Abstract:
There is disclosed a Payload Ejection System (PES) able to store any set of payloads for launch and eject that set of payloads at a controlled speed with a low tumble rate while accommodating any offset centre of mass within a restricted volume. The need for ballasting or balancing is eliminated thus freeing up the design space for these payloads. Insensitivity to centre of mass location is enabled by the use of a deployment hinge assembly arrangement which uses two or more non-parallel folding hinge arrangements that allow for linear motion of the output link in one direction while restricting the motion all other directions. One embodiment of the current PES concept uses four (4) hinge panel assemblies, selected to provide optimal stiffness around the entire mechanism. The stiffness of the PES is integral to managing offset centre-of-mass locations by allowing the mechanism to translate the effective force vector to the center of mass location.

Description:
FIELD 
       [0001]    The present disclosure relates to systems for hosting payloads on a host spacecraft for the purpose of carrying the payload to orbit and ejecting it from the host spacecraft in a controlled manner. The payload ejection system is the entire system which enables the ejection of a payload from a host spacecraft in a controlled manner. Ejection in a controlled manner is defined as an ejection whereby the ejected payload has no angular momentum or linear momentum transverse to the ejection axis at the time of release. 
       BACKGROUND 
       [0002]    When satellites are launched to orbit (regardless of orbit type) there is often some launch vehicle mass and volume capacity that is not used. One purpose of the system disclosed herein is to use this surplus volume and mass capacity to deliver additional and separate payloads to orbit, from where the payload can proceed with its intended mission. This concept of delivering hosted payloads to particular orbits is described in “DARPA Phoenix Payload Orbital Delivery (POD) System: “FedEx to GEO”, Dr. Brook Sullivan et al, AAAIAA Space 2013 Conference and Exposition, Sep. 10-12, 2013, San Diego, Calif. As described in this paper, a payload includes but is not limited to such space systems as another small (micro or nano) spacecraft, replacement materials (e.g. fuel) to replenish another satellite, replacement components for on-orbit servicing repair of another spacecraft, components for in-space assembly of a new space system or spacecraft. 
         [0003]    Current orbital payload ejection systems require that the payload centre of mass be closely aligned with the centre of force of the ejection mechanism or else significant tumble rates (undesired angular rates and translational velocities transverse to the ejection axis at the time of release) are created at ejection, which is almost always considered a very negative condition. Accommodating an offset between the mechanism centre of force and the payload centre of mass that remains unknown, but within a prescribed volume, at launch allows for increased flexibility in accommodating payloads. This flexibility is particularly beneficial if there are multiple payload parts that may have specific packaging requirements or irregular shapes. Similarly, endeavouring to make the prescribed volume for the centre of gravity as large as possible maximises the payload accommodation flexibility. 
         [0004]    The current state of the art either uses an array of separation springs (e.g. the commercially available Lightband™) that can induce a significant tumble rate if the center of mass is spaced from the ejection mechanism geometric center, or a guide rail system (i.e. Pico-Satellite Orbital Deployer PPOD) for very small payloads (nano-sats) that does not scale well to larger payloads—in excess of tens of kilograms up to a few hundred kilograms—and, further, would be at risk of jamming or binding upon release. 
         [0005]    Existing ejection methods are unable to eject a payload with an offset center of mass without causing the payload to tumble. This is a result of the ejection technique; many existing methods exert a force or forces that are on, or average to, the geometric center of the ejection device. If the center of mass of the payload is offset from this geometric center of the ejection device, the payload will tumble. A common technique in the industry is to use springs to eject a payload. If the payload center of mass is offset from the geometric center, the force upon the springs is not evenly distributed. This results in the payload tumbling when the springs are released. 
       SUMMARY 
       [0006]    The present disclosure provides a system and method of ejecting a payload from a host spacecraft in a microgravity environment. The system and method does not require the payload have a geometrically centralized center of mass. It also minimizes the tumble rate of the ejected payload while being insensitive to the location of the centre of mass of that payload. 
         [0007]    There is disclosed herein a payload ejection system (PES) device able to store any set of payloads for launch and eject that set of payloads at a controlled speed with a low tumble rate while accommodating any offset centre of mass within a restricted volume. The need for ballasting or balancing is eliminated thus freeing up the design space for these payloads. Insensitivity to centre of mass location is enabled by the use of a deployment hinge assembly arrangement which uses two or more non-parallel folding hinge arrangements that allow for linear motion of the output link in one direction while restricting the motion all other directions. One non-limiting embodiment of the PES disclosed herein uses four (4) hinge panel assemblies, selected to provide optimal stiffness around the entire mechanism. However it will be appreciated that, depending on the mass of payload being ejected, a PES with as few as two non-parallel folding hinge arrangements may be used and for larger mass payloads the number of hinge arrangements may be scaled up, to three (3), four (4), five (5), to as many as needed for the particular payload size and mass. The stiffness of the PES is integral to managing offset centre-of-mass locations by allowing the mechanism to translate the effective force vector to the center of mass location. In other words, the payload ejection mechanism has a preselected stiffness to translate an effective force vector generated by the at least two deployment hinge assemblies to a center of mass location of the payload assembly. 
         [0008]    Thus, there is disclosed herein a system for hosting and controllably ejecting a payload from a host spacecraft. The device comprises: a base plate attached to the host spacecraft, a payload ejection mechanism, a plurality of ejection devices, such as springs to store ejection energy to enable the ejection velocity of the payload, and a plurality of hinge panel assemblies attached to both the base plate and the payload release plate and arranged such that at least two of the hinge panel assemblies have hinge axes that are non-parallel so the payload release plate can only propagate parallel to the base plate. The payload, or payloads, are placed on the payload chassis and are releasably secured to the host spacecraft. Ejection occurs when a retention device(s) is released and the payload ejection mechanism is free to enable ejection, either actively (a commanded device with powered actuators) or passively (a stored energy device such as springs). The deployment hinge assemblies are deployed, guiding and accelerating the payload release plate in the ejection direction until the ejection speed is reached and the deployment hinge assemblies reach their full stroke, at which point the payload and attached payload chassis becomes completely separated from the ejection system. 
