US 20090211225A1

(19) United States

(12) Patent Application Publication (10) Pub. No.: US 2009/0211225A1
Nyberg et al. (43) Pub. Date: Aug. 27, 2009
(54) SYSTEMS AND METHODS FORVARYING (21) Appl. No.: 11/699.617
THE THRUST OF ROCKET MOTORS AND
ENGINES WHILE MANTAINING HIGHER (22) Filed: Jan. 29, 2007
EFFICIENCY USING MOVEABLE PLUG
NOZZLES Publication Classification

(75) Inventors: Donald Gerrit Nyberg, Redmond,

(51) Int. Cl.
FO2K 9/86 (2006.01)
WA (US); Thomas Adrian
Groudle, Redmond, WA (US);
Richard Doyle Smith, Kirkland,
t 3. :08:
WA (US); John A. Shuba, (52) U.S. Cl. .................. 60/242: 60/235; 60/240; 60/771
Spokane, WA (US); Richard T.
Smith, Kirkland, WA (US) (57) ABSTRACT
Correspondence Address: The thrust of a rocket motor can be varied to optimize Nozzle
JENSEN & PUNTIGAM, P.S. Pressure Ratio (NPR) using a design that allows for adjusting
2033 SIXTHAVENUE, SUITE 1020 the relative position of a plug and a combustion chamber exit.
SEATTLE, WA 98121-2527 (US) The plug or the exit may be attached to an adaptive control
system for position modification. The relative position of the
(73) Assignee: GHKN Engineering, LLC, plug and exit may be adjusted to optimize NPR to account for
Redmond, WA (US) changing propellant flow and/or changing ambient pressure.

Liquid propellant
rocket engine 401
Combustion
chamber 403 Moveable Exit 405
plug 404
Propellant flow
control device
402

22
%2 Spike
406

Insulation 408
Chamber Support Actuator
Altimeter / Automated barometer Structure 409
barometer control 410 407
41 component
42
Patent Application Publication Aug. 27, 2009 Sheet 1 of 6 US 2009/0211225A1

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Patent Application Publication Aug. 27, 2009 Sheet 2 of 6 US 2009/0211225A1

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Patent Application Publication Aug. 27, 2009 Sheet 3 of 6 US 2009/0211225A1

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Patent Application Publication Aug. 27, 2009 Sheet 4 of 6 US 2009/0211225A1

quod nS
Patent Application Publication Aug. 27, 2009 Sheet 5 of 6 US 2009/0211225A1

Predict / Sense
Decrease propellant decrease in chamber Predict / Sense increase
flow rate in ambient pressure
pressure
501 511 521

Move plug to decrease Move plug to decrease Move plug to decrease

throat area throat area throat area
502 512 522

500 510 52O

Predict / Sense
increase propellant flow Predict / sense increase
decrease in ambient
rate in chamber pressure pressure
551 561 571

Move plug to increase Move plug to increase Move plug to increase

throat area throat area throat area
552 562 572

550 560 570

FIG. 5
Patent Application Publication Aug. 27, 2009 Sheet 6 of 6 US 2009/0211225A1

