BACKGROUND OF THE INVENTION

The present invention relates generally to spacecraft engineering and design. More particularly, the present invention relates to a method of realigning a spacecraft frame or relieving stress within a spacecraft frame, while in space. The present invention also relates to the activation and deactivation of structural connectors of spacecraft modules.

A spacecraft experiences substantial tensile, compressive and shearing stresses during its launch to orbit which can have a detrimental effect on the spacecraft's frame. In addition, a spacecraft undergoes additional cyclic tensile, compressive and shearing stresses once in space due thermal fluctuations caused by the spacecraft's rotation relative to the sun, and due to vibration, such as produced by the rotation of momentum wheels and control moment gyros. The movement and rotation of a spacecraft payload and the deployment of a spacecraft's solar panels can also induce minor stress into the spacecraft frame. Individually or cumulatively, these stresses can cause the alignment of the spacecraft's structural members to change which, in turn, can induce stress within the spacecraft's structural members or have detrimental effects on the orientation of the spacecraft payload. Unfortunately, there is no known method of relieving the stress within the spacecraft's structural members or of realigning the structural members in the event that they have become misaligned.

Modular spacecraft, homogenous cellular spacecraft, and CubeSats and the like are particularly prone to having components shift in space. Modular spacecraft are attempts to transition away from custom designed spacecraft toward multi-use designs and mass production in an effort to reduce the cost of spacecraft development. A modular spacecraft includes a bus structure which provides a general-purpose spacecraft platform. Various payloads can be mounted to the spacecraft platform utilizing a standard structural, fuel and electric interface. Again, the mechanical connector is of paramount importance.

Even more recently, spacecraft have been designed by combining homogeneous cells. Each cell possesses the traditional architecture of a spacecraft including structure, power, fuel, attitude control and determination, satellite processing, etc. Each cell is substantially identical so as to be manufactured inexpensively and quickly. These cells are combined to create larger and larger platforms to support payload functions such as communications and surveillance. Though there is substantial redundancy by each cell incorporating all spacecraft sub-system capabilities, the extra costs are more than made up for by mass production savings and rapid assembly. Of course, the satellite reconfigurable cell concept also requires that cells be interconnected by a structural, fluid and electrical interface.

CubeSats are a class of small satellites that use a standard size and form factor. The standard CubeSat size uses a “one unit” or “1 U” measuring 10×10×10 cm. To form larger spacecraft, two or more CubeSat modules are connected prior to launch to form a spacecraft of larger sizes such as 2, 3, 6, and even 12 U.

There are now several companies that provide off-the-shelf exoskeleton structures (often called frames or chassis) which mechanically encompass the CubeSat modules. Most of the frames are machined from 6061-T6 or 7075 aluminum and are designed with several mounting locations for mounting components in an attempt to offer configuration flexibility for spacecraft designers. 1 U, 3 U and 6 U sized frames are mose prevalent; however, 12 U frames are available as well. The structural designs come are either: 1) formed from a solid block of aluminum, or 2) those assembled from multiple frames. There are pros and cons associated with each design approach. Solid body designs tend to be lighter and more rigid because they do not experience concentrated stresses due to fasteners during assembly. However, forming thin shapes from solid blocks of aluminum can leave residual internal stresses in the structure, which can be difficult to detect.

Unfortunately, due to the higher number of attachment points, components within and upon the frames of modular, homogenous and CubeSat satellites can become misaligned due to the environmental factors described above. Also, there is no known manner of realigning a spacecraft's frame or adjusting the position of spacecraft's modules once the spacecraft is in space.

Thus, there is a need for a method and apparatus for relieving mechanical stress within a spacecraft frame.

There is also a need for a method and apparatus for realigning the structural components of a spacecraft frame.