         [0009]    An embodiment of a system for hosting and controllably ejecting a payload from a host spacecraft, comprises a payload ejection mechanism attached to the host spacecraft, with the payload ejection mechanism including a base plate attached to the host spacecraft and a payload release plate. A payload assembly is releasibly attached to the payload release plate and a payload is attachable to the payload assembly. The payload ejection mechanism is configured to eject the payload assembly away from the host spacecraft such that forces acting upon the payload assembly are substantially equal across the interface between the payload assembly and the payload release plate at a moment of release of payload assembly from the payload ejection mechanism regardless of a location of the payload center of mass of the payload with respect to a geometric center of the payload ejection mechanism. 
         [0010]    A further understanding of the functional and advantageous aspects of the disclosure can be realized by reference to the following detailed description and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0011]    Embodiments will now be described, by way of example only, with reference to the drawings, in which: 
           [0012]      FIG. 1  is an underneath oblique view of the payload ejection system (PES)  100  in the stowed configuration. It shows the relative positions of the payload assembly  200 , mechanical mounting assembly  400 , with the base plate  430  and the attached launch lock assemblies  410  and release mechanism  460 . 
           [0013]      FIG. 2  is an end view, taken along arrow  2  in  FIG. 1 , of the payload ejection system  100  in the stowed configuration illustrating where the sections for subsequent figures are taken. 
           [0014]      FIG. 3  is an isometric view of the payload ejection system  100  in the deployed configuration. It shows the relative positions of the payload assembly  200  and the payload ejection mechanism (PEM)  300  which comprises the mechanical mounting assembly  400 , the deployment hinge assemblies  500  and the payload release plate assembly  600 . 
           [0015]      FIG. 4  is an end view of the payload ejection system  100  along the arrow  4  in  FIG. 3 , in the deployed configuration illustrating where the sections for subsequent figures are taken. 
           [0016]      FIG. 5  shows an isometric view of possible locations for the payload ejection system  100  to reside on the hosting spacecraft  700 ,such as an unused battery bay  720  or an unused portion of an outer surface  710 . 
           [0017]      FIGS. 6   a  and  6   b  illustrate the rotation or tumble  770  caused by a lateral displacement  790  of the payload  800  centre of mass  750  from the centre of an applied ejection force  760  or from the geometric centre  785  of a plurality of ejection springs  765 . 
           [0018]      FIG. 7  shows an isometric view of a simplified linkage used in the present system using two (2) pairs of hinge plates, with simplified depictions of the upper hinge plate  501 , lower hinge plate  502 , upper hinge pin  503 , mid hinge pin  504 , lower hinge plate  505 , base plate  430  and payload release plate  610 . 
           [0019]      FIG. 8  is a section along the line  8 - 8  in  FIG. 2  illustrating the relative positions of the major assemblies in the stowed configuration including the payload  800 , payload attachment features  225 , payload chassis  210 , mounting plate  430 , PES to host connectors  440 , launch lock assemblies  410 , release mechanism  460 , and deployment hinge assemblies  500 . 
           [0020]      FIG. 9  is a section along the line  9 - 9  in  FIG. 4  illustrating the relative positions of the major assemblies in the deployed configuration including the payload  800 , payload assembly  200 , payload attachment features  225 , payload chassis  210 , payload electrical box  240 , mounting plate  430 , PES to host connectors  440 , launch lock assemblies  410 , release mechanism  460 , lock mounting plates  413 , payload release plate  610 , snubber arms  650 , snubbers  651  and deployment hinge assemblies  500  with upper hinge plates  501 , lower hinge plates  502 , upper hinge pins  503 , mid hinge pins  504 , lower hinge pins  505 , deployments springs  508 , and snubber shafts  540 . 
           [0021]      FIG. 10  is a view along arrows  10  in  FIG. 9  and shows an underside view of the payload assembly  200  with the payload chassis  210 , retaining bolts  414 , payload contacts  221 , connector alignment features  220 , payload electrical connectors  230 , and payload electrical box  240 . 
           [0022]      FIG. 11  is a view along arrows  11  in  FIG. 9  and shows an overhead view of the mechanical mounting assembly  400  with the deployment hinge assemblies  500  and the payload release plate assembly  600  omitted for clarity. It shows the relative positions of the base plate  430 , lock mounting plates  413 , release plate contact  433 , connector alignment pins  434 , payload to PEM connectors  435 , circuit board  436  and PES to host connectors  440   a  and  440   b . In this embodiment a system of electrical redundancy has been instituted for reliability resulting in two equal and separate electrical circuits for each task. In addition, the payload harness  432  connecting the payload to the host  700  through the PES to host connector  440   a  is kept separate from the launch lock harness  431  that connects the various mechanisms and sensors on the Payload Ejection system  100  to the host spacecraft  700  through PES to host connector  440   b.    
           [0023]      FIG. 12  is an oblique view of one of the deployment hinge assemblies  500  in the deployed configuration. It shows the relative positions of the base plate  430 , with the release plate contact  433  and lock mounting plate  413  and the lower hinge bracket  507 . Mounted to the lower hinge bracket  507  are the upper hinge plate  501 , lower hinge plate  502 , upper hinge pin  503 , mid hinge pin  504 , lower hinge pin  505 , upper hinge bracket  506  as part of the payload release plate  610 , the deployment springs  508 , upper hinge bearings  509 , mid hinge bearings  510 , lower hinge bearings  511  and the snubber shaft  540 . The payload release plate  610  also includes the snubber arms  650  and snubbers  651 . 