AREA RAO AJA,

FIG. 6
US 2009/0211225 A1 Aug. 27, 2009

SYSTEMS AND METHODS FORVARYING and the rocket becomes less efficient. The centerbody, how
THE THRUST OF ROCKET MOTORS AND ever, increases the pressure of the exhaust gases by Squeezing
ENGINES WHILE MANTAINING HIGHER the gases into a smaller area thereby virtually eliminating any
EFFICIENCY USING MOVEABLE PLUG loss in thrust at low altitude.”
NOZZLES 0007 Liquid propellant engines have improved perfor
mance over a wide range of Nozzle Pressure Ratios (NPRs)
FIELD OF THE INVENTION using systems such as those described in Sutton and Huzel
0001. This invention relates to rocket propulsion, and and Huang. A recent improvement is described in U.S. Pat.
more particularly to controlling the thrust of a rocket engine No. 6,591,603 B2, granted Mar. 13, 2003 to Gordon A.
or rocket motor and maintaining the thrust efficiency of the Dressler, Thomas J. Mueller, and Scott J. Rotenberger,
system. entitled “Pintle Injector Rocket With Expansion-Deflection
Nozzle' (hereinafter “Dressler'). Dressler describes a liquid
BACKGROUND OF THE INVENTION rocket engine with a variable thrust injector and an ED nozzle
to improve performance. In the Dressler system, a throat is
0002 Modern rocket propulsion systems can be classified formed at one end of a combustion chamber through which
according to the type of energy source: chemical, nuclear, and hot gases escape. A rod runs through the throat, and a deflec
Solar. Chemical rocket propulsion uses the energy from a tor is formed at the end of the rod, downstream of the throat.
high-pressure combustion reaction of propellant chemicals, A nozzle exit cone extends from the throat. Thus, exiting
which heats reaction product gases to very high temperatures. gases pass through the throat and are deflected by the deflec
These gases are then expanded in a nozzle and accelerated to tor. The deflected gases then pass along the walls of the nozzle
very high Velocities, which, in turn, bring rockets to high exit cone, which direct them in a direction opposite the tra
Velocities in an opposite direction. Nuclear propulsion, using jectory of the rocket.
a fission reactor, a fusion reactor, or directed radioactive 0008 Liquid rocket engines employing efficient variable
isotope decay, has been investigated but remains largely propellant flow into the combustion chamber have been used
undeveloped. Solar propulsion may use Solar panels to heat a effectively for many years but have suffered from perfor
gas. The expanded gas can be expelled through an exhaust mance inefficiencies inherent in the use of cone or bell
noZZle, as with chemical propulsion. nozzles over the wide NPR range which results from the
0003 Chemical propulsion techniques are typically variable chamber pressure and resulting variable thrust when
divided among those using liquid propellants and those using using a rocket engine with a fixed throat area.
Solid propellants. Gaseous propellants and hybrid propellant 0009 While systems such as the above have improved
systems also exist. Typically, liquid propellant rocket engines liquid engine rocketry, no liquid rocket engine design has
feed a propellant under pressure from tanks into a combustion adequately leveraged improved techniques to provide a
chamber. Solid propellant engines, in contrast, store a propel simple and powerful engine with both high efficiency over a
lant "grain' in the combustion chamber, the exposed Surface wide range of backpressures and easily controlled thrust.
of which bums smoothly at a predetermined rate. Combustion Such an efficient and versatile rocket engine would provide
chamber conditions therefore vary with propellant type. The significant gains in many rocketry applications.
techniques applied to control thrust of the various types of 0010 Techniques such as those described above are less
rocket engines historically vary to accommodate for the dif developed in the field of solid propellant rocket motors.
ferent mechanics of liquid versus solid propellants. Methods Designs for use in future generation Army tactical missiles
for optimizing nozzle efficiency are more developed in the have been investigated and tested, as reported in Burroughs,
field of liquid propellant engines than in Solid propellant Susan L. et al., “Pintle Motor Challenges for Tactical Mis
motorS. siles', AIAA Paper 2000-3310, July 2000. These designs use
0004 Methods for initiating and stopping liquid propel a pintle that extends into the throat or just upstream of the
lant rocket engines and for varying the thrust of these liquid throat of a conical expansion nozzle. The pintle is attached to
engines during operation and flight are described in U.S. Pat. a control system that can move the pintle forwards and back
No. 3,897,008; granted Jul. 29, 1975, to Donald G. Nyberg wards within the combustion chamber, thereby varying the
and Ronald F. Dettling entitled “Liquid Fuel Injector System” throat area. The size of the throat area is related to chamber
which is hereby incorporated by reference in its entirety. pressure and thrust of the Solid rocket motor. After passing
0005 Systems providing improved efficiency for liquid through the variable throat area, the exhaust gases are
rocket engines using expansion-deflection (ED) nozzles and expanded in a conventional nozzle (e.g. conical, bell, Rao,
plug nozzles are described in Huzel, Dieter K. and Huang, etc) to produce thrust against the walls of the nozzle. The NPR
David H. Design of Liquid Propellant Rocket Engines. is commonly used to characterize the conditions under which
Washington D.C.: NASA Science and Technical information a rocket operates is the ratio of internal chamber pressure to
Office, 1967, pp. 89-95. The plug nozzle replaces a traditional external (ambient) pressure against which the rocket
noZZle exit cone with a spike centerbody. Exiting gases pass exhausts.
through a throat, and then travel down the surface of the spike 0011 Conventional rocket nozzles must be designed to
to converge in a direction opposite that of rocket trajectory. optimize nozzle efficiency at a given NPR. Nozzle perfor
0006. The use of an ED nozzle is elaborated in Sutton, mance (i.e. the efficiency with which a nozzle converts ther
George P.; Rocket Propulsion Elements, 6th Edition, John mal energy of the heated gases in the chamber into thrust
Wiley and Sons (1992). As stated therein, “this behavior is producing, directed kinetic energy of the exhausted gases)
desirable at low altitudes because the atmospheric pressure is typically degrades at NPR's other than the “design.” or opti
high and may be greater than the pressure of the exhaust mal, pressure ratio.
gases. When this occurs, the exhaust is forced inward and no 0012. As an example, consider a rocket with a constant
longer exerts force on the nozzle walls, so thrust is decreased chamber pressure, a fixed throat area and a conical nozzle
US 2009/0211225 A1 Aug. 27, 2009

