Patent Description:
In rocket propulsion systems having a traditional bell rocket nozzle, optimized thrust performance can be achieved only at an altitude where the pressure of the airflow exiting the nozzle equals the ambient pressure. If the ambient pressure is greater than the pressure of the airflow exiting the nozzle, the exhaust plume of the rocket will be under-expanded. Alternatively, if the ambient pressure is less than the pressure of the airflow exiting the nozzle, the exhaust plume of the rocket will be over-expanded. In either of these described instances where the ambient pressure is not equal to the pressure of the airflow exiting the nozzle, thrust efficiency and rocket propulsion performance is lost. Instead, for a rocket having a traditional bell rocket nozzle that flies across various altitudes, thrust performance will only be optimized at one specific altitude.

Rocket propulsion systems having an aerospike rocket nozzle are capable of achieving optimal expansion of an exhaust plume of a rocket under a wide range of altitudes, improving thrust efficiency and rocket propulsion performance across an entire flight of the rocket. Specifically, instead of initially being encompassed by a bell rocket nozzle, the airflow exiting the rocket engine blast tube or combustion chamber flows across the aerospike and creates an exhaust plume that is contained by the ambient atmospheric pressure. However, conventional aerospike nozzles have an annular or linear geometry, neither of which are practical for use in tactical missile designs due to packing, producibility and heating constraints. For example, for an aerospike nozzle having an annular geometry, struts are required to hold the aerospike in a center of the exit opening of the combustion chamber. These struts pose a structural liability as they have the potential to cause blockage to the airflow exiting the combustion chamber, overheat and cause various other problems with the rocket propulsion system.

<CIT> discloses an apparatus including a slotted multi-nozzle grid with a plate having multiple elongated slotlettes through the plate. Each of at least some of the slotlettes has a convergent input, a divergent output, and a narrower throat portion separating the convergent input and the divergent output. At least some of the slotlettes are arranged in multiple rows. The plate further includes multiple cooling channels through the plate. At least some of the cooling channels are located between the rows of slotlettes. Each cooling channel is configured to transport coolant through the plate in order to cool the plate, such as to cool the plate as hot combustion gases pass through the plate. Each of at least some of the rows may include at least two slotlettes, and two adjacent slotlettes in one row may be separated by a structural ligament (which may have a tear-drop cross-sectional shape).

<CIT> discloses an aerial missile having a propelling jet, a jet vane pivotally mounted in the jet orifice, a trailing vane pivotally mounted on said vehicle and exposed to the medium through which the vehicle moves, said trailing vane being arranged to trail rearwardly of said pivot line, and means mechanically connecting said trailing vane to said jet vane so that said trailing vane controls said jet vane in a contra directionto the direction of movement of said trailing vane, whereby whenever said trailing vane is deflected by said medium because the axis of said vehicle does not coincide with its direction of niotion, a corresponding contra movement of said jet vane will occur, said jet vane, in turn, deflecting said jet such that the deflecting jet will tend to return said vehicle to its proper orientation.

In a general embodiment, a rocket engine nozzle manufacturable and applicable to tactical missile designs includes an aerospike having a plurality of airfoil fins distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The airfoil fins and an inner perimeter of the combustion chamber define a plurality of apertures which choke an airflow exiting the combustion chamber, causing the airflow to expand supersonically along the airfoil fins. The aerospike rocket engine nozzle requires less machine precision and achieves packing benefits over conventional bell and aerospike nozzle geometries. The configuration of the aerospike rocket engine nozzle also removes the producibility and heating constraints typically encountered with conventional aerospike nozzles in tactical missile applications while improving thrust performance of the rocket engine across a wide range of altitudes.

According to an aspect of the invention, a rocket engine comprising: a rocket engine combustion chamber, and a rocket engine nozzle including an aerospike including: a plurality of airfoil fins disposed at an exit end of the rocket engine combustion chamber and extending across an exit plane of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis; and a central airfoil hub from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction; wherein the plurality of airfoil fins and an inner perimeter of the rocket engine combustion chamber define a plurality of apertures between adjacent airfoil fins at the exit plane, the plurality of apertures being configured to choke an airflow exiting the rocket engine combustion chamber and cause the airflow to expand supersonically along the plurality of airfoil fins to create thrust.

According to an embodiment of any paragraph(s) of this summary, the rocket engine combustion chamber is cylindrical.

