Spacecraft systems and methods

Systems and methods for configuring, packaging, and deploying spacecraft are provided. More particularly, spacecraft are configured with statically mounted instruments. An end of the spacecraft to which the instruments are mounted is relatively distant from a spacecraft bus, and has a narrow width relative to the spacecraft bus. During launch multiple overlapping and interleaved spacecraft are disposed radially about a longitudinal axis of the launch vehicle.

FIELD

Systems and methods for locating sensors on spacecraft and facilitating mounting multiple spacecraft to a launch vehicle are provided.

SUMMARY

Spacecraft, including but not limited to satellites, are expensive to produce and launch. In an effort to drive down the costs associated with deploying spacecraft, efforts have been made to simplify their design. In addition, systems have been developed that allow multiple spacecraft to be deployed from a single launch vehicle.

In order to accommodate multiple spacecraft within a single launch vehicle, various arrangements have been proposed or implemented. For example, spacecraft have been stacked one on top of the other. This configuration allows the centers of mass of the spacecraft to be aligned with the center axis of the launch vehicle. However, a vertically stacked arrangement severely limits the height of each spacecraft. As another example, spacecraft have been arranged within a launch vehicle in a side-by-side configuration. As still another example, small spacecraft have been disposed radially about the center axis of the launch vehicle. A particular implementation of this type of arrangement is the Evolved Expendable Launch Vehicle Secondary Payload Adaptor or ESPA ring. Moreover, multiple ESPA rings can be stacked, one on top of the another. Although such arrangements can facilitate the deployment of multiple spacecraft from a single launch vehicle, they do not provide for efficient packaging of certain spacecraft configurations.

In addition to considerations related to enabling the efficient deployment of multiple spacecraft, individual spacecraft must be capable of carrying desired and necessary payload components. This often includes the need to provide deployable booms that enable sensitive instruments to be positioned away from sources of interference, such as power supplies and propulsion units, while maintaining as compact an overall structure as possible. However, the inclusion of deployable booms adds to the expense and complexity of the spacecraft.

Accordingly, it would be desirable to provide a spacecraft that was capable of being efficiently packaged within a launch vehicle fairing, while providing suitable instrument operating environments.

DETAILED DESCRIPTION

FIG.1depicts portions of a spacecraft system100in accordance with embodiments of the present disclosure. More particularly, the figure depicts a payload portion of a launch vehicle104and multiple spacecraft108carried by the launch vehicle104in accordance with embodiments of the present disclosure in a side cross section view. As depicted in the figure, the launch vehicle104can include a nose cone or fairing112defining a payload volume116in which the spacecraft108or other payloads can be carried from a launch site to a location in orbit or in interstellar space. As examples but without limitation, the spacecraft108can include satellites, space vehicles, or the like that are intended for separation from the launch vehicle104once the launch vehicle104has arrived at a predetermined location. The spacecraft108are interconnected to either of a first payload or mounting ring120aor a second payload or mounting ring120b. In accordance with at least some embodiments of the present disclosure, the first mounting ring120ais interconnected to a launch vehicle interface128provided as part of the launch vehicle104structure, and the second mounting ring120bis stacked on top of the first mounting ring120a.

With reference now toFIGS.2,3A and3B, particular features of the arrangement of the spacecraft108within the payload volume116are illustrated. In particular,FIG.2depicts an interior of the launch vehicle104payload volume116and multiple spacecraft108carried by the launch vehicle104as in the example configuration ofFIG.1in an end view taken from the front or top of the launch vehicle104.FIGS.3A and3Bdepict the interior of the launch vehicle104payload volume116and the interconnection between the spacecraft108and the payload rings or fixtures120in accordance with embodiments of the present disclosure in side views. As can be seen from these figures, this example configuration includes six spacecraft108disposed radially about a center or longitudinal axis124of the launch vehicle104. A first set108aof three spacecraft108a1-108a3are connected to the first mounting ring120ausing fixtures spaced radially about the mounting ring120a, and are oriented in a first direction. A second set108bof three spacecraft108b1-108b2are connected to the second mounting ring120busing fixtures spaced radially about the mounting ring120b, and are oriented in a second direction that is opposite the first direction. Moreover, longitudinal axes304of each of the spacecraft108are generally aligned with the longitudinal axis124of the launch vehicle124. For instance, some or all of the longitudinal axes304of the spacecraft108can be parallel to the longitudinal axis124of the launch vehicle104. As another example, the longitudinal axis304of some or all of the individual spacecraft108can be non-parallel to the longitudinal axis124of the space craft104, but can be aligned with the longitudinal axis124of the launch vehicle104in that the axis304of an individual spacecraft108and the longitudinal axis124of the space craft104can define an individual plane. Although the figures depict the first108aand second108bsets of spacecraft108as each containing three spacecraft108, other configurations are possible.

