Patent Description:
Launch vehicles often employ a space vehicle payload adapter to attach multiple satellites to enable a shared launch for multiple satellites (e.g., a shared launch including a primary satellite along with several small secondary satellites) from the surface of Earth into space. Currently, conventional space vehicle payload adapters often employ a solid monocoque ring design (e.g., refer to the space vehicle payload adapter ring <NUM> of <FIG>). Since this conventional design comprises a solid ring structure, the design is able to provide for stiffness at the interfaces of the secondary payload adapters such that high frequency modes are maintained (e.g., during the duration of the launch). However, the solid ring structure of this conventional design has the disadvantages of being heavy, expensive to manufacture, and not providing for easy access to components located internal to the ring.

<CIT> describes, in accordance with its abstract, a cargo carrier for the efficient delivery of cargo to space, such as to support the International Space Station (ISS). Both pressurized and unpressurized cargo may be delivered into space on an expendable launch vehicle, such as the Delta-IV rocket. The cargo carrier may utilize a slightly modified Delta-IV second stage to provide on-orbit station keeping of the payload until it is transferred to the ISS. The cargo carrier can include an unpressurized section having a rigid central structure supporting a frame to which unpressurized cargo modules are coupled. In addition, a pressurized cargo section may be coupled to the unpressurized section. The cargo carrier may utilize existing on-orbit assets such as the European Automated Transfer Vehicle (ATV) to transfer the ISS cargo from a rendezvous orbit to the ISS.

<CIT> describes, in accordance with its abstract, a modular spacecraft incorporating a launch vehicle interface design that also functions as a berthing or docking interface for the spacecraft. The vehicle interface utilizes existing launch vehicle components and is said to to minimize cost and time to flight. In addition, the spacecraft is constructed from off-the- shelf launch vehicle components that is said to enable a low cost construction as well as a secondary payload carrying functionality.

<CIT> describes, in accordance with its abstract, a shock damping element having a closed cross section. The shock damping element comprises a first shock damping structure comprising a multitude of first elements, a multitude of second elements and a first central member. The first elements and the second elements are located at opposite sides of the first central member. The first and second elements extend towards the first central member at a first acute angle relative the first central member and are connected to the first central member. The first and second elements extend in a first direction circumferentially.

In light of the foregoing, there is a need for an improved space vehicle payload adapter design that provides for a reduction in weight and cost, and allows for an easier access to internal components, while also maintaining high frequency modes.

The present disclosure relates to a method, system, and apparatus for the direct mount of secondary payload adapters to a truss structure common to a space vehicle payload adapter. The invention to which this European patent relates is defined in the appended claims.

A method for reacting loads into a space vehicle payload adapter is provided in claim <NUM>.

In one or more embodiments, the reacting of the loads maintains high frequency (e.g., greater than (>) thirty (<NUM>) gigahertz (GHz)) modes for the space vehicle payload adapter.

In one or more embodiments, the secondary payloads are mounted onto the space vehicle payload adapter via secondary payload adapters. In some embodiments, the secondary payloads are mounted onto the secondary payload adaptors via kinematic mount bolts and/or easy ride adapters. In at least one embodiment, each of the secondary payload adapters are releasably attached to various different locations on at least one of the interstitial rings. In some embodiments, adapter port openings of the secondary payload adapters are of different sizes. In one or more embodiments, each of the adapter port openings comprises one of a circular shape, a rectangular shape, a triangular shape, or a polygon shape. In one or more embodiments, the adapter port openings comprise shapes complementary to interfaces of the secondary payloads. In some embodiments, the secondary payload adapters are manufactured from aluminum, titanium, and/or a composite material.

In the present invention the struts connect the forward ring to the aft ring. In one or more embodiments, the struts are oriented at consistent angles to form alternately inverted isosceles triangle-shaped openings within the truss structure.

In one or more embodiments, the interstitial rings are connected to the struts via a nested joint configuration. In some embodiments, the interstitial rings are located between the forward ring and the aft ring. In at least one embodiment, at least one of the interstitial rings is a partial interstitial ring. In some embodiments, at least one of the interstitial rings comprises a plurality of segments. In one or more embodiments, each of the segments comprises an inner portion and an outer portion.

