Patent Description:
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 "<NPL>. 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.

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.

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.

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.

<CIT> discloses a method for forming a payload separation system comprising a supporting structure, a payload releasably attached to the supporting structure using a lock, and springs between supporting structure and the payload. The method uses the location of the center of mass of the payload to calibrate the force that each spring must exert such that to torque exerted by each spring in equal and opposite to the sum of the torques of the other springs.

<CIT> discloses a payload ejection mechanism which is similar to that of <CIT> because it uses springs to eject a payload and the location of the center of mass of the payload must be calculated such that the torques from each of the spring impart zero net torque on the payload.

<CIT> discloses a device for jettisoning a payload from a spacecraft comprising: a connection means between the between the payload and the spacecraft; an energy storage device for jettisoning the payload; and a releasable latching device for latching the payload to the spacecraft. The connection means is a single rod in alignment with the axis of symmetry of the jettisoning device and having a first end fixed to the payload where the energy storage device acts on the first end of the connection rod and the latching device acts on the second end of the connection rod.

<CIT> discloses a payload ejection mechanism that includes a cylindrical core that guides a larger cylinder where the ejection of a payload is constrained by the complementary cylindrical surfaces which slide off each other as the payload is ejected. This patent discloses at least two of these "complementary cylinders" being used in an embodiment.

The present disclosure provides a system according to claim <NUM> and method according to claim <NUM> 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.

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 (<NUM>) 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 (<NUM>), four (<NUM>), five (<NUM>), 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.

A further understanding of the functional and advantageous aspects of the disclosure can be realized by reference to the following detailed description and drawings.

Embodiments will now be described, by way of example only, with reference to the drawings, in which:.

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.

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.

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.

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.

As used herein, the term "tumble rate" is a toppling rotational rate about any axis of a <NUM>-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.

Referring to <FIG>, host spacecraft <NUM> often have surplus mass and volume capacity on their exterior. As shown in <FIG>, this can include unused battery bays <NUM> or an unused exterior surface area <NUM> of the spacecraft <NUM> that could be used to host a payload ejection system <NUM>, see <FIG>. In one embodiment of this mechanism, it is proposed to use these vacant spaces to house the payload ejection system <NUM> and its attached payload <NUM>. Other embodiments 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.

As mentioned above, some existing methods of ejecting a payload from a host spacecraft <NUM> 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.

As illustrated in <FIG>, the ejection force <NUM> is applied along the direction <NUM>. The centre of mass <NUM> of the payload <NUM> is offset some distance <NUM> from the direction <NUM>. This combination of force at a distance produces a moment or couple <NUM> that causes the payload <NUM> 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.

Similarly, other payload ejection methods rely upon the action of a plurality of springs <NUM> spread over a known area to provide the ejection force as shown in <FIG>. In this case any distance between the centre of mass <NUM> and the geometric centre of the group of springs <NUM> means that springs closer to the centre of mass <NUM> exert their force against proportionally more of the payload <NUM> 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 <NUM> to the payload <NUM>. And, again, the ejection mechanism itself cannot correct this effect with the same deleterious impacts on the payload.

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.

The present payload ejection system uses a plurality of deployment hinge assemblies <NUM> (<FIG>) 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 <NUM> 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.

The present system is readily scaled up to handle larger payloads by using more than two deployment hinge assemblies <NUM>. The payload ejection system <NUM> disclosed herein and illustrated has four (<NUM>) deployment hinge assemblies <NUM> but for larger payloads five, six, seven and larger numbers of deployment hinge assemblies <NUM> 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 <NUM> and the payload release plate <NUM> 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 <NUM>.

More specifically, <FIG> shows a simplified diagram of a deployment hinge assembly <NUM> used in the payload ejection system <NUM>. To minimise torsional effects and reduce the required stiffness of the deployment hinge assembly <NUM> the payload ejection system <NUM> uses a pair of opposed linkages with each pair consisting of two upper hinge plates <NUM> and two lower hinge plates <NUM> arranged orthogonally to each other. The mechanism in the figure shows each linkage hinge to bend outward about the mid hinge pin <NUM>, 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.