In accordance with embodiments of the present disclosure, the first mounting ring120ais connected to a launch vehicle interface or platform128, and the second mounting ring120bis stacked on top of and is connected to the first mounting ring120a. In accordance with further embodiments of the present disclosure, a spacer element122or set of spacer elements122can be disposed between the first120aand second120bmounting rings, to maintain a desired axial spacing (seeFIG.3A). Alternatively or in addition, a spacer element122can be disposed between the launch vehicle platform128and a mounting ring. For example, a first spacer element122acan be disposed between the launch vehicle platform128and a first mounting ring120a, and a second spacer element122bcan be disposed between the first mounting ring120aand a second mounting ring120b(seeFIG.3B). Moreover, the spacecraft108overlap and are interleaved with one another while connected to the launch vehicle, such that a section taken transverse to the longitudinal axis124of the launch vehicle104and through the interface of the first120aand second120bmounting rings intersects all of the spacecraft. In addition, a first end404of each of the spacecraft in the first set108ais adjacent the second mounting fixture120bor closer to the second mounting fixture120bthan to the first mounting fixture120a, and a first end404of each of the spacecraft in the second set108bis adjacent the first mounting fixture120aor closer to the first mounting fixture120athan to the second mounting fixture120b. The first end404of each of the spacecraft108aconnected to the first mounting ring120agenerally points toward a first end of the launch vehicle104, which includes the nose cone112, while a first end404of each of the spacecraft108bconnected to the second mounting120bgenerally points in the opposite direction, towards the launch vehicle104main engines (not shown). As depicted, spacecraft solar panels132can be maintained in a folded state when the spacecraft108are stowed in the payload volume116. Also as depicted the spacecraft108associated with a particular launch vehicle104can be identical or similar to one another. Alternatively, the spacecraft108can differ from one another.

With reference also now toFIG.4, each spacecraft108includes a first end404and a second end408. An instrument platform412to which one or more instruments or sensors416are mounted is located at or towards the first end404of the spacecraft108. A spacecraft bus420portion is located at or towards the second end408of the spacecraft108. In accordance with embodiments of the present disclosure, the spacecraft bus420is connected to a mounting fixture or port of a mounting ring120during launch and delivery of the spacecraft108to a deployment location by the launch vehicle104. Accordingly, the spacecraft bus420can include fixtures or features for mating with the mounting fixture of the mounting ring. The spacecraft bus420portion can include, for example, a power supply424and hinge points428for solar panels132. Note that inFIG.4the solar panels132of the illustrated spacecraft108are depicted in a deployed state. In addition, an antenna or instrument432, for example an Earth facing antenna or instrument, can be located on an Earth facing or base platform436at or adjacent the second end408of the spacecraft108. The instrument platform412can be connected to the spacecraft bus420portion by a structural or spacer member440. In general, the spacer member440extends from the spacecraft bus420to enable the instruments416to be positioned far enough away from the spacecraft bus420to ensure that electromagnetic interference from components located in the spacecraft bus420does not adversely affect the operation of the instruments416. A single spacecraft108can include one or more instruments416of one or more types. The location of particular instruments416relative to other instruments416and to the spacecraft108components can be determined by the particular operating requirements of the instruments416, including isolation requirements, field of view requirements, weight or the like.

In accordance with embodiments of the present disclosure, the spacer member440is fixed relative to the spacecraft bus420and the instrument platform412. Thus, a location of the instrument platform412relative to the spacecraft bus420is fixed. In accordance with at least some embodiments of the present disclosure, the spacer member440forms some or all of the instrument platform412. In accordance with still further embodiments of the present disclosure, the instruments416are fixed to the instrument platform412. Moreover, the instruments416can have locations relative to the spacecraft bus420that are the same during operation of those instruments416as during launch. That is, the instruments416in accordance with at least some embodiments of the present disclosure are not deployed on booms. The spacer member440can include one or more stiffener members444. Moreover, the spacer member440can be formed using various structures, including but not limited to aluminum panels having honeycomb cores, composite panels, interconnected struts, an integral beam, or any other material or structure that provides the required strength, stiffness, and weight requirements. The spacer member440and instrument platform412can also be configured in view of its interaction with the operational requirements of the instruments416, such as instrument416operating wavelengths, shielding or shading requirements, or the like.

