Patent Description:
Aircraft-like reusability for rockets has long been the "holy grail" of rocketry due to the potential for huge cost benefits. The ability to recover and reuse an upper stage rocket of a multi-stage rocket system (e.g., the second stage rocket of a two-stage rocket system) remains a significant technical gap that has not yet been solved by the industry. Reusing the upper stage of a multi-stage rocket is challenging due to the harsh re-entry environment and the performance penalties associated with increased structural mass required for robust reuse. Upper stage rockets are typically constructed with the minimum structure and complexity since any mass addition to the second stage is a <NUM>:<NUM> reduction in payload capacity. Reusing an upper stage rocket therefore requires significant additional functionality but with minimal mass addition.

Rockets and other vehicles that travel at or above hypersonic speeds (e.g., space re-entry vehicles, aircraft, missiles, etc.) require a means to protect themselves from the heating that occurs at such high speeds. Conventional solutions for mitigating such heating include use of one or more of the following: (i) ablative materials, which undergo pyrolysis and generate gases that move downstream in a boundary layer to form a protective film layer; (ii) high-temperature materials (e.g., ceramics, carbon-carbon, etc.); (iii) composite materials, which insulate a base material and radiate heat away therefrom; and (iv) transpiration cooling, which involves use of a thin protective film that is provided by a gas passing through a semi-porous wall.

Existing heat management solutions have cost, operations, and/or mass impacts that may not trade favorably in certain applications, such as reusable vehicles. For example, ablative materials and fragile ceramics are incompatible with a highly reusable system. Transpiration cooling of a heat shield is costly and difficult to control. What is needed is a cooling system that is highly robust, highly controllable, and well suited for long term reusability.

Aspects of the present invention are directed to these and other problems.

<CIT> relates to cooling the airframe of an aircraft having an air intake engine. The heating experienced by such aircraft is distinguished from the heating experienced by atmospheric re-entry vehicles. <CIT> relates to launch vehicles for transporting goods and personnel from the surface of the earth to low earth orbit and to the trajectory or path which the launch vehicle follows in achieving earth orbit. <CIT> discloses a two-stage launch vehicle having an upper stage with a heat shield. The upper stage is intended to be reusable.

According to an aspect of the present invention, an upper stage rocket of a multi-stage rocket system includes an actively-cooled heat shield system. The heat shield system includes a heat shield, a tank, a pump, a heat exchanger, and a turbine. The heat shield defines a windward side of an upper stage rocket. The tank is onboard the upper stage rocket and is configured to store a coolant. The pump is onboard the upper stage rocket and is configured to receive the coolant from the tank and output a pressurized coolant. The heat exchanger is onboard the upper stage rocket and is integrally connected with the heat shield. The heat exchanger is configured to receive the pressurized coolant from the pump, transfer heat from the heat shield to the pressurized coolant to generate a heated fluid, and output the heated fluid. The turbine is onboard the upper stage rocket and includes an inlet, a shaft, and an outlet. The inlet is configured to receive the heated fluid output from the heat exchanger. The shaft is coupled to the pump and includes turbine blades mounted thereon. The shaft is configured to rotate and thereby power the pump when the heated fluid received from the heat exchanger acts on the turbine blades. The outlet is configured to output the heated fluid.

A reusable upper stage rocket of a multi-stage rocket system includes an actively-cooled heat shield system that converts heat from a high Mach number flow environment into energy to drive a liquid coolant pump.

According to another aspect of the present invention, a method for actively cooling a windward side of an upper stage rocket of a multi-stage rocket system during atmospheric re-entry includes the steps of: initiating driving of a pump onboard the upper stage rocket to initiate output of a pressurized coolant from the pump; flowing the pressurized coolant output by the pump through a heat exchanger integrally connected with a heat shield that defines at least a portion of the windward side of the upper stage rocket; transferring heat from the heat shield to the pressurized coolant to generate a heated fluid; inputting the heated fluid to a turbine onboard the upper stage rocket, the turbine including a shaft coupled to the pump and turbine blades mounted to the shaft; and exposing the turbine blades to the heated fluid to drive the shaft and thereby continue driving the pump.

In addition to one or more of the features described above, further aspects of the present invention can include one or more of the following features, individually or in combination:.

These and other aspects of the present invention will become apparent in light of the drawings and detailed description provided below.

Referring to <FIG> and <FIG>, the present disclosure describes an actively-cooled heat shield system <NUM> and a vehicle <NUM> including the same. The heat shield system <NUM> converts heat from a high Mach number flow environment <NUM> into energy to drive a liquid coolant pump <NUM> (see <FIG>).