For launch and any powered transit in the stowed configuration shown in <FIG> to the ejection site, the payload assembly <NUM> is secured to the base plate <NUM> of the payload ejection mechanism <NUM> by one or more launch lock assemblies <NUM>. The payload ejection mechanism <NUM> is in the stowed configuration and the deployment springs <NUM> (<FIG>) are stowed in their maximum stored energy state.

When it is decided to eject the payload the launch lock assemblies <NUM> are commanded to release and then the payload ejection mechanism <NUM> and the deployment springs <NUM> are held by the release mechanism <NUM>. At the appropriate time, the release mechanism <NUM> is commanded to release and when it does, the stored energy in the deployment springs <NUM> starts to force the upper hinge panel <NUM> and lower hinge panel <NUM> to straighten up. The connector alignment pins <NUM> ensure that the payload electrical connector <NUM> slides cleanly out of engagement with the payload to PEM connector <NUM> before coming out of contact with the connector alignment features <NUM> themselves.

The actions of the pair of deployment hinge assemblies <NUM> drive the payload release plate <NUM> away from the base plate <NUM> at a rate determined by the spring forces, the mechanism frictional drag and the mass of the payload and with the payload release plate <NUM> remaining parallel to the base plate <NUM>.

At the end of the travel of the deployment hinge assemblies <NUM> as shown in <FIG>, the upper hinge plate <NUM> and the lower hinge plate <NUM> come into contact when the upper hard stop <NUM> contacts the lower hard stop <NUM>. The deployment spring <NUM> force then drops to zero and the payload release plate <NUM> advances no further. The payload assembly <NUM> and the attached payload <NUM> are not physically attached to the release plate assembly <NUM>, but payload assembly <NUM> is adjacent to payload release plate <NUM> in physical contact to form an interface between them but not in any way fixed to the payload release plate <NUM> so that payload assembly <NUM> simply experiences the uni-axial ejection force created by the deploying mechanism. At the point that the deployment hinge assemblies <NUM> reach their hard stops <NUM> and <NUM>, the payload assembly <NUM> becomes free of the payload release mechanism <NUM> which continues on the ejection vector due to its own inertia, where its motion is perpendicular to the payload ejection plate <NUM> at time of release.

The mechanism will now be described in more detail with reference to the figures.

At any time after the launch of the host spacecraft <NUM> and prior to the time it is desired to eject the payload <NUM> and payload assembly <NUM> the computer control system <NUM> either determines through internal programming or is commanded by a signal <NUM> from earth <NUM> to initiate the payload ejection sequence. Prior to the issuance of the command to eject the payload being given by the computer control system <NUM>, the payload ejection system <NUM> is in the stowed configuration as shown in <FIG>, <FIG>, <FIG> and <FIG>.

In this configuration, any power or data required by the payload is passed from the host spacecraft <NUM> through the PES to host connectors 440a, the payload harness <NUM>, the circuit board <NUM> to the PEM harness sockets <NUM> held by the payload to PEM connectors <NUM>. The power and data then crosses to the payload assembly <NUM> via the payload harness pins <NUM> held by the payload electrical connectors <NUM>. A harness connects the payload harness pins <NUM> to the payload <NUM> and the payload assembly <NUM>. This harness is not shown because it is specific to each combination of payload <NUM> and payload assembly for each use of the payload ejection system <NUM>.

The commands to initiate payload <NUM> ejection are provided to or generated by the computer control system <NUM> and passed to the payload ejection mechanism <NUM> via PES to host connectors 440b and launch lock harness <NUM>. The launch lock harness <NUM> provides a means to provide power and data connectivity to the launch lock assemblies <NUM> and the release mechanism <NUM> and any sensors (not present in this embodiment) that may required for the operation and monitoring of the payload ejection mechanism <NUM>.

Upon the command to operate the launch lock assemblies <NUM> and referring to <FIG>, <FIG> and <FIG> the signal and power from the launch lock harness <NUM> passes to each the lock control harnesses <NUM>. In this embodiment, the launch lock assemblies <NUM> are commercially available separable nut devices. Upon command, the lock release mechanism <NUM> causes the nut within the lock release mechanism <NUM> to separate releasing the retaining bolt <NUM>. The retraction spring <NUM> is also released and moves the spring housing <NUM> and the retaining bolt <NUM> away from the base plate <NUM> and up into the lock bolt housing <NUM>, preventing the retaining bolt from causing the payload ejection system <NUM> from binding or fouling.