           [0024]      FIG. 13  is a section view along line  9 - 9  of  FIG. 4  of one of the deployment hinge assemblies  500  in the deployed configuration. It shows the relative positions of the base plate  430 , and lock mounting plate  413  and the lower hinge bracket  507 . Mounted to the lower hinge bracket  507  are the upper hinge plate  501 , lower hinge plate  502 , upper hinge pin  503 , mid hinge pin  504 , lower hinge pin  505 , upper hinge bracket  506  as part of the payload release plate  610 , the deployment springs  508 , upper hinge bearings  509 , mid hinge bearings  510  lower hinge bearings  511  and the snubber shaft  540 . The payload release plate  610  also includes the snubber arms  650  and snubbers  651 . 
           [0025]      FIG. 14  is a section view along line  8 - 8  of  FIG. 2  of one of the deployment hinge assemblies  500  in the stowed configuration. It shows the relative positions of the payload chassis  210 , payload mounting features  225 , launch lock assembly  410 , lock control harness  412 , base plate  430 , and the lower hinge bracket  507 . Mounted to the lower hinge bracket  507  are the upper hinge plate  501 , lower hinge plate  502 , upper hinge pin  503 , mid hinge pin  504 , lower hinge pin  505 , upper hinge bracket  506  as part of the payload release plate  610 , the deployment springs  508 , upper hinge bearings  509 , lower hinge bearings  511  and the snubber shaft  540 . 
           [0026]      FIG. 15  is an oblique section view along line  15 - 15  of  FIG. 2  of one of the deployment hinge assemblies  500  in the stowed configuration with the payload assembly  200  omitted for clarity. It shows how the snubbers  651  engage the snubber shaft  540  and the relative positions of the lock mounting plate  413 , base plate  430 , upper hinge plate  501 , lower hinge plate  502 , upper hinge pin  503 , mid hinge pin  504 , the deployment springs  508 , mid hinge bearings  510 , payload release plate  610  and snubber arms  650 . 
           [0027]      FIG. 16  is a partial section view along line  16 - 16  of  FIG. 2  of one of the launch lock assemblies  410  in the stowed configuration. It shows the relative positions of the payload chassis  210 ,launch lock assembly  410 , lock release mechanism  411 , lock control harness  412 , lock mounting plate  413 , retaining bolt  414 , retraction spring  415  in compressed configuration, lock bolt housing  416 , spring housing  417 , load cell  418 , load cell housing  419 , load cell harness  420 , base plate  430  and launchlock harness  431 . 
           [0028]      FIG. 17  is a partial section view along line  17 - 17  of  FIG. 4  of one of the launch lock assemblies  410  in the deployed configuration. It shows the relative positions of the lock release mechanism  411 , lock control harness  412 , lock mounting plate  413 , base plate  430  and launch lock harness  431 . 
           [0029]      FIG. 18  is a partial section view along line  18 - 18  of  FIG. 4  of one of the launch lock assemblies  410  in the deployed configuration. It shows the relative positions of the payload chassis  210 , retaining bolt  414 , retraction spring  415  in extended configuration, lock bolt housing  416 , spring housing  417 , load cell  418 , load cell housing  419  and load cell harness  420 . 
           [0030]      FIG. 19  is a partial section view along line  19 - 19  of  FIG. 2  of one of the deployment load paths in the stowed configuration. It shows the relative positions of the payload chassis  210 , payload contact  221 , payload release plate  610 , release plate contact  433  and base plate  430 . 
           [0031]      FIG. 20  is a partial section view along line  8 - 8  of  FIG. 2  showing the release mechanism  460  in the stowed configuration. It shows the relative positions of the release mechanism  460 , release shaft  461 , release nut  462 , washer  463 , mounting plate  464 , lock control harness  412 , base plate  430 , payload to PEM connectors  435 , circuit board  436 , PEM harness sockets  437 , payload electrical connector  230 , payload harness pins  231  and payload electrical box  240 . 
           [0032]      FIG. 21  is a partial section view along line  9 - 9  of  FIG. 4  of the payload side of the release mechanism  460  in the deployed configuration. It shows the relative positions of the payload chassis  210 , release shaft head  461   a , release nut  462 , washer  463 , payload electrical connector  230 , payload harness pins  231  and payload electrical box  240 . 
           [0033]      FIG. 22  is a partial section view along line  9 - 9  of  FIG. 4  of the host side of the release mechanism  460  in the deployed configuration. It shows the relative positions of the base plate  430 , release mechanism  460 , release shaft stub  461   b , mounting plate  464 , lock control harness  412 , payload to PEM connectors  435 , circuit board  436  and PEM harness sockets  437 . 
           [0034]      FIG. 23  is an oblique view of the host-side electrical connections between the PEM 300  and the payload assembly  200  showing the payload harness  432 , connector alignment pins  434 , payload to PEM connectors  435  and circuit board  436 . 
           [0035]      FIG. 24  is an oblique view of the payload-side electrical connections between the PEM  300  and the payload assembly  200  showing the payload electrical connectors  230 , connector alignment features  220  and the payload electrical box  240 . 
           [0036]      FIGS. 25   a ,  25   b ,  25   c ,  25   d , and  25   e  are a sequence of partial section views along line  8 - 8  of  FIG. 2  showing the deployment sequence as the deployment hinge assemblies  500  accelerate the payload release plate assembly  600  and payload assembly  200  away from the mechanical mounting assembly  400  until the deployment hinge assemblies  500  reach the end of their travel and the payload assembly  200  separates from the payload release plate assembly  600  and continues onward under its own inertia. 
           [0037]      FIG. 25   a  shows the mechanism ready to be released. The launch locks  410  are released, the deployment springs  508  are held in place by the retaining action of the release mechanism  460 . 