which is used to launch a payload through the earth's atmo plug in a plug nozzle configuration, while a second configu
sphere. As the rocket ascends, the ambient pressure into ration provides a moveable plug in an expansion-deflection
which the motor exhausts (atmospheric pressure) will (ED) configuration. The plug and spike operate to achieve
decrease, thus increasing the NPR. Nozzle efficiencies at greatly improved efficiency over a wide range of NPRs. Other
NPR's other than the design ratio will be lower than optimal, advantages and features of the invention are described below.
So rocket designers must choose the pressure ratio “design
point to give the best average performance over the range of BRIEF DESCRIPTION OF THE DRAWINGS
expected NPRs.
0013. A class of nozzles called “plug nozzles or “aero 0018. The systems and methods for improved thrust effi
spikes, with a fixed-position centerbody, or spike, that ciency and control in accordance with the present invention
extends downstream of the combustion chamber throat, have are further described with reference to the accompanying
the characteristic that nozzle efficiency remains relatively drawings in which:
high as a rocket motor with a constant chamber pressure 0019 FIG. 1 illustrates an exemplary embodiment of a
moves through varying ambient pressure conditions. These rocket motor 1 with a moveable plug 4 in a plug nozzle
noZZles are therefore known as “altitude compensating configuration. The position of a plug 4 can be modified with
noZZles. respect to a combustion chamber exit 5. A combustion cham
0014 Nozzles with moveable pintles affect NPR in a dif ber 3 is illustrated, and an exit 5 is formed at a nozzle end of
ferent way, but suffer nonetheless from loss of nozzle effi the combustion chamber 3. The position of the plug 4 is
ciency at “off-design” NPRs. In this class of nozzles, the modifiable by the adaptive control system 9.
pintle is used to vary the throat area, and thus the thrust of 0020 FIG. 2 illustrates a variation of the rocket motor
Solid propellant motors. In varying throat area, these nozzles introduced in FIG.1. The moveable plug 4 is truncated so that
vary the chamber pressure, and thus the propellant burn rate, it is flattened rather than spiked at the downstream end.
with the ultimate effect of varying thrust. However, because 0021 FIG. 3 illustrates an exemplary embodiment of a
the pintle is used in combination with a cone nozzle, varying rocket motor 301 with a moveable plug 304 in an ED nozzle
NPRs force rockets of such a design to operate at sub-optimal configuration. A moveable plug 304 is positioned substan
NPRs. Thus thrust control, or “throttling' is achieved at the tially downstream of the exit 305. A nozzle cone 310 is added.
cost of nozzle efficiency. 0022 FIG. 4 illustrates an exemplary liquid propellant
0015 Thus, theory and test results demonstrate that the rocket engine equipped to adjust a relative position of a plug
tested designs cannot maintain high performance over a wide 404 and an exit 405.
range of NPRs. This is largely because such designs suffer (0023 FIG. 5 illustrates a variety of methods that may be
from efficiency losses due to expansion problems in a fixed employed to optimize NPR by adjusting a relative position of
noZZle exit cone or bell nozzle configuration. Regardless of a plug and an exit.
whether the change in NPR occurs because of decreasing 0024 FIG. 6 illustrates variation of thrust coefficient as a
exhaust pressure (increasing altitude) or decreasing chamber function of nozzle area ratio.
pressure (thrust throttling), nozzle efficiency suffers due to
non-optimal nozzle expansion at off-design NPRs. Perfor DETAILED DESCRIPTION OF ILLUSTRATIVE
mance losses of up to 30% off of optimal efficiency can occur EMBODIMENTS
at off-nominal NPRs. To date, no method has been identified 0025 Certain specific details are set forth in the following
for maintaining near-optimal nozzle efficiency while varying description and figures to provide a thorough understanding
thrust over a wide range. of various embodiments of the invention. Certain well-known
0016. In summary, both liquid and solid rocket motor details often associated with the design and manufacture of
designs have failed to realize their full potential in providing rocket motors are not set forth in the following disclosure,
both high efficiency over a wide range of NPRs, and thrust however, to avoid unnecessarily obscuring the various
control. Such an efficient and versatile solid, liquid, or other embodiments of the invention. Further, those of ordinary skill
propellant type rocket would provide significant gains for in the relevant art will understand that they can practice other
rockets used for commercial and military spacecraft embodiments of the invention without one or more of the
launches, as well as missile launches used for both conven details described below. Finally, while various methods may
tional and anti-terrorism warfare.
be described with reference to steps and sequences in the
SUMMARY OF THE INVENTION following disclosure, the description as Such is for providing
a clear implementation of embodiments of the invention, and
0017. In consideration of the above-identified aspects of the steps and sequences of steps should not be taken as
rocket design, the present invention provides systems and required to practice this invention.
methods for varying the thrust of a rocket while maintaining 0026 FIG. 1 demonstrates a cross-sectional view of an
significantly higher nozzle efficiency over the thrust range. A exemplary rocket motor 1 that employs various features for
moveable plug design is provided for use in rocket motors and improved thrust control and efficiency. The exemplary motor
engines. The plug may be part of a “moveable plug nozzle, 1 depicted in FIG. 1 has features of a solid-propellant rocket
where a combustion chamber exit, such as a cowl, and plug motor. Namely, the motor 1 has a solid propellant grain 2
are moveable with respect to one another. A plug or combus depicted by the left-to-right diagonal shading. While a solid
tion chamber exit may be attached, or otherwise operably propellant rocket motor is used herein for illustration pur
coupled, to an adaptive control system for modifying their poses, it will be recognized that many aspects of the invention
position with respect to one another. The adaptive control are applicable to liquid engines, Solar engines, or indeed any
system may thus control the thrust force and thrust direction rocket engine or motor that makes use of a combustion cham
ofa rocket. At least two configurations employing a moveable ber 3 and throat 5 arrangement to provide thrust. To empha
plug are described: a first configuration provides a moveable size the wide applicability of the invention, the traditional
US 2009/0211225 A1 Aug. 27, 2009