Referring to <FIG>, <FIG> and <FIG>, prior to initiation, the payload ejection system <NUM> is held together by that action of the release mechanism <NUM> that prevents the deployment springs <NUM> from ejecting the payload <NUM>. At the appropriate time, as determined by programming within the central computer system <NUM> (see <FIG>) or passed to the central computer system <NUM> from earth <NUM> by signals <NUM> to the host satellite <NUM>. The ejection command from the central computer system <NUM> is passed to the payload ejection mechanism <NUM> via the PES <NUM> to host connectors 440b (see <FIG>) and the launch lock harness <NUM> which connects to the release mechanism <NUM>.

In this embodiment, the release mechanism <NUM> is a commercially available frangible bolt device. Upon command the release mechanism <NUM> causes the release shaft <NUM> to fracture in a precise manner leaving the bulk of the release shaft 461b within the release mechanism <NUM> attached to the mounting plate <NUM> and then to the base plate <NUM>. The remaining portion of the release shaft 461a remains attached to the release nut and attached to the payload assembly <NUM> during the ejection sequence.

The deployment hinge assemblies <NUM> (described in detail below) push the payload assembly <NUM> away from the payload ejection mechanism <NUM>. Referring to <FIG>, <FIG>, <FIG>, <FIG> and <FIG> in order to provide a clean release of the release mechanism <NUM>, payload to PEM connectors <NUM> and circuit board <NUM> are fixed to mounting plate <NUM> which is attached to base plate <NUM> in such a way to permit limited movement in the plane of the base plate <NUM> and perpendicular to that plane. This movement removes any stresses on the release mechanism <NUM> and electrical connectors <NUM> and <NUM> that might prevent them from disengaging easily. To further guide the disengagement of the connectors <NUM> and <NUM> during ejection, the alignment of the payload electrical connectors <NUM> to the payload to PEM connectors <NUM> is maintained by the connector alignment pins <NUM> that are mounted releasibly within the connector alignment features <NUM> that form a part of the payload electrical box <NUM>. By the combined action of close manufacturing tolerances and lubricated surfaces the connector alignment pins <NUM> slide easily within the connector alignment features <NUM> and yet restrain unwanted movement between the payload electrical connectors <NUM> and the payload to PEM connectors <NUM>. Upon ejection, as the payload assembly <NUM> moves away from the payload ejection mechanism, the payload harness pins <NUM> that are part of the payload electrical connectors <NUM> slide out of engagement of the PEM harness sockets <NUM> that are part of the PEM connectors <NUM> while the connector alignment pins <NUM> are still engaged within the connector alignment features <NUM>. After the payload harness pins <NUM> have completely moved out of engagement with the PEM harness sockets <NUM> the connector alignment pins <NUM> then move out of engagement with the connector alignment features <NUM>.

The deployment hinge assemblies <NUM> provide the force that enables the ejection of the payload <NUM> and payload assembly <NUM>. Referring to <FIG>, <FIG>, <FIG> and <FIG> the deployment hinge assemblies <NUM> work in the following manner. As explained above, when the release mechanism <NUM> (<FIG>) is activated the payload release plate <NUM> is then free to be acted upon by the deployment hinge assemblies <NUM>. Specifically, the deployment springs <NUM> are configured to act upon the upper hinge plate <NUM> and the lower hinge plate <NUM> in such a way as to force them from the collapsed or stowed configuration (<FIG>) to the extended or deployed configuration (<FIG>). The configuration of the deployment hinge assemblies <NUM>, 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 <NUM> use at least two sets of hinges connected to two parallel planes, the base plate <NUM> and the payload release plate <NUM>, 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 (<NUM>) deployment hinge assemblies <NUM> and any pair of adjacent, non-parallel deployment hinge assemblies <NUM> are sufficient to constrain the motion of the payload release plate <NUM> to be parallel to the base plate <NUM>, however the use of additional deployment hinge assemblies <NUM> reduces the torsional loads within the mechanism and reduces the required stiffness of the deployment hinge assemblies <NUM> advantageously reducing the mechanism mass and increasing reliability.