In at least one embodiment, the interstitial rings, the struts, the forward ring, and/or the aft ring are manufactured from aluminum, titanium, and/or a composite material.

In an example not being covered by the appended claims a space vehicle payload adapter comprises a forward ring and an aft ring. The space vehicle payload adapter further comprises a truss structure comprising a plurality of struts, where the struts connect the forward ring to the aft ring. Also, the space vehicle payload adapter comprises more than two interstitial rings connected to the struts, and positioned between the forward ring and the aft ring. Further, the space vehicle payload adapter comprises a plurality of secondary payload adapters each releasably attached to at least one of the interstitial rings.

The features, functions, and advantages can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments.

These and other features, aspects, and advantages of the present disclosure will become better understood with regard to the following description, appended claims, and accompanying drawings where:.

The methods and apparatuses disclosed herein provide operative systems for the direct mount of secondary payload adapters to a truss structure common to a space vehicle payload adapter. In one or more embodiments, the system of the present disclosure employs a plurality of interstitial rings within a space vehicle payload adapter, which comprises a lightweight truss structure, to provide stiffness at the interfaces of the secondary payload adapters to maintain high frequency (e.g., greater than (>) thirty (<NUM>) gigahertz (GHz)) modes. The disclosed system also employs secondary payload adapters that are clockable to allow for the rotation of the secondary payloads to different locations around the circumference of the space vehicle payload adapter to optimize the center of gravity (CG) for the launch vehicle.

As previously mentioned above, launch vehicles often employ a space vehicle payload adapter to attach multiple satellites to enable a shared launch for multiple satellites (e.g., a shared launch including a primary satellite (e.g., a primary payload) along with several small secondary satellites (e.g., secondary payloads)) from the surface of Earth into space. Currently, conventional space vehicle payload adapters often employ a solid monocoque ring design (e.g., refer to the space vehicle payload adapter ring <NUM> of <FIG>). Since this conventional design comprises a solid ring structure, the design is able to provide for stiffness at the interfaces of the secondary payload adapters (refer to 110a, 110b, 110c, 110d, 110e, 110f of <FIG>) such that high frequency modes (e.g., > <NUM>) are maintained (e.g., during the duration of the launch). However, the solid ring structure of this conventional design has the disadvantages of being heavy, expensive to manufacture, and not providing for easy access to components located internal to the ring.

The system of the present disclosure provides a means to add secondary payloads to a space vehicle payload adapter, which fits between the top of a rocket (e.g., a launch vehicle) and the primary payload, that maintains the high frequency modes of a conventional monocoque ring design, while also utilizing a sparse truss structure that allows for an easier installation of the secondary payloads as well as easy access to internal components (e.g., during payload buildup and test). The disclosed space vehicle payload adapter allows for a reduction in weight and cost as compared to the conventional monocoque ring designs, while maintaining the stiffness at the interfaces of the secondary payload mounts. In particular, the disclosed space vehicle payload adapter employs multiple interstitial rings (e.g., more than two interstitial rings) that interface with the secondary payload adapters as well as with the truss structure. With this disclosed design, the interstitial rings provide a load path to react the bending loads (e.g., loads caused by the secondary payloads) into the truss structure.

In the following description, numerous details are set forth in order to provide a more thorough description of the system. It will be apparent, however, to one skilled in the art, that the disclosed system may be practiced without these specific details. In the other instances, well known features have not been described in detail, so as not to unnecessarily obscure the system.

Embodiments of the present disclosure may be described herein in terms of functional and/or logical components and various processing steps. It should be appreciated that such components may be realized by any number of hardware, software, and/or firmware components configured to perform the specified functions. In addition, those skilled in the art will appreciate that embodiments of the present disclosure may be practiced in conjunction with other components, and that the systems described herein are merely example embodiments of the present disclosure.