According to another aspect of the invention, a method of operating a rocket propulsion system, the method comprising: providing a rocket engine including a rocket engine combustion chamber and a rocket engine nozzle, the rocket engine nozzle including: an aerospike having a plurality of airfoil fins disposed at an exit end of the rocket engine combustion chamber and extending across an exit plane of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis; a central airfoil hub from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction; wherein the plurality of airfoil fins and an inner perimeter of the rocket engine combustion chamber define a plurality of apertures between adjacent airfoil fins at the exit plane, and operating the rocket engine such that an airflow exits the rocket engine combustion chamber at the exit plane and the plurality of apertures choke the airflow exiting the rocket engine combustion chamber at the exit plane, causing the airflow to expand supersonically along the plurality of airfoil fins to create thrust.

The following description and the annexed drawings set forth in detail certain illustrative embodiments of the invention. These embodiments are indicative, however, of but a few of the various ways in which the principles of the invention may be employed. Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings.

The annexed drawings show various aspects of the invention.

According to a general embodiment, an aerospike rocket nozzle <NUM> manufacturable and applicable to tactile missile systems achieves optimal expansion and thrust generation at a plurality of altitudes and ambient pressures with a smaller amount of hardware, in terms of both mass and volume, compared with conventional bell rocket nozzles. Additionally, the configuration of the aerospike rocket nozzle <NUM> disclosed herein reaps both performance and packaging benefits over conventional aerospike rocket nozzles and, unlike conventional aerospike rocket nozzles, is producible specifically for tactical missile systems. Specifically, for example, the aerospike rocket nozzle <NUM> disclosed herein has geometry having a shorter length and smaller diameter compared to conventional bell and aerospike nozzles, allowing for larger combustion chambers, blast tubes, and warheads.

Referring now to the figures, and initially to <FIG>, an exemplary embodiment is depicted of an aerospike rocket engine nozzle <NUM> manufacturable and applicable for use in a rocket engine <NUM> of (for example) a tactical missile. The rocket engine <NUM> includes a rocket engine combustion chamber <NUM> or rocket motor chamber. The rocket engine nozzle <NUM> is disposed at an exit end <NUM> of the rocket engine combustion chamber <NUM>. The rocket engine nozzle <NUM> includes an aerospike <NUM> having a configuration that achieves improved expansion and exit velocity of an airflow exiting the combustion chamber <NUM>, and having a reduced dimension compared to conventional bell nozzles. Specifically, the aerospike <NUM> includes a plurality of airfoil fins <NUM> disposed at the exit end <NUM> of the combustion chamber <NUM> and extending across an exit plane <NUM> of the combustion chamber <NUM>. As illustrated in <FIG>, the exit plane <NUM> is a plane that spans an exit opening <NUM> at the exit end <NUM> of the combustion chamber <NUM>.

The plurality of airfoil fins <NUM> are distributed around a central longitudinal axis <NUM> at the exit plane <NUM>. For example, in an embodiment in which the combustion chamber is cylindrical, the plurality of adjustable airfoil fins <NUM> may be distributed radially around the central longitudinal axis <NUM>. Although in the illustrated embodiment, the combustion chamber <NUM> is cylindrical and the exit opening <NUM> is circular, the combustion chamber <NUM> and exit opening <NUM> may be of different shapes and sizes, for example polygonal or otherwise non-axisymmetric. In another embodiment, there may be multiple combustion chambers <NUM> wherein each of the plurality of airfoil fins <NUM> are disposed at an exit plane <NUM> of each combustion chamber <NUM>. Stated differently, each airfoil fin <NUM> may be designated to one combustion chamber <NUM>. In any embodiment, the central longitudinal axis <NUM> is an axis that extends along a center line of the rocket motor combustion chamber <NUM> and is perpendicular to the exit plane <NUM> at a center point of the exit opening <NUM>.

The plurality of airfoil fins <NUM> and an inner perimeter <NUM> of the combustion chamber <NUM> at the exit opening <NUM> define a plurality of apertures <NUM> between adjacent airfoil fins <NUM> at the exit plane <NUM>. These apertures <NUM> act as a nozzle throat, which chokes the airflow as it exits the combustion chamber <NUM> at the exit plane <NUM>, causing the airflow to expand supersonically along the plurality of airfoil fins <NUM>. A simplified two-dimensional representation of such supersonic expansion over an airfoil fin <NUM> is depicted in <FIG>. The supersonically expanded airflow imparts a force on the aft surface <NUM> of each of the plurality of airfoil fins <NUM>, generating a thrust that propels the missile in a forward direction. As the plurality of airfoil fins <NUM>, and therefore the plurality of apertures <NUM> acting as the nozzle throat, are disposed at the exit plane <NUM> of the combustion chamber <NUM>, the aerospike rocket nozzle <NUM> is configured for optimal expansion and thrust generation at a wide range of altitudes and ambient pressures.