As the deployment springs act upon the upper hinge plate <NUM> and the lower hinge plate <NUM> they rotate about the mid hinge pins <NUM> which then causes the upper hinge plate <NUM> to rotate around the upper hinge pin <NUM> and the lower hinge plate <NUM> to rotate around the lower hinge pin <NUM>. The physical arrangement of one deployment hinge assembly <NUM> in relationship to any adjacent deployment hinge assembly <NUM>, is characterized by the two deployment hinge assemblies <NUM> being attached to the payload deployment plate <NUM> and the base plate <NUM> such that.

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 <NUM> or the payload release plate <NUM> from being pushed out of parallel as the deployment springs <NUM> act to extend the individual deployment hinge assemblies <NUM>. This constrained motion is what forces the payload release plate <NUM> to move in a manner parallel to the base plate <NUM> when (referring to <FIG>) even when the center of mass <NUM> of the payload <NUM> is a significant distance <NUM> from the total ejection force vector <NUM> as applied by the deployment hinge assemblies <NUM>.

As the deployment hinge assemblies <NUM> reach their desired limit of travel (refer to <FIG> and <FIG>) the upper hard stop <NUM>, which is a feature on the upper hinge plate <NUM>, comes into contact with the lower hard stop <NUM> which is a feature on the lower hinge plate <NUM>, and the extension of that deployment hinge assembly <NUM> comes to a stop. Due to the arrangement of angularly arranged deployment hinge assemblies <NUM>, each deployment hinge assembly <NUM> will come to its end of travel at substantially the same time therefore ending the ejection acceleration of the payload release plate <NUM> away from the base plate <NUM>.

Referring to <FIG>, in the stowed configuration there is no direct loading between the payload contacts <NUM>, the payload release plate <NUM> and the release plate contact <NUM>. Operation vibrations and loads may cause some contact between all three components and the release plate contact <NUM> is designed to restrict any excessive movement between the payload release plate <NUM> yet remaining free of the payload release plate <NUM> in nominal conditions. Upon release mechanism <NUM> activation, as the deployment hinge assemblies <NUM> act to push the payload release plate <NUM> away from the base plate <NUM>, the payload release plate <NUM> now comes into firm contact with the payload contacts <NUM> at four places. The acceleration of the payload assembly <NUM> provided by the actions of the deployment hinge assemblies <NUM> provides a force that keeps payload contacts <NUM> on the payload assembly <NUM> in controlled contact with the payload release plate <NUM> during the ejection sequence. When the deployment hinge assemblies <NUM> have reached their full range of motion and no longer provide an acceleration, then the payload contacts <NUM> simply move away from the payload release plate <NUM> and the payload assembly <NUM> and payload <NUM> are then independent of the host satellite <NUM>.

It should be emphasised that the current payload ejection system <NUM> does not require an additional latch device between the payload release plate <NUM> and the payload assembly <NUM> which would have to be timed to release at or just before full extension of the PEM hinge assemblies <NUM>. This lack of a need for additional latches is enabled by the deployment hinge assemblies <NUM> providing the uni-axial ejection force and is predicated on the center of mass of the payload <NUM> and the payload assembly <NUM> lying within the rectangle formed by the four payload contacts <NUM>.

As described above, the connection of the PEM <NUM> to the payload assembly <NUM> once the final release mechanism has been released is between the payload release plate <NUM> and the payload contacts <NUM>. 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 <NUM> is within the area contained by the payload contacts <NUM> on the payload assembly <NUM> and payload release plate <NUM>. 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 <NUM> at the moment of separation from the PEM <NUM>.

<FIG> shows the payload ejection system <NUM> in its stowed configuration. The payload <NUM> is 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 <NUM> can also be a unitary item or an aggregate of items fastened individually to the payload chassis <NUM> using the payload attachment features <NUM>. The payload attachment features <NUM> are simple threaded holes in this embodiment, however, depending upon the mission or the is 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) <NUM> plus providing access to power, data and heat from the host <NUM> via cable harnesses that originate in the host <NUM> and pass to the payload via the payload to host connectors 440b, the payload harness <NUM>, the circuit board <NUM>, the payload to PEM connectors <NUM>, the payload electrical connectors <NUM> and the mission specific harness(es) that would lead from the payload electrical connectors <NUM> to the payload <NUM>. This is not shown as it would be unique to each payload.