For the sake of brevity, conventional techniques and components related to space vehicle payload adapters, and other functional aspects of the system (and the individual operating components of the systems) may not be described in detail herein. It should be noted that many alternative or additional functional relationships or physical connections may be present in one or more embodiments of the present disclosure.

<FIG> is a diagram showing a perspective view of a conventional space vehicle payload adapter <NUM>. In this figure, the conventional space vehicle payload adapter <NUM> is shown to comprise a solid monocoque ring design. The conventional space vehicle payload adapter (e.g., referred to as a ring) <NUM> comprises a single aluminum forging that is machined down to its flanges 120a, 120b and its secondary payload adapters (e.g., in the form of webs) 110a, 110b,110c, 110d, 110e, 110f, which are each typically around a half an inch (ca. <NUM>) thick. The monocoque structure of the conventional space vehicle payload adapter <NUM> inherently defines a shell <NUM>. The shell <NUM> of the ring <NUM> carries the stress of any loads exerted upon the space vehicle payload adapter <NUM>.

It should be noted that, since the secondary payload adapters (e.g., webs) 110a, 110b,110c, 110d, 110e, 110f are machined into the ring <NUM>, the locations of the secondary payload adapters 110a, 110b,110c, 110d, 110e, 110f on the exterior circumference of the ring <NUM> are fixed. As such, the conventional space vehicle payload adapter <NUM> design does not provide the ability to clock (e.g., rotate) the secondary payloads to different locations on the exterior circumference ring <NUM> to optimize the center of gravity (CG) for the space vehicle (e.g., a launch vehicle combined with the payloads).

In addition, the conventional space vehicle payload adapter <NUM> design makes access to the inside of the ring <NUM> very difficult because the interior of the ring <NUM> cannot be accessed from the secondary payload adapters 110a, 110b,110c, 110d, 110e, 110f located on the periphery of the ring <NUM>, after the secondary payloads are mounted onto the ring <NUM>. Additionally, the thickness (e.g., typically around a half an inch, ca. <NUM>) of the secondary payload adapters 110a, 110b,110c, 110d, 110e, 110f provides a relatively stiff interface at the cost of considerable weight.

<FIG> is a diagram (not to scale) showing an exemplary launch sequence <NUM> for a spacecraft (e.g., a primary satellite, referred to as a primary payload) <NUM> that employs the disclosed space vehicle payload adapter <NUM>, in accordance with at least one embodiment of the present disclosure. In this figure, the primary payload <NUM> as well as a plurality of secondary satellites (e.g., referred to as secondary payloads) 220a, 220b, 220c are mounted onto a launch vehicle (e.g., the launch vehicle upper stage (LVUS)) <NUM> via the disclosed space vehicle payload adapter <NUM>. The primary payload <NUM>, the secondary payloads 220a, 220b, 220c, and the space vehicle payload adapter <NUM> are all initially housed within the payload fairing <NUM> of the launch vehicle (e.g., LVUS) <NUM>.

At the beginning of the launch sequence <NUM>, during pre-launch phase of the launch sequence <NUM>, a payload van (PVAN) <NUM> comprising electrical ground support equipment (EGSE) is in communication (e.g., transmitting auxiliary (AUX) payload (PL) telemetry (TLM)) with a spacecraft operations center (SOC) <NUM> as well as with the launch vehicle (e.g., LVUS) <NUM>. The launch vehicle (e.g., LVUS) <NUM> is connected to a launch vehicle lower stage (LVLS) <NUM>, and is located on a launch pad <NUM> on the ground.

During launch, the primary payload <NUM> as well as the secondary payloads 220a, 220b, 220c experience high levels of vibration. As such, it is important that the space vehicle payload adapter <NUM> is manufactured and designed to have sufficient stiffness such that high frequency modes (e.g., > <NUM>) are maintained during all phases of the launch sequence.

During the launch and ascent phase of the launch sequence <NUM>, the LVLS <NUM> separates from the launch vehicle (e.g., LVUS) <NUM>, and the payload fairing <NUM> separates from the launch vehicle (e.g., LVUS) <NUM>. Also during the launch and ascent phase, the launch vehicle (e.g., LVUS) <NUM> transmits telemetry information to the SOC <NUM> via a ground station antenna 270a.