In the illustrated embodiment the plurality of airfoil fins <NUM> are distributed axisymmetrically around the longitudinal axis <NUM>. In this embodiment, therefore, the apertures <NUM> formed between adjacent airfoil fins <NUM> and the inner perimeter <NUM> of the combustion chamber <NUM> are equal in size. Accordingly, the thrust that is created will propel the missile in a relatively straight forward direction.

The plurality of airfoil fins <NUM> may include two or more distinct airfoil fins <NUM>. For example, in an embodiment the plurality of airfoil fins <NUM> may include three distinct airfoil fins <NUM>. In the illustrated embodiment the plurality of airfoil fins <NUM> include four distinct airfoil fins <NUM>. In the illustrated embodiment, having four airfoil fins <NUM> distributed axisymmetrically around the longitudinal axis, the aerospike <NUM> has a cruciform configuration. In another embodiment, the plurality of airfoil fins <NUM> may include five or more distinct airfoil fins <NUM>.

The plurality of airfoil fins <NUM> are fixed to the exit end <NUM> of the combustion chamber <NUM> such that they do not rotate or translate in any direction. The plurality of airfoil fins <NUM> may be, for example, welded to the inner perimeter <NUM> of the combustion chamber <NUM> at the exit end <NUM>. The plurality of airfoil fins <NUM> are additionally be welded, or otherwise fixed, to each other at the central longitudinal axis, such that the only contact between the aerospike <NUM> and the combustion chamber <NUM> is where each of the plurality of airfoil fins <NUM> are fixed to the inner perimeter <NUM>.

The aerospike <NUM> may include a central airfoil hub <NUM>, to which each of the plurality of airfoil fins <NUM> are fixed and from which each of the plurality of airfoil fins <NUM> extend radially outward. Each of the plurality of airfoil fins <NUM> may be welded, or otherwise fixed, to the central airfoil hub <NUM> at the central longitudinal axis <NUM>. The central airfoil hub <NUM> may be configured such that a maximum length of the central airfoil hub <NUM> is less than or equal to a maximum length of the plurality of airfoil fins <NUM> in the longitudinal direction (the direction in which the longitudinal axis <NUM> extends). In this way, a majority of the airflow exiting the exit end <NUM> of the combustion chamber <NUM> is configured to supersonically expand across the plurality of airfoil fins <NUM> rather than the central airfoil hub <NUM>. Therefore the supersonic expansion created by the plurality of airfoil fins <NUM> generates a majority of the thrust that propels the missile forward.

The aerospike <NUM>, including the plurality of airfoil fins <NUM> and the central airfoil hub <NUM>, may be made of high temperature alloys such as titanium-zirconium-molybdenum (TZM), tungsten, carbon-carbon, or silica-filled ethylene propylene diene monomer (EPDM). The material thickness of the airfoil vanes <NUM> may be dependent on the specific implementation and environment in which they are to be used, such as whether they will be exposed to high temperatures, as this would affect the rate of erosion.

With reference to <FIG>, a method <NUM> of operating a rocket propulsion system is depicted. In step <NUM> the method includes providing a rocket engine. The rocket engine has a rocket engine combustion chamber and a rocket engine nozzle. The rocket engine combustion chamber may be cylindrical. The rocket engine nozzle includes an aerospike having a plurality of airfoil fins disposed at an exit end of a rocket engine combustion chamber and extending across an exit plane of the rocket engine combustion chamber. The plurality of airfoil fins are distributed around a central longitudinal axis, the central longitudinal axis being an axis that extends along a center line of the rocket motor combustion chamber, perpendicular to the exit plane at a center point of an exit opening at the exit plane.

In an embodiment, the aerospike further includes a central airfoil hub from which each of the plurality of airfoil fins extend radially outward. A maximum length of the central airfoil hub in a longitudinal direction (the direction in which the longitudinal axis extends) is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction.

Claim 1:
A rocket engine (<NUM>) comprising:
a rocket engine combustion chamber (<NUM>), and
a rocket engine nozzle (<NUM>) including an aerospike (<NUM>) including:
a plurality of airfoil fins (<NUM>) disposed at an exit end (<NUM>) of the rocket engine combustion chamber and extending across an exit plane (<NUM>) of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis (<NUM>); and
a central airfoil hub (<NUM>) from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction;
wherein the plurality of airfoil fins and an inner perimeter (<NUM>) of the rocket engine combustion chamber define a plurality of apertures (<NUM>) between adjacent airfoil fins at the exit plane, the plurality of apertures being configured to choke an airflow exiting the rocket engine combustion chamber and cause the airflow to expand supersonically along the plurality of airfoil fins to create thrust.