nomenclature that refers to liquid propellant rocket thrust specific propellant types and operational requirements, how
providers as “engines', while referring to the thrust provider ever Such shape can remain generally similar to the shape
of solid propellant rockets as a “motor” is dispensed with illustrated in the figures.
here. Hereafter, the terms “engine' and “motor will be used 0033. In FIG. 2, the numeral 206 refers to the truncated
interchangeably to refer to all types of rocket thrusters using elongated downstream portion of plug 204, while the numeral
all types of propellants. 210 refers to the “missing spike tip that is present in FIG.1.
0027. The solid propellant grain 2 and its burning surface A moveable plug may also have shape contours Suited to an
are contained within a combustion chamber 3. When ignited, ED configuration as illustrated in FIG. 3. In FIG. 3, the
the Surface of the grain 2 burns, providing hot gases from the numeral 306 refers to the elongated downstream portion of
burning surface. The burn rate of the propellant 2 affects the plug 304, and the plug 304 also has a tapered front portion for
flow rate of gas through the throat formed by the exit 5. A the purpose of changing throat area as plug 304 moves with
faster burn rate will force more gases through the throat. The respect to exit 305.
burn rate is dependent on the pressure in the combustion 0034. The plug 4 may be manufactured as a single piece
chamber 3. At higher pressures, the propellant 2 burns faster. with rod 8, or may be separately fabricated and attached to rod
0028. The chamber pressure, in turn, is dependent on the 8. In embodiments without a rod 8, other solutions may be
nozzle throat area. Nozzle throat area is defined as the small adapted to fit the needs of the particular configuration. The
est space through which exhaust gases must pass to exit the plug 4 may be made from the same material as rod 8 or from
combustion chamber 3. In the embodiment of FIG. 1, nozzle Some other material; a sturdy heat-resistant material best
throat area is the smallest annular space between the exit 5 Suited to the propellant and mission is desirable.
and plug 4. The term “throat plane' refers to the plane that
passes through the throat. Note that the position of the throat 0035. The term “exit as used herein refers to the sidewall
plane may change, in some embodiments, when the position Substantially overlapping and adjacent to the throat. A portion
of the plug 4 is modified with respect to the exit 5. More of an exit 5 may form an outer boundary of a throat. In some
importantly, the throat area will change in size and/or shape embodiments where the exit 5 is a very thin piece, the exit 5
when the plug 4 position, or the exit 5 position is modified. may form the throat without any mass upstream or down
This causes a decrease or increase in combustion chamber 3 stream thereof. The exemplary exit 5 in FIG. 1 may form a
pressure, which causes a decrease or increase in burn rate of round throat opening for exhaust gasses to pass through, but
the grain 2, which forces less or more gas through the throat may also form a throat of any other shape. For example, exits
105 and thereby decreases or increases thrust. that form rectangular throats are known in the art and may be
0029. The thrust of the rocket motor 1 is based upon the used. Similarly, exits may be a variety of sizes and may be
specific impulse of the given propellant 2, the chamber 3 manufactured from a variety of materials.
pressure, the area of the throat and the thrust coefficient. The 0036 Referring back to FIG. 1, the movement of a plug 4
thrust coefficient is the measure of efficiency of the expansion and/or an exit 5 may be controlled via a range of mechanisms.
of the exhaust gases and the transfer of their energy to the In the illustrated embodiment, the plug 4 position is con
rocket 1, i.e. the efficiency of the nozzle. trolled by a moveable rod 8. The rod 8 positions the plug 4
0030 The thrust coefficient may change when the rocket within the exit 5. A spike portion 6 may be located down
motor 1 is operating in different ambient pressures. Differing stream of the exit 5. "Downstream, as the term is used here,
ambient pressures will effect the ratio of the pressure inside refers to the stream of exhaust gasses when a rocket engine is
the chamber 3 to the pressure outside the chamber (the NPR), in operation. By “in operation, it is understood that the rocket
which affects the dynamics of the gas flow exiting the rocket propellant 2 is burning.
motor 1. For example, when the rocket motor 1 operates at 0037 Note that rod 8 may be a single straight shaft of any
higher altitudes, the atmospheric pressure decreases, chang suitable material, as illustrated in FIG. 1. Rod 8 may also be
ing the NPR and the corresponding thrust coefficient. Con configured in some other fashion employing curvature or
versely when the chamber pressure is decreased by increasing multiple converging shafts. The rod 8 is one example of a
the nozzle throat area and thus decreasing the propellant burn means for controlling the position of the plug 4, or a portion
rate, the NPR will decrease, thus affecting the nozzle thrust of Such a means, which may be replaced in various embodi
coefficient. The thrust coefficient can be controlled, and ments with other means for controlling plug 4 position. Some
maintained at higher levels if desired, in both the high and low embodiments may employ electromagnetic Suspension and
backpressure situations by using a moveable plug in a plug control mechanisms, flexible disk diaphragms capable of sus
nozzle configuration as illustrated in FIGS. 1 and 2 or in an pending plug 4, flexible meshes, or other means. In embodi
ED nozzle configuration as shown in FIG. 3. ments where the plug 4 remains in a fixed position with
0031. In rocket motor designs contemplated by various respect to engine 1, while the position of the exit 5 is move
embodiments of the invention, a plug 4 is moveable with able with respect to the plug 4, additional techniques may be
respect to a combustion chamber exit 5. The relative change in available for holding the plug 4 in place and modifying the
position can be achieved either by moving the plug 4, or by position of the exit 5.
moving the exit 5, or both. 0038. In FIG. 1, as the control system 9 and rod8 move the
0032. The plug 4 is defined herein as a shaped object plug 4 upstream, the annular restricted throat area is
roughly in the shape illustrated in FIGS. 1, 2, 3, and 4. This increased. This results in decreased chamber 3 pressure and
shape may be referred to as a “pregnant plum' shape. The corresponding decrease in thrust. In the case of a solid pro
plug 4 may comprise an elongated downstream portion 6. pellant, the decreased chamber 3 pressure results in a
which may come to a point, as illustrated in FIG. 1, or may be decreased burn rate of the propellant 2 according to the
truncated as shown in FIG. 2. The exact shape of plug 4 and empirical relation:
spike 6 portion of plug 4 is determined by and optimized for
US 2009/0211225 A1 Aug. 27, 2009

