increased combustion and exhaust plume temperatures, 
while enabling a thrust level reduction. 
SUMMARY OF THE INVENTION 
The present invention is directed to a rocket motor nozzle that avoids the 
problems and disadvantages of the prior art. The invention accomplishes 
this goal by providing a nozzle construction that increases combustion 
efficiency, while permitting thrust reduction. The nozzle has an outer 
generally cylindrical wall and a cup-shaped member disposed therein. The 
cup-shaped baffle includes a generally cylindrical portion which is spaced 
radially inward from the generally cylindrical outer wall of the nozzle, 
thereby forming a channel or plenum chamber therebetween. The cylindrical 
portion also includes multiple ports to provide fluid communication 
between the plenum chamber and the interior of the cup-shaped member for 
mixing and subsequent discharge. In operation, the nozzle is positioned 
downstream from a conventional rocket combustion chamber with the closed 
end of the cup-shaped member facing the chamber. As propellant burns in 
the combustion chamber the combustion products flow into the outer channel 
of the nozzle, then radially inward through the multiple nozzle jet ports 
and into the mixing chamber. In the mixing chamber, the jet streams 
impinge upon each other to cause chaotic mixing, which improves combustion 
efficiency, but does not enhance thrust. Thrust in fact is reduced. The 
combustion products then exit from the mixing chamber through the open end 
of the cup-shaped member. 
Another important feature of the invention is that the cylindrical portion 
of the cup-shaped member is concentrically positioned within the outer 
wall of the nozzle. This ensures flow equilibrium. Additionally, the ports 
are substantially equidistantly spaced in the circumferential direction of 
the cylindrical portion. With this configuration the net direction of the 
combustion jets in the mixing chamber is zero. This advantageously 
minimizes the thrust of the rocket decoy. 
As evident from the foregoing, the nozzle of the present invention improves 
decoy performance by inducing turbulent mixing of the combustion products, 
resulting in increased reaction temperature and exhaust plume intensity. 
It also allows the propellant to burn at selected design pressures 
consistent with reliable and reproducible operation, while producing lower 
thrust levels than can be achieved with conventional nozzles 
The above is a brief description of some deficiencies in the prior art and 
advantages of the present invention. Other features, advantages and 
embodiments of the invention will be apparent to those skilled in the art 
from the following description, accompanying drawings and appended claims.

DESCRIPTION OF THE PREFERRED EMBODIMENT 
Referring to the drawings in detail, wherein like numerals indicate like 
elements, rocket motor assembly 1 is illustrated in accordance with the 
principles of the present invention. Rocket motor assembly 1 generally 
includes combustion chamber portion 2 and nozzle portion 4 through which 
high pressure combustion products or gases generated in combustion chamber 
2 exit the rocket motor assembly. 
Referring to FIG. 1, combustion chamber 2 and nozzle 4 are positioned in a 
conventional rocket motor casing 5 and secured thereto as is known in the 
art (nozzle 4 can be adhesively bonded to casing 5, for example). 
Combustion chamber 2, which is of conventional construction, includes 
cup-shaped liner 6 for containing an energetic composition 10 and 
insulating casing 5 as is known in the art. Although energetic composition 
10 is shown as being in tube form, other conventional configurations and 
arrangements can be used, for example, without departing from the scope of 
the present invention. The materials used to make up cop-shaped liner 6 
can vary widely as will be evident to one skilled in the art. Materials 
used to construct cup-shaped liner 6 include phenolics, for example. 
Referring to FIGS. 1 and 2, nozzle assembly 4 includes an outer preferably 
cylindrical shell 8 and a cup-shaped baffle 12 positioned therein. Shell 8 
abuts cup-shaped liner 6 as is conventional in the art. Cup-shaped member 
12 includes a preferably cylindrical portion 14 and an imperforate end 
wall or baffle 16 which form mixing chamber 18. Cylindrical portion 14 is 
spaced radially inward from outer shell 8 to form annular channel or 
plenum 20. Cylindrical portion 14 also includes multiple ports 22 which 
provide fluid communication between annular channel 20 and mixing chamber 
18. Annular wall 24 extends between outer shell 8 and cylindrical portion 
14 at a location downstream from ports 22. Annular wall 24 ensures that 
the combustion products entering annular channel 20 do not escape the 
motor assembly before entering mixing chamber 18. 
