This application claims the benefit of U.S. Provisional Application No. 63/266,274, filed Dec. 30, 2021, which is hereby incorporated by reference in its entirety.

The technology relates generally to reusable rockets, in particular to reusable upper stage rockets using an aerospike engine.

Single use rockets increase the cost for access to space. Reusable rockets provide for multiple uses and thus reduce the cost. Typical reusable rockets for a multi-stage space launch system are for the lower (first) stage only. It is therefore desirable to have reusability of the upper stage(s) of the space launch system.

The embodiments disclosed herein each have several aspects no single one of which is solely responsible for the disclosure’s desirable attributes. Without limiting the scope of this disclosure, its more prominent features will now be briefly discussed. After considering this discussion, and particularly after reading the section entitled “Detailed Description” one will understand how the features of the embodiments described herein provide advantages over existing approaches to space launch systems by presenting features for reusability of the upper stage rocket.

Described herein are systems and devices for a fully reusable upper stage for a multi-stage launch vehicle. The upper stage uses an aerospike engine for main propulsion and for vertical landing. At least a portion of the aerospike engine may be integrated into a heat shield configured to dissipate heat encountered upon reentry of the upper stage rocket. At least a portion of the heat shield may be actively cooled.

In some configurations, a reusable upper stage for a multi-stage launch vehicle may include: a rocket body extending from a forward end to an aft end and defining a longitudinal axis; a heat shield at the aft end of the rocket body, the heat shield may be configured to spherically cap the aft end of the rocket body and dissipate heat encountered upon reentry of the upper stage into the atmosphere; the heat shield may include a plurality of scarfed nozzles arranged radially around the longitudinal axis of the rocket body, wherein at least an inboard portion of each scarfed nozzle defines at least a portion of an aerospike contour; and an aerospike engine located at the aft end of the rocket body, the aerospike engine may include: at least one powerpack configured to pump propellant at high pressure to a plurality of thruster modules, wherein at least a portion of each thruster module of the plurality of thruster modules is inset into a respective scarfed nozzle of the plurality of scarfed nozzles of the heat shield, the plurality of thruster modules configured to eject exhaust along at least the inboard portions of the scarfed nozzles defining portions of the aerospike contour.

The reusable upper stage may further include secondary fluid injectors configured to eject fluid that joins with the exhaust ejected by the plurality of thruster modules.

The secondary fluid injectors may be arranged radially within the heat shield.

The plurality of thruster modules may be configured to eject exhaust and the secondary fluid injectors may be configured to eject fluid to form an exhaust plume that is similar in geometry to the exhaust plume from a full-length plug nozzle.

The heat shield may be actively cooled.

The heat shield may include a cooling circuit configured to dissipate heat encountered during reentry of the upper stage.

In some configurations, a multi-stage launch vehicle having a reusable upper stage may include: a lower stage configured to launch the vehicle from ground; an upper stage configured to separate from the lower stage, the upper stage may include: a rocket body extending from a forward end to an aft end and defining a longitudinal axis; a heat shield at the aft end of the rocket body, the heat shield may include a semi-spherical surface configured to cap the aft end of the rocket body and dissipate heat encountered upon reentry of the upper stage into the atmosphere, the heat shield may include: a plurality of scarfed nozzles embedded radially around the semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour; and an aerospike engine located at the aft end of the rocket body, the aerospike engine may include: at least one powerpack configured to pump propellant at high pressure to a plurality of thruster modules, wherein each thruster module of the plurality of thruster modules is inset into a respective scarfed nozzle of the plurality of scarfed nozzles of the heat shield, the plurality of thruster modules configured to eject exhaust along the aerospike contour collectively defined by inboard surfaces of the plurality of scarfed nozzles.

The upper stage may further include secondary fluid injectors configured to eject fluid that joins with the exhaust ejected by the plurality of thruster modules.

The secondary fluid injectors may be arranged radially within the heat shield.

The plurality of thruster modules may be configured to eject exhaust and the secondary fluid injectors may be configured to eject fluid to form an exhaust plume that is similar in geometry to the exhaust plume from a full-length plug nozzle.