In one or more embodiments, when the launch vehicle (e.g., LVUS) <NUM> has reached a lower earth orbit (LEO) or a geostationary transfer orbit (GTO), at least one of the secondary payloads 220a, 220b, 220c mounted onto the disclosed space vehicle payload adapter <NUM> is deployed into space. Also, when the launch vehicle (e.g., LVUS) <NUM> has reached a lower earth orbit (LEO) or a geostationary transfer orbit (GTO), the launch vehicle (e.g., LVUS) <NUM> transmits telemetry information to the SOC <NUM> via the ground station antenna 270a.

Then, the launch vehicle (e.g., LVUS) <NUM> performs a first trans-launch injection (TLI) maneuver, which is a population maneuver. During the first TLI maneuver, the launch vehicle (e.g., LVUS) <NUM> transmits telemetry information to the SOC <NUM> via a ground station antenna 270b.

After the first TLI maneuver, the primary payload <NUM> mounted onto the disclosed space vehicle payload adapter <NUM> is deployed into space. During the deployment, the launch vehicle (e.g., LVUS) <NUM> transmits telemetry information to the SOC <NUM> via the ground station antenna 270b.

In one or more embodiments, the launch vehicle (e.g., LVUS) <NUM> performs a secondary TLI maneuver. During this secondary TLI maneuver, at least one of the secondary payloads 220a, 220b, 220c mounted onto the disclosed space vehicle payload adapter <NUM> is deployed into space. Also during this secondary TLI maneuver, the launch vehicle (e.g., LVUS) <NUM> transmits telemetry information to the SOC <NUM> via a ground station antenna 270b. After the secondary TLI maneuver, it is the end of the mission (EoM) for the launch sequence <NUM>.

It should be noted that the launch sequence <NUM> depicted in <FIG> is merely one exemplary launch sequence of phases that may be employed for a spacecraft (e.g., a primary payload) <NUM> utilizing the disclosed space vehicle payload adapter <NUM>. As such, in one or more embodiments, launch sequences other than the specific launch sequence <NUM> shown in <FIG> may be employed by a spacecraft (e.g., a primary payload) <NUM> utilizing the disclosed space vehicle payload adapter <NUM>. For example, in one or more embodiments, during the launch sequence, after the spacecraft (e.g., a primary payload) <NUM> separates from the space vehicle payload adapter <NUM>, the space vehicle payload adapter <NUM> may then separate from the launch vehicle (e.g., LVUS) <NUM>. In another example, during the launch sequence, at least one of the secondary payloads 220a, 220b, 220c mounted onto the disclosed space vehicle payload adapter <NUM> may be hosted on the space vehicle payload adapter <NUM> for the duration of the life of the secondary payload(s) 220a, 220b, 220c. For another example, during the launch sequence, the secondary payloads 220a, 220b, 220c may be deployed from the space vehicle payload adapter <NUM> into different specific orbital states and/or locations.

<FIG> is a diagram showing details of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. In this figure, the space vehicle payload adapter <NUM> is located in beneath a primary payload <NUM>, and is located above a launch vehicle (e.g., LVUS) <NUM>. In particular, the primary payload <NUM> is attached (e.g., releasably attached) to a forward ring 450a of the space vehicle payload adapter <NUM>, and the launch vehicle (e.g., LVUS) <NUM> is attached to an aft ring 450b of the space vehicle payload adapter <NUM>. In addition, a plurality of secondary payloads 220a, 220b, 220c are attached (e.g., releasably attached) to the periphery of the space vehicle payload adapter <NUM> via secondary payload adapters <NUM>.

<FIG> is a diagram showing a perspective view of the disclosed space vehicle payload adapter <NUM>, in accordance with at least one embodiment of the present disclosure. The space vehicle payload adapter <NUM> comprises a forward ring 450a, an aft ring 450b, three interstitial rings 430a, 430b, 430c, multiple secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f, and an open truss structure comprising a plurality of struts <NUM>. It should be noted that, in one or more embodiments, all of the components of the disclosed space vehicle payload adapter <NUM> are manufactured from a low cost and lightweight material(s) (e.g., a lightweight metal, such as aluminum or titanium, and/or a composite material).