           [0038]      FIG. 25   b  shows that the mechanism has been activated. The release mechanism  460  has activated and the deployment springs  508 , no longer restrained, are causing the upper hinge panel  501  and lower hinge panel  502  to rotate. This causes the payload release plate assembly  600  to accelerate away from the mechanical mounting assembly  400 . At this point, release plate assembly  600  and the payload assembly  200  are held together only by the acceleration of the mechanism. 
           [0039]      FIG. 25   c  shows that the mechanism continues to accelerate the payload release plate assembly  600  away from the mechanical mounting assembly  400 . 
           [0040]      FIG. 25   d  shows the mechanism at the instant the deployment hinge assemblies  500  are at their full extension and have come to a complete stop. 
           [0041]      FIG. 25   e  shows the mechanism after the deployment hinge assemblies  500  are at their full extension and have come to a complete stop and the payload assembly  200  has separated from the deployment plate assembly  600  and is free to move under its own inertia. 
           [0042]      FIG. 26  is a block diagram showing host spacecraft  700  with a communication antenna  701  for communicating with Earth  703  with spacecraft  700  having payload ejection system  100  mounted thereon in compartments such as unused battery bays  720  or an unused exterior surface area  710  of the spacecraft  700 . 
           [0043]      FIG. 27  is a block diagram of a non-limiting exemplary computer system coupled to the payload ejection system  100  which contains a central processor  1210  interfaced with a memory storage device  1220 , input/output devices and interfaces  1230 , a power supply  1260 , an internal storage  1240  and a communications interface  1250 . 
       
    
    
     DETAILED DESCRIPTION 
       [0044]    Various embodiments and aspects of the disclosure will be described with reference to details discussed below. The following description and drawings are illustrative of the disclosure and are not to be construed as limiting the disclosure. Numerous specific details are described to provide a thorough understanding of various embodiments of the present disclosure. However, in certain instances, well-known or conventional details are not described in order to provide a concise discussion of embodiments of the present disclosure. 
         [0045]    As used herein, the terms, “comprises” and “comprising” are to be construed as being inclusive and open ended, and not exclusive. Specifically, when used in the specification and claims, the terms, “comprises” and “comprising” and variations thereof mean the specified features, steps or components are included. These terms are not to be interpreted to exclude the presence of other features, steps or components. 
         [0046]    As used herein, the term “exemplary” means “serving as an example, instance, or illustration,” and should not be construed as preferred or advantageous over other configurations disclosed herein. 
         [0047]    As used herein, the terms “about” and “approximately”, when used in conjunction with ranges of dimensions of particles, compositions of mixtures or other physical properties or characteristics, are meant to cover slight variations that may exist in the upper and lower limits of the ranges of dimensions so as to not exclude embodiments where on average most of the dimensions are satisfied but where statistically dimensions may exist outside this region. It is not the intention to exclude embodiments such as these from the present disclosure. 
         [0048]    As used herein, the term “operably connected” refers to a means of communication between two devices. This can be either a wired or non-wired communication. 
         [0049]    As used herein, the term “tumble rate” is a toppling rotational rate about any axis of a 3-axis orthogonal reference frame associated with the centre of mass of the payload or payload assembly that is detrimental to operation and/or recapture of the ejected payload. 
         [0050]    Referring to  FIG. 5 , host spacecraft  700  often have surplus mass and volume capacity on their exterior. As shown in  FIG. 5 , this can include unused battery bays  720  or an unused exterior surface area  710  of the spacecraft  700  that could be used to host a payload ejection system  100 , see  FIG. 26 . In one embodiment of this mechanism, it is proposed to use these vacant spaces to house the payload ejection system  100  and its attached payload  800 . Other embodiments can include spacecraft that are designed specifically to carry and eject a plurality of payloads as part of their primary function as opposed to carrying payloads in addition to their primary function. 
         [0051]    As mentioned above, some existing methods of ejecting a payload from a host spacecraft  700  in a microgravity environment such as orbit, apply the ejection force along a single vector and as such any displacement of the centre of mass of the payload from the vector of the ejection force produces a moment that is directly related to the distance between the centre of mass and the vector and mass of the payload. 
         [0052]    As illustrated in  FIG. 6   a , the ejection force  760  is applied along the direction  780 . The centre of mass  750  of the payload  800  is offset some distance  790  from the direction  780 . This combination of force at a distance produces a moment or couple  770  that causes the payload  800  to rotate or tumble. The ejection mechanism itself cannot correct this effect and it requires that the payload either be manufactured with very strict control of the location of the centre of mass, frequently compromising aspects of the payload, or the payload itself must expend resources to correct the tumble. 
         [0053]    Similarly, other payload ejection methods rely upon the action of a plurality of springs  765  spread over a known area to provide the ejection force as shown in  FIG. 6   b . In this case any distance between the centre of mass  750  and the geometric centre of the group of springs  795  means that springs closer to the centre of mass  750  exert their force against proportionally more of the payload  800  mass. This again causes a moment or couple as the springs further from the centre of mass extend faster and impart a rotation or tumble  770  to the payload  800 . And, again, the ejection mechanism itself cannot correct this effect with the same deleterious impacts on the payload. 
         [0054]    There are several methods to mitigate the tumbling effect of an ejected payload. These include: ballasting the payload to collocate the centre of mass with the ejection force vector, and guiding the payload. Ballasting the payload is mass and volume expensive and requires accurate and specific knowledge of the mass properties of the payload. It must also be done uniquely for each payload decreasing operational flexibility. Guiding the payload through the entire acceleration to the ejection speed, as in the case of a PPOD, requires guides. Linear guides are prone to jamming or binding as the payload approaches the end of the guides and the effective engagement of the guides is reduced to zero. 
         [0055]    The present payload ejection system uses a plurality of deployment hinge assemblies  500  ( FIG. 7 ) to eject a payload with a negligible amount of induced rotational rate or tumble even though the centre of mass of the payload is, or may be significantly distant from the overall ejection force vector. A key to this mechanism is the use of two or more system linked hinges that produce parallel motion of one plane versus another. The payload ejection mechanism  300  uses at least two pairs of hinges connected to two parallel planes and placed at an angle to each other thus constraining the possible motion of the two planes relative to each other to be parallel. 