In order to ensure that the mechanism does not bind during activation, several features have been incorporated in the payload ejection system <NUM>. Referring to <FIG>, which shows the general arrangement of the deployment hinge assemblies <NUM>, the combination of deployment force applied by the deployment springs <NUM> coupled with maximum offset distance <NUM> the payload <NUM> centre of gravity <NUM> can be from the geometric centre of the payload ejection mechanism <NUM> creates a moment or couple <NUM> that must be resisted by the deployment hinge assembly <NUM>. Through the choice of manufacturing tolerances and the stiffnesses of the hinge plate <NUM> and <NUM> and hinge bearing <NUM>, <NUM> and <NUM>, the inevitable flexing that happens within the mechanism can be accommodated while minimising system mass and maximising the payload offset distance <NUM>, thereby maximising the system's utility.

Referring to <FIG> the launch lock assemblies <NUM> are configured to minimise the chances of the lock release mechanism <NUM> failing to release the payload assembly <NUM> from the payload ejection mechanism <NUM>. In the stowed configuration the exact clamping force needed to hold the payload assembly <NUM> to the payload ejection mechanism <NUM> is established during assembly by the use of a load cell <NUM> as one of the clamped components. The data from the load cell can be read during assembly and the load cell harness <NUM> can be trimmed at that point if continuous monitoring is not needed or the harness can be integrated into the payload electrical connector <NUM> via the payload harness pins <NUM>.

When activated, the lock release mechanism <NUM> releases the retaining bolt <NUM> and the retraction spring <NUM> pulls the retaining bolt <NUM> back and up into the lock bolt housing <NUM>, out of the way and minimising the chances of these bolts jamming the mechanism.

Referring to <FIG>, to ensure the electrical connectors <NUM> and <NUM> between the payload assembly <NUM> and the payload ejection mechanism <NUM> separate cleanly the release mechanism <NUM> is rigidly fastened to the mounting plate <NUM> but the mounting plate <NUM> has limited freedom of movement in the radial and axial directions. This permits the assembly of parts rigidly held by the release mechanism <NUM> to accommodate the movement of the other parts of the payload ejection mechanism <NUM>. This assembly of rigidly held parts includes the payload electrical box <NUM> with the attached payload electrical connectors <NUM>, payload harness pins <NUM>, payloadto PEM connectors <NUM> with the attached PEM harness sockets <NUM>. To further ensure alignment of the connectors <NUM> and <NUM> during separation the two connector alignment pins <NUM> slide within two connector alignment features <NUM> that are manufactured to tight tolerances to ensure binding does not occur.

Referring to <FIG>, when the deployment hinge assemblies <NUM> 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 <NUM>. This embodiment uses a series of compliant snubbers <NUM> to restrict and damp out potential element movement prior to payload assembly <NUM> ejection. The snubbers <NUM> are attached to the snubber arms <NUM> which are attached to the payload release plate <NUM> and are configured such that when the payload release mechanism is in the stowed configuration, there is a nominal interference between the snubber shaft <NUM> and the snubbers <NUM>. The compliant nature of the snubbers <NUM> results in a spring force being applied to the snubber shafts <NUM> which acts to restrict the motion of the mid hinge pins <NUM> and thereby restricts and secures the rest of the components of the deployment hinge assemblies <NUM> preventing potential damage prior to the initiation of the command by the central computer system <NUM>.

Referring to <FIG> and <FIG>, the launch lock assemblies <NUM> are the primary structural connection between the payload assembly <NUM> and the payload ejection mechanism <NUM> that withstands the forces generated during the hosting spacecraft's launch from earth and orbital manoeuvres up to the time that payload ejection is initiated in the desired orbit.

In an alternate embodiment, the release mechanism assembly <NUM> can be designed to be capable of bearing the launch loads entirely, such that launch lock assemblies <NUM> would not be necessary. In this case,the structure of the release mechanism assembly <NUM> would be configured to act as the primary structural load path and bear the loads generated in the plane of the base plate <NUM>during spacecraft launch while the lock release mechanism <NUM> would provide the clamping load to react the launch loads perpendicular to the base plate <NUM>.