The invention is directed to an upper stage rocket of a multi-stage rocket system. In the embodiment illustrated in <FIG>, the vehicle <NUM> is a second stage rocket of a two-stage rocket system (not shown). The vehicle <NUM> (hereinafter the "second stage rocket <NUM>") extends along a centerline <NUM> between a forward end <NUM> and an opposing aft end <NUM> thereof. The second stage rocket <NUM> includes a payload <NUM> toward the forward end <NUM>, and an engine <NUM> toward the aft end <NUM>. The aft end <NUM> defines the windward side <NUM> of the second stage rocket <NUM>. In the illustrated embodiment, the engine <NUM> is an augmented aerospike nozzle engine as disclosed in <CIT> by the same inventors, and in the International Patent Application claiming priority to <CIT>. In other embodiments, the engine <NUM> is a bell nozzle engine or another type of rocket engine, or the vehicle may not include an engine at all.

Referring to <FIG>, during use the second stage rocket <NUM> moves through an environment <NUM> (e.g., the atmosphere, space) at freestream Mach numbers <NUM> that can approach Mach thirty (<NUM>) for space reentry vehicles. A bow shock <NUM> is formed upstream of the second stage rocket <NUM>, and temperature on the vehicle side of the bow shock <NUM> can reach thousands of degrees Kelvin. The windward side <NUM> of the second stage rocket <NUM> is exposed to these high temperatures and therefore cooling and/or otherthermal protection is necessary for reusability.

The actively-cooled heat shield system <NUM> includes a heat shield <NUM>, a tank <NUM>, a pump <NUM>, a heat exchanger <NUM>, and a turbine <NUM>.

The heat shield <NUM> defines an outer surface of the windward side <NUM> of the second stage rocket <NUM>.

The tank <NUM> is onboard the second stage rocket <NUM> and stores a coolant (e.g., an active coolant, a liquid coolant, a cryogenic coolant, etc.).

The pump <NUM> is onboard the second stage rocket <NUM> and receives the coolant from the tank <NUM>. The pump <NUM> outputs a pressurized coolant (e.g., a coolant having a pressure of several hundred psi or higher). That is, the pressure of the coolant is greater after it passes through the pump <NUM> than it is when stored in the tank <NUM>. The coolant is transferred from the tank <NUM> to the pump <NUM> via a coolant conduit <NUM> (e.g., ducting, tubing, etc.).

The heat exchanger <NUM> is onboard the second stage rocket <NUM> and is integrally connected with the heat shield <NUM>. The heat exchanger <NUM> receives the pressurized coolant from the pump <NUM>, transfers heat from the heat shield <NUM> to the pressurized coolant to generate a heated fluid (e.g., a gas, a supercritical fluid, etc.), and outputs the heated fluid. The pressurized coolant is transferred from the pump <NUM> to the heat exchanger <NUM> via a pressurized coolant conduit <NUM> (e.g., ducting, tubing, etc.).

The turbine <NUM> is onboard the second stage rocket <NUM> and includes an inlet <NUM>, a shaft <NUM>, and an outlet <NUM>. The inlet <NUM> receives the heated fluid that is output from the heat exchanger <NUM> via a primary heated fluid conduit <NUM> (e.g., ducting, tubing, etc.). In some embodiments, there will be an excess of energy in the heated fluid which will be bypassed around the turbine <NUM> via a bypass conduit <NUM> (e.g., ducting, tubing, etc.) and used to pressurize or power an auxiliary system <NUM> (e.g., a tank, a gas thruster, a transpiration cooling system, an auxiliary power unit (APU), etc.). The shaft <NUM> of the turbine <NUM> is coupled (e.g., directly coupled, indirectly coupled via a coupler, etc.) to the pump <NUM> and includes turbine blades (not shown) mounted thereon. The shaft <NUM> rotates and thereby powers the pump <NUM> when the heated fluid received from the heat exchanger <NUM> acts on the turbine blades (not shown). The outlet <NUM> of the turbine <NUM> outputs the heated fluid. In some embodiments, the second stage rocket <NUM> includes an exhaust conduit <NUM> through which the heated fluid exits the second stage rocket <NUM> (e.g., for providing thrust). Additionally or alternatively, the heated fluid output from the outlet <NUM> of the turbine <NUM> can be used to pressurize or power an auxiliary system (e.g., a tank, a gas thruster, a transpiration cooling system, an APU, etc.). The heated fluid that is output from the outlet <NUM> of the turbine <NUM> will have a pressure and energy that is less than that of the heated fluid diverted through the bypass conduit <NUM> to the auxiliary system <NUM>. In some embodiments, the heat shield system <NUM> further includes bearings, gears, and/or seals (not shown) that facilitate the coupling of the turbine <NUM> and the pump <NUM> via the shaft <NUM>.

During operation of the heat shield system <NUM>, the pressurized coolant passes into the heat exchanger <NUM> (e.g., into channels or other conduits formed in the heat shield <NUM>) and picks up heat <NUM> at a rate (i.e., a heat flux) that is typical of hypersonic and re-entry vehicles and may be in the range of <NUM>-<NUM> BTU/in<NUM>-s, for example. The heat exchanger <NUM> serves a dual purpose of cooling the windward side <NUM> of the second stage rocket <NUM>, and adding energy to the coolant which is used to drive the turbine <NUM> and then power the pump <NUM>. The pressure of the coolant drops while overall enthalpy increases along the heat exchanger <NUM>, until the coolant exits the heat exchanger <NUM> as a heated fluid. The primary flow of heated fluid enters the turbine <NUM>, where energy is extracted. The heated fluid exiting the turbine <NUM> is expelled out of the second stage rocket <NUM> into the external environment <NUM> or used for another purpose (e.g., re-chilled by onboard systems and passed back through the heat exchanger <NUM> in a closed loop cycle).