         [0056]    The present system can be readily scaled up to handle larger payloads by using more than two deployment hinge assemblies  500 . The payload ejection system  100  disclosed herein and illustrated has four (4) deployment hinge assemblies  500  but for larger payloads five, six, seven and larger numbers of deployment hinge assemblies  500  may be used. Because the two hinges are at an angle they effectively describe a series of parallel planes at each of the upper, mid and lower hinge axes constraining the base plate  430  and the payload release plate  610  to remain parallel even in the presence of variations in the centre of mass of the payload with respect to the geometric centre of the payload release plate  610 . 
         [0057]    More specifically,  FIG. 7  shows a simplified diagram of a deployment hinge assembly  500  used in the payload ejection system  100 . To minimise torsional effects and reduce the required stiffness of the deployment hinge assembly  500  the payload ejection system  100  uses a pair of opposed linkages with each pair consisting of two upper hinge plates  501  and two lower hinge plates  502  arranged orthogonally to each other. The mechanism in the figure shows each linkage hinge to bend outward about the mid hinge pin  504 , however to make the mechanism more compact the current embodiment has one pair that bends inwards and another that bends outwards. The direction of the hinge action has no bearing on the effectiveness of the mechanism other than compactness and reduced mass. 
         [0058]    For launch and any powered transit in the stowed configuration shown in  FIG. 1  to the ejection site, the payload assembly  200  is secured to the base plate  430  of the payload ejection mechanism  300  by one or more launch lock assemblies  410 . The payload ejection mechanism  300  is in the stowed configuration and the deployment springs  508  ( FIG. 9 ) are stowed in their maximum stored energy state. 
         [0059]    When it is decided to eject the payload the launch lock assemblies  410  are commanded to release and then the payload ejection mechanism  300  and the deployment springs  508  are held by the release mechanism  460 . At the appropriate time, the release mechanism  460  is commanded to release and when it does, the stored energy in the deployment springs  508  starts to force the upper hinge panel  501  and lower hinge panel  502  to straighten up. The connector alignment pins  434  ensure that the payload electrical connector  230  slides cleanly out of engagement with the payload to PEM connector  434  before coming out of contact with the connector alignment features  220  themselves. 
         [0060]    The actions of the pair of deployment hinge assemblies  500  drive the payload release plate  610  away from the base plate  430  at a rate determined by the spring forces, the mechanism frictional drag and the mass of the payload and with the payload release plate  610  remaining parallel to the base plate  430 . 
         [0061]    At the end of the travel of the deployment hinge assemblies  500  as shown in  FIG. 12 , the upper hinge plate  501  and the lower hinge plate  502  come into contact when the upper hard stop  530  contacts the lower hard stop  531 . The deployment spring  508  force then drops to zero and the payload release plate  610  advances no further. The payload assembly  200  and the attached payload  800  are not physically attached to the release plate assembly  600 , but payload assembly  200  is adjacent to payload release plate  610  in physical contact to form an interface between them but not in any way fixed to the payload release plate  610  so that payload assembly  200  simply experiences the uni-axial ejection force created by the deploying mechanism. At the point that the deployment hinge assemblies  500  reach their hard stops  530  and  531 , the payload assembly  200  becomes free of the payload release mechanism  300  which continues on the ejection vector due to its own inertia, where its motion is perpendicular to the payload ejection plate  610  at time of release. 
         [0062]    The mechanism will now be described in more detail with reference to the figures. 
         [0063]    At any time after the launch of the host spacecraft  700  and prior to the time it is desired to eject the payload  800  and payload assembly  200  the computer control system  1200  either determines through internal programming or is commanded by a signal  702  from earth  703  to initiate the payload ejection sequence. Prior to the issuance of the command to eject the payload being given by the computer control system  1200 , the payload ejection system  100  is in the stowed configuration as shown in  FIGS. 1 ,  2 ,  8  and  11 . 
         [0064]    In this configuration, any power or data required by the payload is passed from the host spacecraft  700  through the PES to host connectors  440   a , the payload harness  432 , the circuit board  436  to the PEM harness sockets  437  held by the payload to PEM connectors  435 . The power and data then crosses to the payload assembly  200  via the payload harness pins  231  held by the payload electrical connectors  230 . A harness connects the payload harness pins  231  to the payload  800  and the payload assembly  200 . This harness is not shown because it is specific to each combination of payload  800  and payload assembly for each use of the payload ejection system  100 . 
         [0065]    The commands to initiate payload  800  ejection are provided to or generated by the computer control system  1200  and passed to the payload ejection mechanism  300  via PES to host connectors  440   b  and launch lock harness  431 . The launch lock harness  431  provides a means to provide power and data connectivity to the launch lock assemblies  410  and the release mechanism  460  and any sensors (not present in this embodiment) that may required for the operation and monitoring of the payload ejection mechanism  300 . 
         [0066]    Upon the command to operate the launch lock assemblies  410  and referring to  FIGS. 16 ,  17  and  18  the signal and power from the launch lock harness  431  passes to each the lock control harnesses  412 . In this embodiment, the launch lock assemblies  410  are commercially available separable nut devices. Upon command, the lock release mechanism  411  causes the nut within the lock release mechanism  411  to separate releasing the retaining bolt  414 . The retraction spring  415  is also released and moves the spring housing  417  and the retaining bolt  414  away from the base plate  430  and up into the lock bolt housing  416 , preventing the retaining bolt from causing the payload ejection system  100  from binding or fouling. 