There are several commercially available release mechanisms which may be chosen to be used for the launch release mechanism <NUM> or the release mechanism <NUM>. 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 <NUM> are sized to withstand the launch structural loads and the release mechanism needs to be sized only to hold back the deployment springs <NUM> prior to the final command to eject the payload assembly <NUM>.

An alternate embodiment would exchange the stored energy activation of the deployment springs <NUM> for a powered actuator(s) that drive the hinge plates <NUM> and <NUM> to deploy. Using a powered actuator can confer a different acceleration profile to the payload assembly <NUM> which may be advantageous in some situations or environments.

An alternate embodiment would add features to the payload assembly <NUM> suitable to permit the ejected payload <NUM> and attached payload assembly <NUM> 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 <NUM> 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 <NUM> will be captured by another spacecraft where the greatly reduced tumble rates produced by the payload ejection system are especially advantageous. Reduced payload assembly <NUM> tumble rates significantly reduce the difficulty of another spacecraft capturing the ejected payload assembly <NUM>.

Features that enhance or enable the capture of a payload assembly <NUM> 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, LIDAR and radar tracking by the capturing spacecraft, interface mechanisms that enable the captured payload assembly <NUM> to be securely attached to the capturing spacecraft enabling the payload <NUM> to be utilised.

Examples of some of the features usable for a spacecraft to capture the payload assembly <NUM> are those used in the Orbital Express Demonstration Mission (<NPL> and<NPL>).

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.

An alternate embodiment would add features that would permit the payload ejection mechanism <NUM> to be retracted after activation and change the release mechanism <NUM> from a single use device such as the frangible bolt device 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 <NUM> and payload assemblies <NUM> upon the reset payload ejection mechanism <NUM> so that these additional payloads <NUM> and payload assemblies <NUM> 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 <NUM> coupled thereto but where addition payloads <NUM> 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.

Claim 1:
A payload ejection system (<NUM>) for hosting and controllably ejecting a payload (<NUM>) from a host spacecraft (<NUM>), comprising:
a payload ejection mechanism (<NUM>) attachable to the host spacecraft (<NUM>), said payload ejection mechanism (<NUM>) including a base plate (<NUM>) attachable to said host spacecraft (<NUM>) and a payload release plate (<NUM>);
a payload assembly (<NUM>) releasibly attached to said payload release plate (<NUM>), a payload (<NUM>) being attachable to said payload assembly (<NUM>); and said payload ejection mechanism (<NUM>) characterized in that
said payload ejection mechanism (<NUM>) includes at least two deployment hinge assemblies (<NUM>), each including two hinge plates (<NUM>, <NUM>) hinged together, wherein one of the two hinge plates (<NUM>) is hinged to said base plate (<NUM>) and the other of the two hinge plates (<NUM>) is hinged to said payload release plate (<NUM>), wherein at least one of the deployment hinge assemblies (<NUM>) has hinge axes that are non-parallel to those of at least one other of the deployment hinge assemblies (<NUM>) so that the payload release plate (<NUM>) can only propagate parallel to the base plate (<NUM>), wherein, when said payload ejection mechanism (<NUM>) is activated, the one of the two hinge plates (<NUM>) hinges with respect to the base plate (<NUM>) and the other of the two hinge plates (<NUM>) hinges with respect to the payload release plate (<NUM>) to drive the payload assembly (<NUM>) away from the host spacecraft (<NUM>) along an ejection axis to eject the payload assembly (<NUM>) away from the host spacecraft (<NUM>) with forces acting upon the payload assembly (<NUM>) being substantially equal across the interface between the payload assembly (<NUM>) and the payload release plate (<NUM>) at a moment of release of the payload assembly (<NUM>) from said payload ejection mechanism (<NUM>) by maintaining controlled contact during the ejection such that the payload has no angular momentum transverse to the ejection axis or linear momentum transverse to the ejection axis at the time of release regardless of a location of the payload center of mass of the payload (<NUM>) with respect to a geometric center of said payload ejection mechanism (<NUM>) to eject said payload assembly (<NUM>) away from said host spacecraft (<NUM>) with a negligible tumble rate.