In some embodiments, the pressure of the coolant in the tank <NUM> alone provides enough energy to start spinning the turbine <NUM> and pump <NUM>, creating an increasing pressure and increasing power available to the turbine <NUM>. In other embodiments, the pressure of the coolant in the tank <NUM> does not provide enough energy to start spinning the turbine <NUM> and pump <NUM>. In some such embodiments, the heat shield system <NUM> further includes an external starter source, such as a motor connected to the turbine shaft <NUM> or a high pressure gas directed at the turbine <NUM>.

In some embodiments, at least one component (e.g., the tank <NUM>, the pump <NUM>, the turbine <NUM>, etc.) is an already-existing component of the engine <NUM>. For example, in some embodiments the engine <NUM> includes at least a pump and a turbine which push coolant through a heat exchanger of the engine. In such embodiments, the fuel pump and the turbine of the engine <NUM> serve dual-purposes by functioning as the pump <NUM> and the turbine <NUM> of the heat shield system <NUM>, respectively, and the heat exchanger of the engine <NUM> forms at least a portion of the heat exchanger <NUM> of the heat shield system <NUM>.

In some embodiments, the heat shield system <NUM> further includes at least one component that is additionally or alternatively cooled passively (e.g., using high temperature materials, etc.).

Once operation of the heat shield system <NUM> is started, thermal energy added to the coolant is enough to sustain operation. Specifically, the energy added to the coolant is enough to supply the turbine <NUM> with the required power to drive the pump <NUM> after taking into consideration all of the losses in the system <NUM>, including pump <NUM> and turbine <NUM> inefficiencies, pressure losses in the heat exchanger <NUM>, and other losses from friction and other mechanisms.

The coolant flowing through the heat exchanger <NUM> integrated into the windward side <NUM> of the second stage rocket <NUM> is enough to maintain acceptable temperatures on the heat shield <NUM> and other walls of the second stage rocket <NUM> while the second stage rocket <NUM> passes through the severe heat environment (e.g., while the second stage rocket <NUM> re-enters the atmosphere). The heat shield system <NUM> therefore enables the second stage rocket <NUM> to perform a base-first re-entry trajectory. This provides several key advantages over other proposed nose-first or body-first (a/k/a belly flop) strategies: (i) it eliminates the need for challenging in-atmosphere reorientation maneuver required for nose-first or body-first (a/k/a belly flop) reentry vehicles with vertical landing profiles; (ii) it keeps the primary load paths in the axial direction during all phases of flight, allowing fora more efficient structural solution; (iii) the common vertical orientation during ascent and reentry simplifies the cryogenic fluid management challenge by minimizing slosh and associated boil-off; and (iv) it minimizes the heat shield surface area while also maintaining a low ballistic coefficient, minimizing the overall heat load managed by the vehicle during re-entry.

Claim 1:
An upper stage rocket (<NUM>) of a multi-stage rocket system, the upper stage rocket (<NUM>) comprising:
an actively-cooled heat shield system (<NUM>) including:
a heat shield (<NUM>) defining a windward side (<NUM>) of the upper stage rocket (<NUM>); a tank (<NUM>) onboard the upper stage rocket (<NUM>), the tank (<NUM>) configured to store a coolant;
a pump (<NUM>) onboard the upper stage rocket (<NUM>), the pump (<NUM>) configured to receive the coolant from the tank (<NUM>) and output a pressurized coolant;
a heat exchanger (<NUM>) onboard the upper stage rocket (<NUM>), the heat exchanger (<NUM>) integrally connected with the heat shield (<NUM>) and configured to receive the pressurized coolant from the pump (<NUM>), transfer heat from the heat shield (<NUM>) to the pressurized coolant to generate a heated fluid, and output the heated fluid;
a turbine (<NUM>) onboard the upper stage rocket (<NUM>), the turbine (<NUM>) including:
- an inlet (<NUM>) configured to receive the heated fluid output from the heat exchanger (<NUM>);
- a shaft (<NUM>) coupled to the pump (<NUM>) and including turbine blades mounted thereon, the shaft (<NUM>) configured to rotate and thereby power the pump (<NUM>) when the heated fluid received from the heat exchanger (<NUM>) acts on the turbine blades; and
- an outlet (<NUM>) configured to output the heated fluid; and
a propulsion engine (<NUM>) disposed at an aft end (<NUM>) of the upper stage rocket (<NUM>); wherein the aft end (<NUM>) defines the windward side (<NUM>) of the upper stage rocket during operation of the heat shield system (<NUM>).