Employing aluminum (and/or other lightweight material(s)) for the components of the space vehicle payload adapter <NUM> allows for the space vehicle payload adapter <NUM> to be lightweight and low cost to manufacture. In one or more embodiments, when components of the space vehicle payload adaptor <NUM> are manufactured from aluminum, the aluminum components are coated with a conductive Alodine finish to provide for aluminum corrosion protection. It should be noted that although the conventional space vehicle payload adapter (e.g., ring) <NUM> of <FIG> is also manufactured from aluminum (e.g., a lightweight material), the disclosed space vehicle payload adapter <NUM> is much lighter in weight because it employs an open truss structure design as opposed to the conventional solid monocoque ring design (e.g., refer to the ring <NUM> of <FIG>), which adds a considerable amount of weight.

In <FIG>, the disclosed space vehicle payload adapter <NUM> is shown to have an annular profile defining a circumference <NUM>. In one or more embodiments, the space vehicle payload adapter <NUM> comprises a forward open end 465a defined by the forward ring 450a, and comprises an aft open end 465b defined by the aft ring 450b. The forward ring 450a is configured to directly (or indirectly) couple (e.g., releasably attach) the space vehicle payload adapter <NUM> to a primary payload (e.g., refer to <NUM> of <FIG>). And, the aft ring 450b is configured to directly (or indirectly) couple (e.g., releasably attach) the space vehicle payload adapter <NUM> to a launch vehicle (e.g., refer to <NUM> of <FIG>).

In one or more embodiments, the forward ring 450a and the aft ring 450b have the same diameter (D), as is shown in <FIG>. However, in other embodiments, the forward ring 450a and the aft ring 450b may have different diameters. For example, the forward ring 450a may have a first diameter (D1) and the aft ring 450b may have a second diameter (D2), where the first diameter (D1) is larger than (or, alternatively, smaller than) the second diameter (D2). The specific diameters of the forward ring 450a and the aft ring 450b of the space vehicle payload adapter <NUM> are determined based on the packaging requirements of the payload fairing of the launch vehicle (e.g., refer to the payload fairing <NUM> of the LVUS <NUM> of <FIG>).

The space vehicle payload adapter <NUM> also comprises an open truss structure comprising a plurality of struts <NUM> connecting the forward ring 450a to the aft ring 450b and, as such, each of the struts <NUM> extends from the forward ring 450a to the aft ring 450b. A first end of each of the struts <NUM> is connected to the forward ring 450a, and a second end of each of the struts <NUM> is connected to the aft ring 450b. In one or more embodiments, the struts <NUM> are oriented at consistent angles to form alternately inverted isosceles triangle-shaped openings <NUM> around the circumference <NUM> of the space vehicle payload adapter <NUM>. This specific arrangement of the struts <NUM> in the open truss structure of the space vehicle payload adapter <NUM> is referred to in structural engineering as a "Warren truss" or an "equilateral truss".

In addition, the space vehicle payload adapter <NUM> comprises six secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f disposed around the circumference <NUM> of the space vehicle payload adapter <NUM>. The secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f are each configured to secure (e.g., releasably attach) a secondary payload (e.g., refer to 610a, 610b, 610c, 610d, 610e, 610f of <FIG>) onto the space vehicle payload adapter <NUM>. In one or more embodiments, the space vehicle payload adapter <NUM> may comprise more (or, alternatively, less) than six secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f than as shown in <FIG>.