         [0067]    Referring to  FIGS. 20 ,  21  and  22 , prior to initiation, the payload ejection system  100  is held together by that action of the release mechanism  460  that prevents the deployment springs  508  from ejecting the payload  800 . At the appropriate time, as determined by programming within the central computer system  1200  (see  FIG. 27 ) or passed to the central computer system  1200  from earth  703  by signals  702  to the host satellite  700 . see  FIG. 26 . The ejection command from the central computer system  1200  is passed to the payload ejection mechanism  300  via the PES  100  to host connectors  440   b  (see  FIG. 11 ) and the launch lock harness  432  which connects to the release mechanism  460 . 
         [0068]    In this embodiment, the release mechanism  460  is a commercially available frangible bolt device. Upon command the release mechanism  460  causes the release shaft  461  to fracture in a precise manner leaving the bulk of the release shaft  461   b  within the release mechanism  460  attached to the mounting plate  464  and then to the base plate  430 . The remaining portion of the release shaft  461   a  remains attached to the release nut and attached to the payload assembly  200  during the ejection sequence. 
         [0069]    The deployment hinge assemblies  500  (described in detail below) push the payload assembly  200  away from the payload ejection mechanism  300 . Referring to  FIGS. 20 ,  21 ,  22 ,  23  and  24  in order to provide a clean release of the release mechanism  460 , payload to PEM connectors  435  and circuit board  436  are fixed to mounting plate  464  which is attached to base plate  430  in such a way to permit limited movement in the plane of the base plate  430  and perpendicular to that plane. This movement removes any stresses on the release mechanism  460  and electrical connectors  230  and  435  that might prevent them from disengaging easily. To further guide the disengagement of the connectors  230  and  435  during ejection, the alignment of the payload electrical connectors  230  to the payload to PEM connectors  435  is maintained by the connector alignment pins  434  that are mounted releasibly within the connector alignment features  220  that form a part of the payload electrical box  240 . By the combined action of close manufacturing tolerances and lubricated surfaces the connector alignment pins  434  slide easily within the connector alignment features  220  and yet restrain unwanted movement between the payload electrical connectors  230  and the payload to PEM connectors  435 . Upon ejection, as the payload assembly  200  moves away from the payload ejection mechanism, the payload harness pins  231  that are part of the payload electrical connectors  230  slide out of engagement of the PEM harness sockets  437  that are part of the PEM connectors  435  while the connector alignment pins  434  are still engaged within the connector alignment features  220 . After the payload harness pins  231  have completely moved out of engagement with the PEM harness sockets  437  the connector alignment pins  434  then move out of engagement with the connector alignment features  220 . 
         [0070]    The deployment hinge assemblies  500  provide the force that enables the ejection of the payload  800  and payload assembly  200 . Referring to  FIGS. 12 ,  13 ,  14  and  15  the deployment hinge assemblies  500  work in the following manner. As explained above, when the release mechanism  460  ( FIG. 22 ) is activated the payload release plate  610  is then free to be acted upon by the deployment hinge assemblies  500 . Specifically, the deployment springs  508  are configured to act upon the upper hinge plate  501  and the lower hinge plate  502  in such a way as to force them from the collapsed or stowed configuration ( FIG. 14 ) to the extended or deployed configuration ( FIG. 12 ). The configuration of the deployment hinge assemblies  500 , specifically the use of a system of two or more linked hinge pairs produces parallel motion of one plane versus another. The deployment hinge assemblies  500  use at least two sets of hinges connected to two parallel planes, the base plate  430  and the payload release plate  610 , and placed at an angle to each other thus constraining the possible motion of the two planes to be parallel. A preferred embodiment of the payload ejection system disclosed herein has four (4) deployment hinge assemblies  500  and any pair of adjacent, non-parallel deployment hinge assemblies  500  are sufficient to constrain the motion of the payload release plate  610  to be parallel to the base plate  430 , however the use of additional deployment hinge assemblies  500  reduces the torsional loads within the mechanism and reduces the required stiffness of the deployment hinge assemblies  500  advantageously reducing the mechanism mass and increasing reliability. 
         [0071]    As the deployment springs act upon the upper hinge plate  501  and the lower hinge plate  502  they rotate about the mid hinge pins  504  which then causes the upper hinge plate  501  to rotate around the upper hinge pin  503  and the lower hinge plate  502  to rotate around the lower hinge pin  505 . The physical arrangement of one deployment hinge assembly  500  in relationship to any adjacent deployment hinge assembly  500 , is characterized by the two deployment hinge assemblies  500  being attached to the payload deployment plate  610  and the base plate  430  such that
       a) all of the upper hinge pins  503  are in one plane,   b) all of the mid hinge pins  504  are in a second plane and   c) all of the lower hinge pins  505  are in a third plane and that   d) the axes of all of these hinge pins ( 503 ,  504  and  505 ) form a non-zero angle (in this case they are orthogonal) with those of the adjacent deployment hinge assembly  500 .       
 
         [0076]    This means that the minimum two adjacent hinge assemblies effectively describe a series of parallel planes at each of the upper, mid and lower hinge axes, preventing the base plate  430  or the payload release plate  610  from being pushed out of parallel as the deployment springs  508  act to extend the individual deployment hinge assemblies  500 . This constrained motion is what forces the payload release plate  610  to move in a manner parallel to the base plate  430  when (referring to  FIG. 6   a  or  6   b ) even when the center of mass  750  of the payload  800  is a significant distance  790  from the total ejection force vector  760  as applied by the deployment hinge assemblies  500 . 
         [0077]    As the deployment hinge assemblies  500  reach their desired limit of travel (refer to  FIGS. 13 and 14 ) the upper hard stop  531 , which is a feature on the upper hinge plate  501 , comes into contact with the lower hard stop  532  which is a feature on the lower hinge plate  502 , and the extension of that deployment hinge assembly  500  comes to a stop. Due to the arrangement of angularly arranged deployment hinge assemblies  500 , each deployment hinge assembly  500  will come to its end of travel at substantially the same time therefore ending the ejection acceleration of the payload release plate  610  away from the base plate  430 . 