SUMMARY OF THE INVENTION

The present invention addresses the aforementioned disadvantages by providing a method of realigning or relieving stress within a spacecraft frame. The spacecraft includes a frame having at least two frame subsections, wherein the term “frame” is intended to be interpreted broadly to include any structure or housing of two units to be connected. The frame subsections are connected by a mechanical connector which connects the two frame subsections. Where the spacecraft includes more than two frame subsections, preferably each subsection is connected by the mechanical connector. Importantly, the mechanical connector is capable of providing both a soft dock connection and a hard dock connection between the two frame subsections. The soft dock connection mechanically connects the two frame subsections, but allows relatives movement between the two frame sections. Conversely, the hard dock connection mechanically connects the two frame subsections, but does not allow relative movement between the two frame sections. Where the spacecraft includes more than two frame subsections, preferably each subsection is connected by a mechanical connector which is capable of providing both a soft dock connection and a hard dock connection.

To control the selective activation of a hard dock connection or soft dock connection, the mechanical connector is connected to one or more control processors, referred to herein as a controller. The controller may be a general purpose computer or microprocessor including hardware and software as can be determined by those skilled in the art to provide automated or directed control of mechanical connector. Preferably, but not necessarily, the spacecraft includes a sensor for determining whether spacecraft frame is aligned or misaligned, or determines whether the spacecraft frame is undergoing strain, which in turn indicates the spacecraft frame is undergoing mechanical stress.

It is anticipated that the spacecraft will be launched while the one or more mechanical connectors are in a hard dock connection. However, due launch vibration, payload vibration, or thermal effects in space, etc., the spacecraft frame alignment may change for which realignment is preferred. Misalignment of the spacecraft may be detected by an alignment/stress sensor. Alternatively, such as where the spacecraft does not possess such a sensor, the spacecraft may undergo periodic maintenance realignment procedures.

To realign the spacecraft frame and/or relieve stress within the spacecraft frame while the spacecraft is in space, the controller causes one or more of the mechanical connectors to transition from the hard dock connection to the soft dock connection. The mechanical connector is maintained in a soft dock connection for a predetermined time period to realign or relieve stress within the spacecraft frame. The soft dock connection may last any length of time as can be determined by one skilled in the art, such as a few milliseconds to ten (10) seconds or longer. Thereafter, the controller causes the mechanical connector to transition from the soft dock connection back to the hard dock connection.

The alignment/stress sensor may be of any type as can be determined by one skilled in the art for determining structural alignment or mechanical strain. A first preferred alignment/strain sensor is simply a strain gauge affixed to the spacecraft frame or affixed to a mechanical connector which transmits deformation measurements to the controller. An alternative preferred alignment/strain sensor includes one or more star trackers located at one or more locations on the spacecraft frame. A star tracker is an optical sensor that measures the positions of stars using photocells or a camera. Since positions of many stars have been measured to a high degree of accuracy, a star tracker may be used to determine the orientation (or attitude) of the spacecraft with respect to the stars. In order to do this, the star tracker must obtain an image of the stars, measure their apparent position in the reference frame of the spacecraft, and identify the stars so their position can be compared with their known position from a star catalog. Meanwhile, to use the star trackers to sense frame alignment (or possible frame stress), the measurements of a plurality of start trackers are compared, and any shift in spacecraft orientation (or attitude) of a first star tracker's measurements relative to a second star tracker's measurements indicates that the frame alignment has changed.

The mechanical connector can be any connector which can be automatically or remotely controlled to switch from a hard connection to a soft connection, and back to a hard connection. A simple connector could consist of a bolt that can be remotely loosened, such as by a motor, and then tightened again by the motor. In the preferred embodiment, the mechanical connector includes a pair of male connector assemblies affixed to a first frame subsection and a pair of mating female mounting connector assemblies affixed to a second frame subsection. Each of the male connector assemblies include a pneumatically controlled piston, and each of the first and second male connector assemblies are mounted to the first frame subsection in outwardly facing coaxially aligned relationship to one another such that extension of both of the pistons causes the pistons to move axially outward and away from each other, and retraction of the pistons causes the pistons to move axially inward and toward each other.