where r is the burn rate at the surface of the propellant, Pc is remainder of rocket 1 via a flexible apparatus, and by con
chamber pressure, and nanda are constants related to specific trolling the motion of the exit 5 using the position control
characteristics of the propellant selected. A decreased burn system 9.
ing rate results in a lower flow rate of propellant and a result 0044 Support brace 7 may be included in various embodi
ing lower thrust. Naturally, reversing the direction of the ments to Support the appropriate position of the moveable
control system and the movement of the plug increases cham plug 4 with respect to the exit 5. In embodiments where the
ber pressure and corresponding thrust. moveable plug 4 can only move axially, Support brace 7 can fit
0039 Modifying the plug 4 and/or exit 5 position around the rod 8 snugly, but not so tight as to prevent axial
upstream and downstream thus controls the amount of thrust sliding. In embodiments where the plug 4 can move radially
of the rocket engine 1, which as a practical matter affects as well as forward and backward, support brace 7 may be
rocket acceleration and Velocity. Upstream and downstream outfitted with additional apparatus to support the rod 8 in the
desired position. Such additional apparatus may be indepen
position modification of the plug 4 and exit 5 with respect to dent of the position control system 9 or may be operably
one another is referred to hereinas axial motion. Thus if either coupled to 9 to act in concert with the positioning activities of
plug 4 or exit 5 is moved directly upstream or directly down adaptive control system 9.
stream, the movement is considered axial. In contrast, mov 0045. The shape of spike 6 will effect the dynamics of
ing the plug 4 or exit 5 from side-to-side affects direction of exhaust gases and so is a feature for close consideration in
thrust, which correspondingly affects the rocket direction. practicing the invention. In particular, the plug 4 and spike
Such movement will be referred to herein as radial move portion 6 thereof may vary depending on whether a plug
ment. Thus, modification of the axial and radial position of nozzle, truncated plug nozzle, or ED nozzle type is used. A
plug 4 and/or exit 5 can be used to alter both rocket speed and plug nozzle configuration is illustrated in FIG. 1, a truncated
direction. Accordingly, position control system 9 and rod 8 plug nozzle is illustrated in FIG. 2, and an ED nozzle is
may comprise apparatus for moving the plug 4 and/or exit 5 illustrated in FIG. 3. Note that despite the different opera
both axially and radially. tional mechanics of the configuration illustrated in FIGS. 1, 2,
0040 Position changes of the plug 4 may be accomplished and 3, each comprises a plug and exit with modifiable relative
via a position control system 9. The position control system 9 position.
is depicted upstream of the combustion chamber 3 in FIG. 1, 0046 Referring to FIG. 2, when the spike 206 portion of
however various embodiments may place it downstream, to plug 204 is truncated, as indicated by the “missing spike
one side, or in some other location with respect to the com point 210, the function of the spike 206 may be approximated
bustion chamber 3. by fluid-mechanical behavior of the propellant downstream
of the truncated plug 204. Truncation has been used in various
0041. Examples of position control systems such as 9 are fixed-plug nozzle designs, and results in what is known as an
presently in use in connection with rockets that use a pintle to aeroSpike. AeroSpike configurations may work well in the
modify rocket thrust. Any presently used or future developed context of moveable plugs provided herein. The advantage of
position control system 9 is considered appropriate for use in an aerospike is that much of the effect of a pointed spike. Such
connection with practicing the invention. as that illustrated in FIG. 1, may be achieved without the
0042. The function of the position control system.9 may be additional mass of the spike tip 210.
simply to adapt to ambient pressures to provide a predictable 0047. The plug nozzle configuration will maintain nozzle
rocket speed, or may be more Sophisticated. Sophisticated efficiency at low flow rates and/or low altitudes where rela
systems might make use of computerized controls that are tively high back pressure causes boundary layer separation
capable of communicating with a computer operated by a and attendant thrust loss in conventional cone and bell
human or automated response system. In Such configurations, nozzles. By contrast, at a low chamber pressure, low thrust
a human might remotely control the trajectory of a rocket by condition using the standard upstream pintle design, the
sending signals to 9, which in turn modifies the position of the exhaust gases do not expand fully into the nozzle but form a
plug 4 to carry out the human instructions. An automated core in the center of the nozzle. With a moveable plug, how
network could also perform the task of the human. Many ever, the plume does not suffer from efficiency-reducing
scenarios might be constructed in which the benefits of such boundary-layer separation at low chamber pressure (low
a system are evident. One such scenario might involve the NPR), and thus the efficiency of the nozzle can be near
automatic adjustment the position of plug 4 with respect to optimized at these reduced-flow conditions.
exit 5 to account for erosion of the plug 4, exit 5, or other 0048 Referring to FIG.3, the use of a moveable plug 304
noZZle Surfaces as the propellant 2 burns, and thus compen is illustrated in the context of an ED nozzle configuration. The
sate for changes in the nozzle throat area and contours during plug 304 is shaped somewhat differently to accommodate the
rocket motor operation. ED nozzle. Note, however, that several important advantages
0043. Note that position control system 9 can modify the accrue from using a moveable plug 4 with elongated down
position of the plug 4 with respect to the exit 5 of the com stream portion 306 in the place of the traditional fixed ED
bustion chamber 3. Note that when plug 4 is moved from a nozzle deflector.
larger diameter portion of the exit 5 to a smaller diameter 0049 FIG. 3 illustrates embodiments of a design variation
portion of the exit 5, either by moving plug 4 downstream or that employs some of the elements of FIG. 1 in a somewhat
exit 5 upstream, the throat area is reduced, and vice versa. different setting. A cross section of an exemplary Solid pro
Changes in throat area may be accomplished by moving the pellant motor 301 is depicted. The motor 301 employs an
plug 4 or by moving the exit 5. Embodiments in which the exit expansion-deflection configuration with a shaped plug 304.
5 is moved while the plug 4 remains fixed with respect to the The shaped plug 304, like the other plugs depicted herein, is
other components of the rocket, Such as sidewall 1 and Sup both moveable with respect to exit 305, and comprises an
port brace 7 can be implemented by mounting the exit 5 to the elongated downstream portion 306, that is numbered sepa
US 2009/0211225 A1 Aug. 27, 2009