In one or more embodiments, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be of various different sizes. For example, some of the secondary payload adapters 420a, 420b, 420c, 420e may have smaller mountings than the other secondary payload adapters 420d, 420f. In particular, for example in <FIG>, secondary payload adapters 420a, 420b, 420c, 420e each have a small mounting size (e.g., an adapter port opening, which is circular in shape, that is approximately fifteen (<NUM>) inches, ca. <NUM>, in diameter) for mounting smaller secondary payloads than secondary payload adapters 420d, 420f, which have a large mounting size (e.g., an adapter port opening, which is rectangular in shape, that is approximately twenty-four (<NUM>) inches, ca. <NUM>, in height and width) for mounting larger secondary payloads, which are larger in size (volume) and/or weight than the smaller secondary payloads.

It should be noted that, in one or more embodiments, more (or, alternatively, less) than two different sizes (and/or weights) of secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be employed by the disclosed space vehicle payload adapter <NUM>. In addition, in one or more embodiments, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may comprise adapter port openings of various different shapes (e.g., triangular or a polygon) other than the circular and rectangular shapes as are shown in <FIG>. As such, the shapes (as well as the sizes) of the adaptor port openings of the secondary payload adaptors 420a, 420b, 420c, 420d, 420e, 420f may be customized according to (e.g., complementary to) the shapes (and sizes) of the interfaces of the secondary payloads (e.g., refer to 610a, 610b, 610c, 610d, 610e, 610f of <FIG>) to be attached to the space vehicle payload adaptor <NUM>. Thus, the adapter port openings of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f comprise shapes (and sizes) complementary to interfaces of the secondary payloads.

In addition, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f each comprise a plurality of bolt holes for the releasable attachment of the secondary payloads (e.g., refer to 610a, 610b, 610c, 610d, 610e, 610f of <FIG>) via fasteners (e.g., bolts, such as kinematic mount bolts or easy ride adapters). In <FIG>, the bolt hole pattern on each of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f for the releasable attachment of the secondary payloads is shown to be in the form of a circle. However, it should be noted that, in one or more embodiments, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may comprise bolt hole patterns of various different shapes (e.g., triangular or a polygon) other than the circular bolt hole pattern as is shown in <FIG>. The shapes (as well as the sizes) of the bolt hole patterns of the secondary payload adaptors 420a, 420b, 420c, 420d, 420e, 420f may be customized according to (e.g., complementary to) the shapes (and sizes) of the bolt hole patterns on the interfaces of the secondary payloads (e.g., refer to 610a, 610b, 610c, 610d, 610e, 610f of <FIG>). Also, it should be noted that, in <FIG>, each of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f is depicted to be composed of a single component (not including the fasteners). However, it should be noted that, in one or more embodiments, each of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be manufactured to comprise more than one component (not including the fasteners), as is shown in <FIG>.

Additionally, each of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f is releasably attached to at least one of the interstitial rings 430a, 430b, 430c of the space vehicle payload adapter <NUM>. The three interstitial rings 430a, 430b, 430c each comprise a plurality of mounting fixtures <NUM> (e.g., in the form of bolt holes) disposed on the exterior surface around the circumference <NUM>. And, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f each comprise multiple attachment points <NUM> (e.g., in the form of bolt holes). The mounting fixtures <NUM> of the interstitial rings 430a, 430b, 430c are configured to releasably attach to the attachment points <NUM> (e.g., via removable fasteners (e.g., bolts, such as kinematic mount bolts or easy ride adapters)) of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f. As such, in particular in <FIG>, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f are releasably attached to at least one of the interstitial rings 430a, 430b, 430c via removable fasteners (e.g., bolts) (e.g., refer to <NUM> of <FIG>) disposed within the mounting fixtures <NUM> of the interstitial rings 430a, 430b, 430c and within the attachment points <NUM> of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f. For example, refer to <FIG> for details of the attaching of secondary payload adapter 420a to interstitial ring 430a via removable fasteners (e.g., bolts) <NUM>.

However, it should be noted that, the use of fasteners (e.g., bolts) for the attaching of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f to the interstitial rings 430a, 430b, 430c is merely exemplary, and that any other device for releasably attaching the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f to the interstitial rings 430a, 430b, 430c may be employed by the disclosed space vehicle payload adapter <NUM>. As such, in one or more embodiments, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be releasably attached to at least one of the interstitial rings 430a, 430b, 430c by another means (e.g., by clamping, by drilling on the assembly, or by slots) other than by using fasteners (e.g., bolts) as in <FIG>.