The heat shield may be actively cooled.

The heat shield may include a cooling circuit configured to dissipate heat encountered during reentry of the upper stage.

In some configurations, a method of using an upper stage rocket may include: launching a multi-stage space vehicle from earth, the space vehicle may include: a lower stage may be configured to lift the space vehicle off ground; and a reusable upper stage may be configured to separate from the lower stage and to land vertically, the upper stage may include: a rocket body extending from a forward end to an aft end and defining a longitudinal axis; a heat shield at the aft end of the rocket body, the heat shield comprising a semi-spherical surface configured to cap the aft end of the rocket body and dissipate heat encountered upon reentry of the upper stage into the atmosphere, the heat shield may include: a plurality of scarfed nozzles embedded radially around the semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour; and an aerospike engine located at the aft end of the rocket body, the aerospike engine comprising: at least one powerpack may be configured to pump propellant at high pressure to a plurality of thruster modules, wherein each thruster module of the plurality of thruster modules is inset into a respective scarfed nozzle of the plurality of scarfed nozzles of the heat shield, the plurality of thruster modules configured to eject exhaust along the aerospike contour collectively defined by inboard surfaces of the plurality of scarfed nozzles.

The method may include descending the upper stage in an aft-end down orientation after separation from the lower stage; and ejecting the exhaust from the plurality of thruster modules to slow descent of the upper stage.

The method may include actively cooling the heat shield during descent of the upper stage.

The method may include landing the upper stage vertically.

The foregoing and other features of the present disclosure will become more fully apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. Understanding that these drawings depict only several embodiments in accordance with the disclosure and are not to be considered limiting of its scope, the disclosure will be described with additional specificity and detail through use of the accompanying drawings. In the following detailed description, reference is made to the accompanying drawings, which form a part hereof. In the drawings, similar symbols typically identify similar components, unless context dictates otherwise. The illustrative embodiments described in the detailed description, drawings, and claims are not meant to be limiting. Other embodiments may be utilized, and other changes may be made, without departing from the spirit or scope of the subject matter presented here. In some drawings, various structures according to embodiments of the present disclosure are schematically shown. However, the drawings are not necessarily drawn to scale, and some features may be enlarged while some features may be omitted for the sake of clarity. It will be readily understood that the aspects of the present disclosure, as generally described herein, and illustrated in the figures, can be arranged, substituted, combined, and designed in a wide variety of different configurations, all of which are explicitly contemplated and make part of this disclosure.

The following detailed description is directed to certain specific embodiments of the development. Reference in this specification to “one embodiment,” “an embodiment,” or “in some embodiments” means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the present disclosure. The appearances of the phrases “one embodiment,” “an embodiment,” or “in some embodiments” in various places in the specification are not necessarily all referring to the same embodiment, nor are separate or alternative embodiments necessarily mutually exclusive of other embodiments. Moreover, various features are described which may be exhibited by some embodiments and not by others.

Various embodiments will now be described with reference to the accompanying figures, wherein like numerals refer to like elements throughout. The terminology used in the description presented herein is not intended to be interpreted in any limited or restrictive manner, simply because it is being utilized in conjunction with a detailed description of certain specific embodiments of the development. Furthermore, embodiments of the development may include several novel features, no single one of which is solely responsible for its desirable attributes or which is essential to practicing the present disclosure.

The technology relates to a fully reusable upper stage for a multi-stage launch vehicle. The upper stage uses an aerospike engine for main propulsion and reentry and landing. An actively-cooled plug of the aerospike engine is used as a heat shield to enable engine-first re-entry. The aerospike engine can include a plurality of structures configured to direct exhaust in a way that is similar to the way exhaust is directed by a truncated aerospike plug. The plurality of structures may be embedded in a semi-spherical surface or segment of a heat shield at the aft end of the upper stage rocket. Inboard surfaces of the plurality of structures can collectively form a contour that is similar to the contour of a continuous annular ramp or surface of a truncated aerospike plug. Each structure of the plurality of structures can direct exhaust ejected from one thruster of a plurality of thrusters through the heat shield. In one non-limiting example, the plurality of structures includes a plurality of scarfed nozzles formed in a semi-spherical surface or segment of the heat shield at the aft end of the upper stage rocket. Embodiments of the reusable upper stage rocket according to the present disclosure may use the aerospike engine to descend and land vertically upright.