         [0078]    Referring to  FIG. 19 , in the stowed configuration there is no direct loading between the payload contacts  221 , the payload release plate  610  and the release plate contact  433 . Operation vibrations and loads may cause some contact between all three components and the release plate contact  433  is designed to restrict any excessive movement between the payload release plate  610  yet remaining free of the payload release plate  610  in nominal conditions. Upon release mechanism  460  activation, as the deployment hinge assemblies  500  act to push the payload release plate  610  away from the base plate  430 , the payload release plate  610  now comes into firm contact with the payload contacts  221  at four places. The acceleration of the payload assembly  200  provided by the actions of the deployment hinge assemblies  500  provides a force that keeps payload contacts  221  on the payload assembly  200  in controlled contact with the payload release plate  610  during the ejection sequence. When the deployment hinge assemblies  500  have reached their full range of motion and no longer provide an acceleration, then the payload contacts  221  simply move away from the payload release plate  610  and the payload assembly  200  and payload  800  are then independent of the host satellite  700 . 
         [0079]    It should be emphasised that the current payload ejection system  100  does not require an additional latch device between the payload release plate  610  and the payload assembly  200  which would have to be timed to release at or just before full extension of the PEM hinge assemblies  500 . This lack of a need for additional latches is enabled by the deployment hinge assemblies  500  providing the uni-axial ejection force and is predicated on the center of mass of the payload  800  and the payload assembly  200  lying within the rectangle formed by the four payload contacts  221 . 
         [0080]    As described above, the connection of the PEM  300  to the payload assembly  200  once the final release mechanism has been released is between the payload release plate  610  and the payload contacts  221 . This connection is a ‘push-contact’ only. This is chosen to ensure that once the ejection event has begun there is no risk that separation would not occur. This then requires that the center of mass of the payload assembly  200  is within the area contained by the payload contacts  221  on the payload assembly  200  and payload release plate  610 . This applies to all embodiments disclosed herein. Otherwise a tipping effect would occur regardless of the parallel motion provided by the ejection linkage. It is possible to add a latch feature that would prevent this tipping if the center of mass was outside of this contact pattern, but the release of the latches would need to be timed so as not to interfere with the payload assembly  200  at the moment of separation from the PEM  300 . 
         [0081]      FIG. 8  shows the payload ejection system  100  in its stowed configuration. The payload  800  can be virtually anything compatible with the space environment. This includes, but is not limited to small satellites, satellite subcomponents, space system consumables such as propellant or tools, components for the construction or maintenance of space systems, etc. The payload  800  can also be a unitary item or an aggregate of items fastened individually to the payload chassis  210  using the payload attachment features  225 . The payload attachment features  225  are simple threaded holes in this embodiment, however, depending upon the mission or the payload these features may also be a plurality of passive or active (motorized) attachment mechanisms each of which facilitates the mechanical attachment of the payload(s)  800  plus providing access to power, data and heat from the host  700  via cable harnesses that originate in the host  700  and pass to the payload via the payload to host connectors  440   b , the payload harness  432 , the circuit board  436 , the payload to PEM connectors  435 , the payload electrical connectors  230  and the mission specific harness(es) that would lead from the payload electrical connectors  230  to the payload  800 . This is not shown as it would be unique to each payload. 
         [0082]    In order to ensure that the mechanism does not bind during activation, several features have been incorporated in the payload ejection system  100 . Referring to  FIG. 12 , which shows the general arrangement of the deployment hinge assemblies  500 , the combination of deployment force applied by the deployment springs  508  coupled with maximum offset distance  790  the payload  800  centre of gravity  750  can be from the geometric centre of the payload ejection mechanism  300  creates a moment or couple  770  that must be resisted by the deployment hinge assembly  500 . Through the choice of manufacturing tolerances and the stiffnesses of the hinge plate  501  and  502  and hinge bearing  509 ,  510  and  511 , the inevitable flexing that happens within the mechanism can be accommodated while minimising system mass and maximising the payload offset distance  790 , thereby maximising the system&#39;s utility. 
         [0083]    Referring to  FIG. 16  the launch lock assemblies  410  are configured to minimise the chances of the lock release mechanism  411  failing to release the payload assembly  200  from the payload ejection mechanism  300 . In the stowed configuration the exact clamping force needed to hold the payload assembly  200  to the payload ejection mechanism  300  is established during assembly by the use of a load cell  418  as one of the clamped components. The data from the load cell can be read during assembly and the load cell harness  418  can be trimmed at that point if continuous monitoring is not needed or the harness can be integrated into the payload electrical connector  230  via the payload harness pins  231 . 
         [0084]    When activated, the lock release mechanism  411  releases the retaining bolt  414  and the retraction spring  415  pulls the retaining bolt  414  back and up into the lock bolt housing  416 , out of the way and minimising the chances of these bolts jamming the mechanism. 
         [0085]    Referring to  FIG. 20 , to ensure the electrical connectors  230  and  435  between the payload assembly  200  and the payload ejection mechanism  300  separate cleanly the release mechanism  460  is rigidly fastened to the mounting plate  464  but the mounting plate  464  has limited freedom of movement in the radial and axial directions. This permits the assembly of parts rigidly held by the release mechanism  460  to accommodate the movement of the other parts of the payload ejection mechanism  300 . This assembly of rigidly held parts includes the payload electrical box  240  with the attached payload electrical connectors  230 , payload harness pins  231 , payloadto PEM connectors  435  with the attached PEM harness sockets  437 . To further ensure alignment of the connectors  230  and  435  during separation the two connector alignment pins  434  slide within two connector alignment features  220  that are manufactured to tight tolerances to ensure binding does not occur. 