In general, it is desirable for a spacecraft to be relatively compact, in order to reduce the associated moment of inertia (MOI). In conventional spacecraft designs, this has often resulted in the adoption of cube or rectangular forms, with sensitive instruments positioned at the end of relatively flexible booms, in order to provide those instruments with a desired field of view and to ensure isolation from sources of EMI. In accordance with embodiments of the present disclosure, expensive, delicate, and failure prone booms are avoided or eliminated. Instead, instruments416are statically mounted to an instrument platform412that is spaced apart from the spacecraft bus420by a structural member440. Moreover, the structural member440is itself static or fixed relative to the spacecraft bus420and is relatively stiff. In order to reduce the MOI of the spacecraft108, the structural member440, any stiffener members444, and/or the platform412can feature a transverse dimension that decreases with increasing distance from the spacecraft bus420. Accordingly, a width of the spacecraft108at the first end404can be several times or more less than a width of the spacecraft108at the second end408. Alternatively or in addition, the mass and/or number of instruments416can be reduced with increasing distance from the spacecraft bus420. Accordingly, the center of gravity of a spacecraft108in accordance with embodiments of the present disclosure can be maintained at a normal location. For example, but without limitation, the center of gravity of a spacecraft108can be centered within the spacecraft bus420.

In addition to various benefits related to providing a relatively inexpensive, simple, robust, and reliable spacecraft108assembly, embodiments of the present disclosure provide packaging benefits. In particular, as described herein, multiple spacecraft108can be interleaved with one another and carried within the payload volume116of conventional launch vehicles104using standards compliant mounting structures or fixtures120. As examples, but without limitation, sensitive instruments416can be located at least 2.5 meters from the spacecraft bus420. Moreover, while a conventional spacecraft will have relative dimensions in L:W:H of about 1:1:1, a spacecraft in accordance with embodiments of the present disclosure can have, as an example but without limitation, relative dimensions of 5.9:1.3:1 (L:W:H).

FIGS.5A and5Bdepict sensor416fields of view504in accordance with embodiments of the present disclosure. In particular, in addition to ensuring that EMI from spacecraft bus420components does not interfere with the operation of instruments416carried by the spacecraft108, the locating of instruments416on an instrument platform412that is spaced apart from the spacecraft bus420by a spacer member440can ensure wide and/or multiple fields of view for the sensors416. For example, embodiments of the present disclosure can provide a field of view to an instrument416that extends in multiple directions by using a fixed spacer member440to position the portion of the instrument platform412to which the instrument416is connected at a suitable location relative to the spacecraft bus420. Moreover, if desired or required, the fields of view504of different instruments416can be overlapping, partially overlapping, and/or separated from one another. In addition, a field of view508of an instrument432at the second end408of the spacecraft108is depicted.

FIG.6depicts a method for developing and deploying multiple spacecraft108carrying sensors in accordance with embodiments of the present disclosure. Initially, an instrument or set of instruments416that will be carried on the instrument platform412of a spacecraft108are selected (step604). Next, the requirements of operational support components, such as power supplies424, communications equipment432, control equipment, and other potentially interfering components that will be mounted to or contained within the spacecraft bus420are determined (step608). A required separation to maintain desired isolation levels between the instruments416on the instrument platform412and the components of the spacecraft bus420is then determined (step612). From the set of instruments416, the sizes of those instruments416, the support components, and the sizes of those components, the sizing of the spacecraft108, including the sizing of the structural member440and instrument platform412, is determined (step616).

At step620, the configuration of multiple spacecraft108and the number of mounting fixtures120within the payload volume116of the launch vehicle is determined. As can be appreciated by one of skill in the art after consideration of the present disclosure, the dimensions of the spacecraft108can be selected in view of the space available within the payload volume116of the launch vehicle104. Accordingly, aspects of the dimensioning of spacecraft108and determining a packaging configuration for the spacecraft108within the payload volume116can be iterative. In general, embodiments of the present disclosure feature at least two mounting rings120that each have a plurality of radially disposed ports or fixtures spaced at regular intervals around an outside diameter of the mounting rings120. The ports of a first one of the mounting rings120can be radially aligned with the ports of a second one of the mounting rings120. Alternatively, the ports of one mounting ring120can be rotated relative to the ports of another mounting ring120. For instance, the ports of a first mounting ring120acan be aligned so that they fall between the radial locations of the ports of a second mounting ring120b.