The technology described herein has multiple advantages and addresses multiple problems associated with conventional approaches to reusing upper stage rockets. For example, the systems and methods address the problem of surviving the re-entry heating environment. A re-entry heating profile may be generated with a high heating rate but over a very short pulse. The system may re-enter engine-first at a low angle-of-attack, which generates a high ballistic coefficient. The aerospike nozzle according to embodiments of the present disclosure takes the brunt of the heating and is actively cooled to withstand high temperatures, which is feasible because the heating pulse is short.

Another problem addressed by this system is the high cost and large amount of time typically required to develop new rocket engines. Systems and methods described herein address this problem by, in some embodiments, repurposing turbomachinery (e.g., powerpacks) and thrust chambers designed for other engines. For example, the systems may use several powerpacks and/or dozens of thrust chambers designed for use in other space vehicles. Advantages of repurposing existing hardware include reducing the cost, schedule, and risk of a new engine development.

Advantageously, embodiments of the aerospike engine systems and methods described herein that use multiple powerpacks and/or thrust chambers have the ability to digitally throttle by shutting down powerpacks and/or nozzles, thereby eliminating the risk of a single engine point-of-failure during landing (as typical designs only land on the center engine, which is thus a single point of failure). As another example, it may be advantageous to power down some number of powerpacks and/or nozzles to meet thrust requirements for a first mission, while employing all or substantially all of the powerpacks and/or nozzles for a second mission. Accordingly, the ability to digitally shut down powerpacks and/or nozzles is particular advantageous in the context of the reusable systems and methods described herein.

Another example advantage of the aerospike engine systems and methods described herein is that the upper stage may stay in the “engine down” orientation for the entire re-entry profile. This eliminates the need to reorient the vehicle (e.g., from a high angle-of-attack) just prior to landing. Another example advantage of this approach is that because aerospike engines are altitude compensating (i.e., they have near-optimal performance at both sea-level and in vacuum), the same engine can be used for ascent-to-orbit, deceleration burns during re-entry, and for landing.

Another advantage of the aerospike engine systems and methods according to embodiments of the present disclosure is the ability to be tested at sea-level. Most vacuum-optimized engines require highly specialized vacuum test stands or are mostly tested without high expansion ratio nozzles.

The lower stage 50 may launch the vehicle 10 from the ground. Booster rockets may or may not be used for launch. After launch and at a desired altitude, the lower stage 50, and any boosters, may separate from the vehicle 10 and fall to earth or be controllably landed for reuse. The second or upper stage 100 may then propel the payload to a faster velocity and/or higher altitude. The upper stage 100 may then separate from the payload. The upper stage 100 is then controllably descended for landing and reuse in subsequent flights. The upper stage 100 uses the aerospike engine to propel the payload as well as to controllably descend and land, as further described herein.

The aerospike engine 122 includes an aft end including portion 144. The engine 122 also includes a heat shield 140 located aft of the portion 144 and the continuous annular surface 130.

The aerospike engine 122 also includes a plurality of openings or pockets 131. The plurality of openings or pockets 131 may be arranged circumferentially around a longitudinal axis 133 of a rocket. The openings or pockets 131 direct exhaust from thrusters 150 onto the continuous annular surface 130. The continuous annular surface 130 may meet the heat shield at an aft end of the surface 130, for example at a point 132. The ejected exhaust may flow along the continuous annular surface 130 and join with fluid from secondary fluid injectors (not shown).

The combined flow from the ejected exhaust and secondary fluid may result in a plume 162. The plume 162 may have the general shape of an aerospike plug. The ejected exhaust may follow the continuous annular surface 130 and flow along a contour 160 to form the plume 162. Additional secondary fluid injectors loca