Conversely, each of female connector assemblies possess a receptacle sized for receipt of a male connector assembly piston. The first and second female connector assemblies are mounted to the second frame in an inwardly facing coaxially aligned relation to one another wherein the female connector assemblies' central bores define the same axis. Furthermore, the first and second female connector assemblies are mounted to the second frame subsection in an inwardly facing relationship and spacing so as to accept the male connector assemblies between the female connector assemblies when the male connector assembly pistons are fully retracted. However, the male and female connector assemblies are spaced to lock together when the male connector assembly pistons have been pneumatically made to extend and project into the female connector assembly's receptacles.

Importantly, the mechanical connectors are capable of providing a hard dock connection by extending the male assembly pistons to fully extended positions wherein the pistons engage their corresponding female connector assembly. Furthermore, the mechanical connectors are capable of providing a soft dock connection by only partially retracting each piston so that the pistons no longer physically engage a female connector assembly, but each piston remains partially within a corresponding receptable. Thus, when in a soft dock condition, meaning the pistons have been partially retracted, a first spacecraft frame subsection can move slightly relative to a second spacecraft frame subsection, but the frame sections will not disengage. The extension and retraction of the piston can be controlled by various mechanical apparatus as can be determined by those skilled in the art. However, in a preferred embodiment, the position of the piston is controlled by a helical spring, pneumatic extension and electromagnetic retraction.

Preferably more and more homogenous satellite cells are combined utilizing the mechanical connectors of the present invention to create a larger spacecraft comprised of more and more frame subsections. To form a satellite of homogenous cells, two spacecraft homogeneous cells are connected by a mechanical connector including two male connector assemblies and two female connector assemblies. The first and second male connector assemblies are mounted to a first homogenous satellite cell in an outwardly facing coaxially aligned relation to one another. More specifically, each male connector assembly's central bore and piston have the same central axis and the male connector assemblies are mounted to the first homogenous satellite cell in an opposing relation to one another such that extension of the pistons causes the pistons to move axially outward from one another, and retraction of the piston causes them to move axially inward toward each other.

Meanwhile, the first and second female connector assemblies are mounted to a second homogenous satellite cell. Like the male connector assemblies, the female connector assemblies are mounted to the homogenous satellite cell in a manner wherein their central bores are coaxially aligned. However, instead of facing outward, the female connector assemblies are mounted to the second spacecraft in an inwardly facing relationship to one another. Furthermore, the female connector assemblies are mounted so as to be spaced so as to receive the two male connector assemblies, mounted to the first spacecraft, between the female connector assemblies when the male connector assembly pistons are in a retracted condition. However, the female connector assemblies are mounted to the second spacecraft so as to be sufficiently close together that the male connector assemblies and female connector assemblies will lock together when the male connector assembly pistons have projected into an extended position with the pistons projecting into the female connector assembly receptacles. More specifically, the male and female connector assemblies form mechanical connectors when the male connector assembly pistons are extended into the female connector assembly receptacles so as to engage an O-ring concentrically positioned within the female connector assembly's central bores.

Advantageously, the satellite's framework (comprised of homogenous cells) can be realigned and/or have mechanical stress relieved within the spacecraft framework while the spacecraft is in space. To accomplish this, the controller causes one or more of the mechanical connectors to transition from the hard dock connection to the soft dock connection. After the one or more mechanical connectors have maintained a soft dock connection for a predetermined time period the controller causes the mechanical connector to transition from the soft dock connection back to the hard dock connection.

Advantageously, it is an object of the present invention to provide a method of realigning the structural components of a spacecraft frame while the spacecraft is in space.

In addition, it is an object of the present invention to provide a method of relieving stress within the structural components of a spacecraft frame while the spacecraft is in space.

Other features and advantages of the present invention will be appreciated by those skilled in the art upon reading the detailed description which follows with reference to the attached drawings.