As noted, the process of dimensioning and configuring the packaging of the spacecraft108can be an iterative process. Accordingly, at step624a determination can be made as to whether the design process is complete. If not, for example if changes to the mission parameters have been made in the interim, or if it becomes apparent that revisions to the design may be beneficial for other reasons, the process returns to step604. Otherwise, the process continues to the production of the spacecraft108and the configuration of the launch vehicle104(step628).

At step632, the spacecraft108are joined to the mounting rings120of the launch vehicle104. This can include joining the spacecraft bus420of each spacecraft108in a first set of spacecraft108ato mounting fixtures included as part of the first mounting ring120a, and joining the spacecraft bus420of each spacecraft in a second set of spacecraft108bto mounting fixtures included as part of the second mounting ring120b. Optionally, an instrument platform412of each spacecraft in the first set of spacecraft108bcan be joined to fixtures included as part of the second mounting ring120b, and an instrument platform412of each spacecraft in the second set of spacecraft108bcan be joined to fixtures included as part of the first mounting ring120a.

The launch vehicle104can then be launched (step636). Once the launch vehicle has arrived at the desired location, the spacecraft108can be deployed (step640). As can be appreciated by one of skill in the art after consideration of the present disclosure, the spacecraft108can all be deployed at the same location at or about the same time, or sequentially as the launch vehicle104reaches the desired locations for different ones of the spacecraft108. The process can then end.

In accordance with embodiments of the present disclosure, various standardized mounting fixture120components, such as but not limited to ESPA rings, can be used to mount the spacecraft108to the launch vehicle104. For instance, as depicted inFIGS.1-3, two ESPA rings or ESPA Grande rings can be stacked one on top of the other. Alternatively, one or more spacer elements can be interposed between adjacent mounting rings120, to maintain axial spacing for the spacecraft108. All or a subset of the radial ports of each of the mounting rings120can be utilized. For example, three of the radial ports of each ESPA ring in a pair of six port ESPA rings can be utilized to attach the spacecraft bus420portions of three spacecraft108, enabling six spacecraft108to be carried within the payload volume114. Mounting rings120having any number of ports or fixtures can be used. As examples, but without limitation, the ports can include 380 mm ESPA ring ports or 610 mm ESPA Grande ports. In addition, the length of the spacecraft108can extend to greater than the height of a standard spacer120by interleaving adjacent spacecraft108. That is, embodiments of the present disclosure enable spacecraft108having dimensions that extend beyond standardized payload envelopes to be carried and deployed by a launch vehicle104. In accordance with still other embodiments, more than two mounting fixtures120can be stacked on top of one another.

The spacecraft108carried by a launch vehicle104can be the same as or similar to one another. Alternatively, the spacecraft108in a first set of spacecraft108acan be the same or similar to one another, while the spacecraft108in a second set of spacecraft108bcan be the same or similar to one another, but different from the spacecraft108in the first set of spacecraft108a. In accordance with still other embodiments of the present disclosure, one or more spacecraft108in accordance with embodiments of the present disclosure can be carried in the payload volume116of a launch vehicle104, at the same time that spacecraft or equipment of other configurations, including conventional configurations having deployable booms, cube satellites, etc., are carried in the payload volume116of the launch vehicle104.

Advantages of spacecraft108configured in accordance with embodiments of the present disclosure include eliminating or reducing the number of complex, expensive, and failure prone boom components. In addition, spacecraft108as disclosed herein can be provided with a large MOI, for use with a gravity-gradient, attitude stabilization approach. The area available for antennas, including but not limited to planar array antennas, can be greater on a spacecraft108configured in accordance with embodiments of the present disclosure as compared to other designs. Moreover, greater area can be available for solar panels132for power generation. In addition, spacecraft108in accordance with embodiments of the present disclosure can be efficiently disposed within the payload volume116of a launch vehicle104.

The foregoing description has been presented for purposes of illustration and description. Further, the description is not intended to limit the disclosed systems and methods to the forms disclosed herein. Consequently, variations and modifications commensurate with the above teachings, within the skill or knowledge of the relevant art, are within the scope of the present disclosure. The embodiments described hereinabove are further intended to explain the best mode presently known of practicing the disclosed systems and methods, and to enable others skilled in the art to utilize the disclosed systems and methods in such or in other embodiments and with various modifications required by the particular application or use. It is intended that the appended claims be construed to include alternative embodiments to the extent permitted by the prior art.