rately for the purpose of any specific discussion of that por 0055. The liquid propellant engine 401 may be, for
tion of the plug. In the ED configuration, a tapered rod 312 example, a bipropellant, monopropellant or hybrid type
may be employed upstream of the plug 304. As with the motor engine. Engine 401 may contain a propellant flow control
1 of FIG. 1, the throat area in FIG. 3 can be controlled by device 402, for example a variable flow rate injector which
modifying the position of plug 304 with the tapered rod 312. controls the flow rate of the propellant into the combustion
When such modification results in a decrease in throat area, chamber 403. The propellant flow control device 402 can be
the resulting increased burning rate creates a higher flow rate in an annular design as depicted in FIG. 4 or a multiple orifice
of propellantanda correspondingly higher thrust. The shaped configuration in which upstream metering devices are use to
plug 304 with elongated downstream portion 306 down control the propellant flow rate(s). The propellants ignite and
stream of the exit 305 can serve to maintain higher overall provide hot gases in the combustion chamber 403 which then
nozzle efficiency as the thrust, and therefore NPR, is varied. flow past the moveable plug 404 and exit 405, then out past
the spike 406 to the ambient environment.
0050 Reversing the direction of the control system 309 0056. In FIG. 5, the position control apparatus is illus
and the movement of the plug 304 and tapered rod 312 trated as an actuator 409 system for moving the plug to the
reduces the combustion chamber 303 pressure and the thrust. desired axial position. Actuator 309 can be, in one embodi
The chamber pressure can be reduced to near extinguishment ment, a state of the art mechanism such as those used to
(Smoldering) or to complete extinguishment by including a position a pintle on tactical Solid propellant motors. These
notched area 311 upstream of the tapered rod 312. By posi actuators and the Surrounding insulation are designed to with
tioning the notched area 311 in the exit 305, the throat area stand the intense heat generated in the combustion chamber.
may be increased to a value Sufficient for complete extin Support structure 407 may be a state of the art spider type
guishment of a solid propellant grain. In the illustrated actuator support. The outer walls and insulation 408 shown in
embodiment, the plug 304, tapered rod 312, and notched area FIG. 4 can be optionally replaced with regenerative cooling
311 are controlled via rod 308 and adaptive control system techniques in a flight missile system or with more durable
309, although any other available means may be used to ablative insulation types.
modify the relative position of notch 311, tapered rod 312, 0057 The thrust of the rocket engine 401 is dependent on
and plug 304 with respect to exit 305, as discussed above. the specific impulse of the given propellant system, the cham
0051. In embodiments such as FIG.3, the plug 304 directs ber 403 pressure, the area of the throat through which the
the flow of exhaust products to the outer walls of the nozzle gases exit, and the thrust coefficient. The thrust coefficient is
310 even at low NPR. The hot gases are expanded to the a measure of the efficiency of the expansion of the exhaust
ambient atmosphere around the plug 304 which directs the gases and transfer of their energy to the rocket.
flow to the outer nozzle walls 10 in a cone or bell exhaust 0.058 As the propellant flow rate is decreased employing
nozzle 310 to provide thrust to the rocket. The elongated the propellant flow control device 402, the moveable plug 404
downstream portion, or spike 306, extends downstream of can be intelligently positioned by an automated control com
plug 304 to provide efficient exhaust dynamics at NPR ranges ponent 412 that is communicatively coupled to the actuator
that cause gases to cling to the walls of the spike 306. Thus the 409. Automated control component 412 moves the plug 404
spike 306 and the expansion-deflection arrangement comple Such that the throat area is reduced to a value that just main
ment each other to the extent that they affect exhaust dynam tains the chamber 403 pressure at a substantially constant
ics in overlapping ranges. The spike 306 and the EDarrange level. This higher chamber pressure results in a higher NPR
ment extend the thrust efficiency to the extent that they do not than if the chamber pressure were reduced, as in a conven
affect overlapping NPR ranges. tional variable thrust engine. The higher NPR then provides a
0052 A control system 309 can provide the correct posi higher thrust coefficient, increased efficiency and higher spe
tioning of the moveable plug 304, tapered rod 312, and/or cific impulse, and resulting higher performance of the rocket
notch 311 in the exit 5 to produce the desired thrust. As with engine 401.
the position control system 109 from FIG. 1, system 309 may 0059 Automated control component 412 can intelligently
be upstream or downstream of the combustion chamber 303, position the plug, for example, if it is communicatively
may be similar to presently-used systems to control pintles in coupled to one or more of the propellant flow control device
Solid-propellant rocketry (or some future developed position 402 (or Some other means of measuring propellant flow), a
control technology), may be preconfigured to react predict chamber pressure gauge 410, and an altimeter or barometer
ably to atmospheric conditions or remotely controllable, and 411.
may operate to modify the position of the plug 304 or the 0060. In one embodiment, automated control component
position of the exit 305, or both. 412 can position the moveable plug 404 by linking movement
0053. The remaining elements, e.g. sidewall 301, grain of the plug 404 to rate of propellant flow established by the
302, and support brace 307 are generally analogous to the propellant flow control device 402. Thus, when flow rate is to
corresponding elements from FIG. 1. Please refer to the dis be reduced, for example, the automated control component
cussion of those elements above for a description of the 412 can also instruct the actuator 409 to move the plug 404
function of various embodiments of these features of a rocket downstream so as to maintain optimum performance NPRs.
motor incorporating aspects of the invention. 0061. In another embodiment, automated control compo
0054 FIG. 4 illustrates one embodiment of a moveable nent 412 can position the moveable plug 404 based on mea
plug nozzle in a variable thrust liquid propellant engine. An Surements of a chamber pressure gauge 410 and/or an altim
embodiment Such as FIG. 4 may employ the concepts dis eter or ambient pressure gauge 411. For example, if there is an
cussed above, and may also link propellant flow rate to plug increase in altitude or drop in pressure measured by 411, the
position as discussed below. Alternative embodiments can be plug 404 may be moved upstream to correspondingly
envisioned comprising a moveable expansion-deflection decrease the chamber 403 pressure, thereby maintaining a
device in a conical or bell shaped nozzle. constant NPR. Similarly, if there is an increase in chamber
US 2009/0211225 A1 Aug. 27, 2009