In one or more embodiments, the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f are "clockable" such that they may be moved to different locations (e.g., to different clockable positions) around the circumference of the space vehicle payload adapter <NUM> to balance the center of gravity of the launch vehicle. For example, at least one of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be removed from its location on the circumference <NUM> of the space vehicle payload adapter <NUM> and, then, attached to another location on the circumference <NUM> of the space vehicle payload adapter <NUM>. In particular, one of the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f may be released from its attachment to a first location on at least one of the interstitial rings 430a, 430b, 430c. Then, the secondary payload adapter 420a, 420b, 420c, 420d, 420e, 420f may be releasably attached to a second location on at least one of the interstitial rings 430a, 430b, 430c.

It should be noted that the center of gravity of the launch vehicle should be balanced to ensure controllability during launch. If the secondary payloads (e.g., refer to 610a, 610b, 610c, 610d, 610e, 610f of <FIG>) are not of equal size (volume) and/or mass, and/or are attached to the payload adapter in a non-symmetrical manner, then the center of gravity of the launch vehicle may exceed the controllable offset limit. As a result, a ballast may be required to provide balance to the launch vehicle. However, introducing a ballast reduces the amount of usable payload mass that the launch vehicle can carry.

Each of the interstitial rings 430a, 430b, 430c of the space vehicle payload adapter <NUM> is attached to the struts <NUM> of the open truss structure of the space vehicle payload adapter <NUM>. Each of the struts <NUM> comprises a plurality of mounting apertures <NUM> (e.g., in the form of bolt holes) each configurable to receive a fastener (e.g., a bolt). The mounting fixtures <NUM> of the interstitial rings 430a, 430b, 430c are configured to attach to the mounting apertures <NUM> (e.g., via removable fasteners (e.g., bolts)) of the struts <NUM>. Also, each of the struts <NUM> comprises at least one V-shaped joint (e.g., refer to <NUM> of <FIG>), which comprises a first portion (e.g., refer to 482a of <FIG>) and a second portion (e.g., refer to 482b of <FIG>), configured to receive one of the interstitial rings 430a, 430b, 430c.

For example, for the attaching of an interstitial ring 430b (refer to <FIG>) to a strut <NUM>, the interstitial ring 430b is disposed within the V-shaped joint (e.g., refer to <NUM> of <FIG>) of the strut <NUM>, thereby forming a nested joint. Specifically, the interstitial ring 430b slides between the first portion (e.g., refer to 482a of <FIG>) and the second portion (e.g., refer to 482b of <FIG>) of the V-shaped joint (e.g., refer to <NUM> of <FIG>) of the strut <NUM>. After the interstitial ring 430b is disposed within the V-shaped joint (e.g., refer to <NUM> of <FIG>) of the strut <NUM>, the strut <NUM> is attached to the interstitial ring 430b via fasteners (e.g., bolts) disposed within the mounting apertures <NUM> of the strut <NUM> and within the mounting fixtures <NUM> of the interstitial ring 430b. Refer to <FIG> for details of the attaching of a strut <NUM> to an interstitial ring 430b.

It is important to note that the interstitial rings 430a, 430b, 430c are employed by the the disclosed space vehicle payload adapter <NUM> to provide sufficient stiffness to the space vehicle payload adapter <NUM> such that high frequency modes (e.g., > <NUM>) are always maintained during all aspects of launch (e.g., during all of the phases of the launch sequence). During launch, the primary payload <NUM> as well as the secondary payloads 220a, 220b, 220c experience high levels of vibration. The high levels of vibration experienced by the secondary payloads 220a, 220b, 220c create bending loads in the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f. The interstitial rings 430a, 430b, 430c react the bending loads in the secondary payload adapters 420a, 420b, 420c, 420d, 420e, 420f into the struts <NUM> of the open truss structure. The struts <NUM> then react the loads to the forward ring 450a and aft ring 450b and, as a result, high frequency modes are maintained for the space vehicle payload adapter <NUM>.