In operation, energetic combustion 10 is reacted to produce combustion 
products that include a hot, high-pressure product gas. Baffle 16 causes 
the product gas or combustion products generally designated with reference 
numeral 26 to flow into annular chamber 20 from which the combustion 
products flow through the multiple nozzle jet ports 22 into the mixing 
chamber 18. In the mixing chamber, the jet streams, generally illustrated 
by the arrows 28, impinge upon each other to cause chaotic mixing which 
improves combustion efficiency. The combustion products then pass from the 
mixing chamber through the exit plane or open end 30 of cup-shaped member 
12, producing reduced thrust with enhanced radiation intensity. 
Referring to FIG. 2, cylindrical portion 14 is concentrically positioned 
within outer shell 8 and jet ports 22 are equidistantly spaced along the 
circumference of cylindrical portion 14. This arrangement balances the jet 
stream energy entering mixing chamber 18. As a result, the net direction 
of jets 28 is zero, thereby minimizing thrust. 
Conventional propellants can be used in conjunction with the present 
invention. Typically, such propellants comprise aluminum and an oxidizer. 
However, propellants comprising magnesium as opposed to aluminum provide a 
brighter plume and can burn at a variety of pressures to facilitate 
matching the radiation signature and flight parameters of the aircraft 
that the decoy it is designed to protect. Accordingly, propellants 
comprising magnesium may be preferred. However, it should be understood 
that the specific propellant composition is selected based upon the 
intensity, wavelength and spacial distribution of radiation generated by 
the exhaust plume of the specific aircraft that the decoy is designed to 
closely match. 
The sizes and materials used to make up the nozzle assembly can vary widely 
as will be evident to one skilled in the art. Materials used to construct 
the nozzle assembly may include graphites and reinforced phenolics, for 
example. The exact number, size and placement of the jet ports as well as 
the dimensions of the annular chamber and mixing chamber are determined 
according to the mass flow rate desired to exit the nozzle. That mass flow 
rate should provide the desired flight parameters and the radiation 
signature of the aircraft the decoy is designed to protect. The flow areas 
are optimized as is conventional in the art. In all cases, however, the 
cross-sectional areas of the annular chamber and mixing chamber must each 
be greater than the total jet port flow area. Additionally, the jet ports 
can be aligned with their center axes perpendicular to the nozzle center 
line as illustrated in FIG. 1, or they may be slightly canted rearwardly 
up to about 10 degrees. 
Merely to exemplify a preferred make up of the nozzle assembly, the 
following example may be recited. It is understood that this example is 
given by way of illustration and not intended to limit the scope of this 
invention. 
EXAMPLE 1 
The tested nozzle assembly of this invention was constructed of ATJ 
graphite. Four equally spaced 0.25-inch diameter jet ports were used. The 
cross-sectional area of the annular chamber and exit plane of the 
cup-shaped member were 7 and 12 times greater than the total jet port 
area, respectively. The mixing chamber length to diameter ratio (L/D) was 
0.55. Specifically, the mixing chamber length was about 1 inch and its 
diameter about 1.8 inches. The axial length of the nozzle was about 11/2 
inches. The nozzle was spaced axially from the propellant about 1/2 inch. 
The combustion chamber had an inner diameter of about 3 inches. The 
propellant was tubular in configuration with a wall thickness of about 1 
inch and length of about 61/2 inches. The propellant used comprised 
approximately 40 percent by weight magnesium, approximately 40 percent by 
weight ammonium perchlorate and approximately 20 percent by weight 
hydrocarbon binder to bind the magnesium and ammonium perchlorate powders 
together. It was found that the thrust delivered by this nozzle was 
approximately 1/2 the minimum value that could have been achieved with a 
conventional nozzle. Measurements also indicated that the exhaust plume 
intensity was extraordinarily high. 
The above is a detailed description of a particular embodiment of the 
invention. It is recognized that departures from the disclosed embodiment 
may be made within the scope of the invention and that obvious 
modifications.,will occur to a person skilled in the art. The full scope 
of the invention is set out in the claims that follow and their 
equivalents. Accordingly, the claims and specification should not be 
construed to unduly narrow the full scope of protection to which the 
invention is entitled.