pressure measured by 410, the plug 404 might be advanta highly dependent on the NPR which is the chamber pressure
geously moved upstream to maintain a constant NPR. divided by the exhaust pressure or: NPR=Pc/Pe. The opti
0062. It will be appreciated that the various components mum performance of a rocket engine is thus dependent on the
402, 410, 411, and 412 can be combined in a variety of NPR and thrust coefficient. FIG. 6 illustrates this relationship.
combinations and integrated into existing control systems in FIG. 6 illustrates variation of the thrust coefficient as a func
a wide variety of configurations. Also, as an alternative to tion of nozzle area ratio, Ae/At, and NPR, Pc/Pa.
measuring pressures or altitude, predicted values can be used 0072 For a conventional cone or bell nozzle rocket
based on a known rocket starting position and trajectory. thruster with a variable propellant flow rate to vary the thrust
0063 A novel method may thus be performed utilizing a and a fixed area throat, the C, and therefore the performance
system such as FIG. 4, wherein the thrust of a liquid propel degrades as the chamber pressure moves away from the
lant rocket engine is varied to maintain optimum performance design NPR. This is illustrated in FIG. 6 as the curves at a
over the entire thrust range of a rocket. Various approaches to given Pc/Pe move away on either side of the optimum line.
such a method are illustrated in FIG. 5. As illustrated in FIG. 0073 Part of this degradation in performance has been
5, a moveable plug nozzle configuration is employed to vary corrected with the use of a stationary plug nozzle or expan
throat area in response to changes in NPR and/or changes in sion deflection nozzle with a fixed throat area, as described in
flow rate of the propellant into the chamber. By intelligently U.S. Pat. No. 6,591,603. This results in the C, and thus the
carrying out the methods of FIG. 5, an NPR can be maintained performance following the optimum curve shown in FIG. 6 as
as close as possible to the design NPR for the rocket engine. the chamber pressure and thrust are decreased. Although this
0064. For example, in exemplary method 500, when pro scenario results in improved performance compared to the
pellant flow rate is decreased 501, a moveable plug may be bell nozzle with a fixed throat area, the C, and the specific
moved so as to decrease throat area 502, thereby maintaining impulse are both reduced significantly as the flow rate and
a substantially constant NPR. In exemplary method 510, chamber pressure are reduced.
when a decrease in chamber pressure is sensed or predicted 0074 The performance degradation resulting from both of
511, a moveable plug may be moved so as to decrease throat the above factors can be reduced using a moveable plug
area 512, thereby maintaining a substantially constant NPR. nozzle to vary the throat area Such that a constant high cham
In exemplary method 520, when an increase in ambient pres ber pressure and a constant high optimum NPR are achieved
sure is sensed or predicted 521, a moveable plug may be over the full range of thrust variation. This constant high
moved so as to decrease throat area 522, thereby maintaining chamber pressure results in a constant high C, and a resulting
a substantially constant NPR. optimum high Specific Impulse (ISP) for any thrust level.
0065. In exemplary method 550, when propellant flow rate (0075. The methods of FIG. 5 allow a liquid rocket engine
is increased 551, a moveable plug may be moved so as to to operate at a constant oxidizer to fuel (O/F) ratio over the
increase throat area 552, thereby maintaining a substantially entire variable thrust range. Current variable thrust liquid
constant NPR. In exemplary method 560, when an increase in engines control the flow rate of the two propellants using
chamber pressure is sensed or predicted 561, a moveable plug either variable area orifices or variable area annuluses for
may be moved so as to increase throat area 562, thereby each of the two propellants. For each of the two propellants in
maintaining a Substantially constant NPR. In exemplary any given system, the flow rate of the propellant through the
method 570, when a decrease in ambient pressure is sensed or metering device is dependant in a simplified form on a con
predicted 571, a moveable plug may be moved so as to stant times two variables-the discharge coefficient and the
increase throat area 572, thereby maintaining a substantially square root of the pressure drop. It is very difficult to design
constant NPR.
the metering device in such a way that the oxidizer and fuel
0066. Of course, any of the factors referred to in 501, 511, flow rates remain at a constant ratio over the full range of
521,551,561, and 571 may offset or augment certain other of pressure drops resulting from the change in chamber pressure
Such factors, thereby increasing or decreasing the extent to over the thrust range. Utilizing the methods of FIG. 5, cham
which the throat area is increased or decreased.
0067. In a liquid engine, the chamber pressure (Pc) in the ber pressure can be substantially constant, so the oxidizer to
rocket engine is equal to the propellant flow rate, od, times the fuel ratio remains at the optimum design point over the entire
propellant c (measure of propellant performance) divided by variable thrust operating range of the rocket engine.
At divided by the gravity constant, g. (32.174 ft/sec) or 0076 A second performance benefit is the maintenance of
basically: the oxidizer to fuel ratio over the entire thrust variation range.
As mentioned previously, most variable thrust liquid bipro
PC-constantxeo/At pellant systems have a variation in the oxidizer to fuel ratio
0068. The above equation illustrates that with a constant because of the variation in pressure drop from the tanks,
throat area, the chamber pressure decreases with a decrease in through the metering devices, and into the combustion cham
the flow rate of propellant. ber throughout the thrust variation range when the chamber
0069. The thrust (F) of a rocket engine is equal to the pressure is also varying. This oxidizer to fuel ratio shift is
Throat Area (At) times the chamber pressure (Pc) times the different for the various bipropellant systems based on their
thrust coefficient (C) or propellant densities and the discharge coefficients associated
with the metering devices.
0077 Although exemplary embodiments refer to utilizing
0070 The above equation illustrates that the thrust (F) is the present invention in the context of solid-propellant rocket
decreased not only in proportion to the decrease in chamber motors, the invention is not so limited, but rather may be
pressure but also in proportion to the decrease in the thrust implemented in connection with any rocket motor configura
coefficient. tion in which thermal energy is converted to directed kinetic
(0071) The thrust coefficient, C is based on the efficiency energy, and thus thrust, by means of a nozzle. Therefore, the
of the expansion of the exhaust products in the nozzle and present invention should not be limited to any single embodi
US 2009/0211225 A1 Aug. 27, 2009