         [0086]    Referring to  FIG. 15 , when the deployment hinge assemblies  500  are collapsed in the stowed configuration there is some freedom of movement between the various elements of the mechanism. This freedom of movement can cause deleterious effects during the phases of the mission prior to the desired ejection of the payload assembly  200 . This embodiment uses a series of compliant snubbers  651  to restrict and damp out potential element movement prior to payload assembly  200  ejection. The snubbers  651  are attached to the snubber arms  650  which are attached to the payload release plate  610  and are configured such that when the payload release mechanism is in the stowed configuration, there is a nominal interference between the snubber shaft  540  and the snubbers  651 . The compliant nature of the snubbers  651  results in a spring force being applied to the snubber shafts  540  which acts to restrict the motion of the mid hinge pins  504  and thereby restricts and secures the rest of the components of the deployment hinge assemblies  500  preventing potential damage prior to the initiation of the command by the central computer system  1200 . 
         [0087]    Referring to  FIGS. 8 and 9 , the launch lock assemblies  410  are the primary structural connection between the payload assembly  200  and the payload ejection mechanism  300  that withstands the forces generated during the hosting spacecraft&#39;s launch from earth and orbital manoeuvres up to the time that payload ejection is initiated in the desired orbit. 
         [0088]    In an alternate embodiment, the release mechanism assembly  460  can be designed to be capable of bearing the launch loads entirely, such that launch lock assemblies  410  would not be necessary. In this case, the structure of the release mechanism assembly  460  would be configured to act as the primary structural load path and bear the loads generated in the plane of the base plate  430  during spacecraft launch while the lock release mechanism  411  would provide the clamping load to react the launch loads perpendicular to the base plate  430 . 
         [0089]    There are several commercially available release mechanisms which may be chosen to be used for the launch release mechanism  411  or the release mechanism  460 . The choice of mechanisms depends on the requirements for the mission. These mechanisms include frangible bolt systems, burn through mechanisms, separable nut systems and pyrotechnic systems, which will be well known by those skilled in the art. Key elements in this embodiment are that the launch release mechanisms  411  are sized to withstand the launch structural loads and the release mechanism needs to be sized only to hold back the deployment springs  508  prior to the final command to eject the payload assembly  200 . 
         [0090]    An alternate embodiment would exchange the stored energy activation of the deployment springs  508  for a powered actuator(s) that drive the hinge plates  501  and  502  to deploy. Using a powered actuator can confer a different acceleration profile to the payload assembly  200  which may be advantageous in some situations or environments. 
         [0091]    An alternate embodiment would add features to the payload assembly  200  suitable to permit the ejected payload  800  and attached payload assembly  200  to be grasped or captured by a device attached to a spacecraft for the purpose that this captured payload may be attached to or used by the capturing spacecraft. Payloads  800  where it might be desirable for them to be captured by a separate spacecraft would be payloads consisting spare parts, additional propellant, or mechanisms conferring additional features to the capturing spacecraft. It is in situations where the payload assembly  200  will be captured by another spacecraft where the greatly reduced tumble rates produced by the payload ejection system are especially advantageous. Reduced payload assembly  200  tumble rates significantly reduce the difficulty of another spacecraft capturing the ejected payload assembly  200 . 
         [0092]    Features that enhance or enable the capture of a payload assembly  200  by another satellite include, but are not limited to, things such as grapple features to enable the physical contact and capture between two spacecraft, visual or radar targets that enhance and enable manual or automated visual, LI DAR and radar tracking by the capturing spacecraft, interface mechanisms that enable the captured payload assembly  200  to be securely attached to the capturing spacecraft enabling the payload  800  to be utilised. 
         [0093]    Examples of some of the features usable for a spacecraft to capture the payload assembly  200  are those used in the Orbital Express Demonstration Mission (Ogilvie, A., Autonomous Satellite Servicing Using the Orbital Express Demonstration Manipulator System, Proceedings of the 8th International Symposium on Artificial Intelligence, Robotics and Automation in Space, iSAIRAS, Pasadena, 2008 and Ogilvie, A., Autonomous Robotic Operations for On-Orbit Satellite Servicing, Sensors and Systems for Space Applications, Proc. Of SPIE Vol 6958, 695809, 2008). 
         [0094]    The present payload ejection system may be retrofitted onto any suitable satellite to be used as a host spacecraft. The system may be under teleoperation by a remotely located operator, for example located on earth, in another spacecraft or in an orbiting space station. The system may also be autonomously controlled by a local Mission Manager with some levels of supervised autonomy so that in addition to being under pure teleoperation there may be mixed teleoperation/supervised autonomy. 
         [0095]    An alternate embodiment would add features that would permit the payload ejection mechanism  300  to be retracted after activation and change the release mechanism  460  from a single use device such as the frangible bolt devise to one that can be reset remotely. Retraction features may include, but are not limited to, cables connected to a winch and motor or a piston and lever arrangement with appropriate hasps and latches. This would allow an additional device (not shown) to place additional payloads  800  and payload assemblies  200  upon the reset payload ejection mechanism  300  so that these additional payloads  800  and payload assemblies  200  may also be ejected. This is a useful embodiment in cases where multiple payloads are being launched with one payload ejection mechanism having a first payload  800  coupled thereto but where addition payloads  800  are stored on the host satellite and can be sequentially retrieved from their stored locations and ejected once the first payload has been ejected. An autoloader mounted on the host satellite may be programmed to fetch the additional payloads and mount them on the payload deployment plate. The autoloader would be pre-programmed to release the addition payloads from their storage berths. Optionally a vision system may be positioned on the host satellite so the re-launch operations may be controlled remotely by a human operator. 
         [0096]    The specific embodiments described above have been shown by way of example, and it should be understood that these embodiments may be susceptible to various modifications and alternative forms. It should be further understood that the claims are not intended to be limited to the particular forms disclosed, but rather to cover all modifications, equivalents, and alternatives falling within the spirit and scope of this disclosure.