ment, but rather should be construed in breadth and scope in 17. The rocket engine of claim 9 wherein said plug is
accordance with the appended claims. positioned within said exit in an expansion-deflection (ED)
We claim: nozzle configuration.
1. A liquid propellant rocket engine with a combustion 18. The rocket engine of claim 9, further comprising a
chamber configured such that a propellant will flow out of the position control apparatus for modifying said relative posi
combustion chamber in a downstream direction, said rocket tion.
engine comprising: 19. A method for optimizing thrust in a liquid propellant
a propellant flow control device for adjusting a rate of rocket engine, comprising:
propellant flow into said combustion chamber; adjusting a rate of propellant flow into a combustion cham
an exit formed at a downstream end of said rocket engine; ber;
a plug with an elongated downstream portion, wherein a adjusting a relative position of a plug with an elongated
relative position of the plug and the exit is axially modi downstream portion and an exit formed at a downstream
fiable during operation of said rocket engine; end of said rocket engine to maintain a Substantially
an automated control component that adjusts said relative constant Nozzle Pressure Ratio (NPR).
position to account for an adjustment of said rate of 20. The method for optimizing thrust in a liquid propellant
propellant flow. rocket engine of claim 19, further comprising sensing or
2. The rocket engine of claim 1 wherein said automated predicting a change in ambient pressure Surrounding said
control component adjusts said relative position to maintain a rocket engine, and accounting for said change in ambient
substantially constant Nozzle Pressure Ratio (NPR). pressure when adjusting said relative position.
3. The rocket engine of claim 1 wherein said automated 21. The method for optimizing thrust in a liquid propellant
control component adjusts said relative position to account rocket engine of claim 19, further comprising sensing or
for a change in ambient pressure. predicting a change in altitude of said rocket engine, and
4. The rocket engine of claim 1 wherein said plug is posi accounting for a corresponding change in ambient pressure
tioned within said exit in a plug nozzle configuration. when adjusting said relative position.
5. The rocket engine of claim 1 wherein said elongated 22. The method for optimizing thrust in a liquid propellant
downstream portion converges to form a spike. rocket engine of claim 19, wherein said adjusting a rate of
6. The rocket engine of claim 1 wherein said elongated propellant flow comprises decreasing said rate of propellant
downstream portion is truncated. flow, and wherein said adjusting a relative position comprises
7. The rocket engine of claim 1 wherein said plug is posi moving said plug closer to said exit.
tioned within said exit in an expansion-deflection (ED) nozzle 23. The method for optimizing thrust in a liquid propellant
configuration. rocket engine of claim 19, wherein said adjusting a rate of
8. The rocket engine of claim 1, further comprising a posi propellant flow comprises increasing said rate of propellant
tion control apparatus for modifying said relative position. flow, and wherein said adjusting a relative position comprises
9. A liquid propellant rocket engine with a combustion moving said plug away from said exit.
chamber configured such that a propellant will flow out of the 24. A method for optimizing thrust in a rocket motor,
combustion chamber in a downstream direction, said rocket comprising:
engine comprising:
a propellant flow control device for adjusting a rate of sensing or predicting a change in ambient pressure Sur
propellant flow into said combustion chamber; rounding said rocket motor;
an exit formed at a downstream end of said rocket engine; adjusting a relative position of a plug with an elongated
a plug with an elongated downstream portion, wherein a downstream portion and an exit formed at a downstream
relative position of the plug and the exit is axially modi end of said rocket motor to maintain a Substantially
fiable during operation of said rocket engine; constant Nozzle Pressure Ratio (NPR).
an automated control component that adjusts said relative 25. The method for optimizing thrust in a rocket motor of
position to account for a change in ambient pressure. claim 24, wherein said rocket motor is a solid propellant
10. The rocket engine of claim 9 wherein said automated rocket motor.
control component adjusts said relative position to maintain a 26. The method for optimizing thrust in a rocket motor of
substantially constant Nozzle Pressure Ratio (NPR). claim 24, wherein said rocket motor is a liquid propellant
11. The rocket engine of claim 9 wherein said automated rocket engine.
control component is communicatively coupled to an ambi 27. The method for optimizing thrust in a liquid propellant
ent pressure barometer. rocket engine of claim 24, wherein said sensing or predicting
12. The rocket engine of claim 9 wherein said automated a change in ambient pressure comprises sensing or predicting
control component is communicatively coupled to a chamber a decreasing ambient pressure, and wherein said adjusting a
pressure barometer. relative position comprises moving said plug away from said
13. The rocket engine of claim 9 wherein said automated exit.
control component is communicatively coupled to an altim 28. The method for optimizing thrust in a liquid propellant
eter. rocket engine of claim 24, wherein said sensing or predicting
14. The rocket engine of claim 9 wherein said plug is a change in ambient pressure comprises sensing or predicting
positioned within said exit in a plug nozzle configuration. an increasing ambient pressure, and wherein said adjusting a
15. The rocket engine of claim 9 wherein said elongated relative position comprises moving said plug closer to said
downstream portion converges to form a spike. exit.
16. The rocket engine of claim 9 wherein said elongated
downstream portion is truncated.