It should be noted that in one or more embodiments, the disclosed space vehicle payload adapter <NUM> comprises more than two full interstitial rings 430a, 430b, 430c. By employing more than two full interstitial rings 430a, 430b, 430c within the disclosed space vehicle payload adapter <NUM>, high frequency modes are able to be maintained for the space vehicle payload adapter <NUM>. And, in one or more embodiments, at least one of the interstitial rings 430a, 430b, 430c employed by the disclosed space vehicle payload adapter <NUM> may be merely a partial interstitial ring 430a, 430b, 430c (e.g., comprising at least one segment of an interstitial ring 430a, 430b, 430c, but not all of the segments for a full interstitial ring 430a, 430b, 430c).

<FIG> is a diagram showing details of a portion <NUM> of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. In particular, the diagram of <FIG> shows details of a portion <NUM> in <FIG> depicting the releasable attachment of secondary payload adapter 420a to interstitial ring 430a. In particular, in this figure, the secondary payload adapter 420a is shown to be releasably attached to interstitial ring 430a via removable fasteners (e.g., bolts) <NUM> disposed within the mounting fixtures <NUM> of the interstitial ring 430a and within the attachment points <NUM> of the secondary payload adapter 420a.

<FIG> is a diagram showing details of one of the secondary payload adapters 420a of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. In particular, in this figure, details of the releasable attachment of secondary payload adapter 420a to all three interstitial rings 430a, 430b, 430c is shown.

<FIG> is a diagram showing details of the nested joint design for the interstitial rings (e.g., interstitial ring 430b) of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. In particular, the diagram of <FIG> shows details of the interior side (e.g., the back side not visible in <FIG>) of the portion <NUM> denoted in <FIG>. In <FIG>, interstitial ring 430b is shown to be attached to a strut <NUM>. For the attaching, the interstitial ring 430b is disposed between the first portion 482a and the second portion 482b of the V-shaped joint <NUM> of the strut <NUM>, resulting in a nested joint configuration. Then, the strut <NUM> is attached to the interstitial ring 430b via fasteners (e.g., bolts) (not shown in <FIG>) disposed within the mounting apertures <NUM> of the strut <NUM> and within the mounting fixtures <NUM> of the interstitial ring 430b.

<FIG> is a diagram showing details of the two-part design for the interstitial rings 430a, 430b, 430c of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. And, <FIG> is an exploded detailed view of the diagram of <FIG> showing details of the two-part design (e.g., showing the inner portions (I) and outer portions (O)) of the interstitial rings 430a, 430b, 430c of the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure.

It should be noted that the two-part design shown in <FIG> and <FIG> may be used in conjunction with (or, alternatively, used without) the nested joint design shown in <FIG> for the disclosed interstitial rings 430a, 430b, 430c. Similarly, it should be noted that the nested joint design shown in <FIG> may be used in conjunction with (or, alternatively, used without) the two-part design shown in <FIG> and <FIG> for the disclosed interstitial rings 430a, 430b, 430c.

In one or more embodiments, for the two-part design shown in <FIG> and <FIG>, each of the interstitial rings 430a, 430b, 430c comprises multiple (e.g., six) sections (e.g., segments (<NUM>), (<NUM>), (<NUM>), (<NUM>), (<NUM>), and (<NUM>)) circumferentially. And, each of the segments (<NUM>), (<NUM>), (<NUM>), (<NUM>), (<NUM>), and (<NUM>) of each interstitial ring 430a, 430b, 430c comprises an inner portion (I) and an outer portion (O). For example, in <FIG>, interstitial ring 430c is shown to comprise multiple segments, such as a first segment (<NUM>) and a second segment (<NUM>). And, each segment (<NUM>), (<NUM>) is shown to comprise an inner portion (I) and an outer portion (O). As such, in the exploded view of <FIG>, interstitial ring 430c is shown to comprise 430c(<NUM>)(O) (i.e. segment (<NUM>), outer portion (O)); 430c(<NUM>)(I) (i.e. segment (<NUM>), inner portion (I)); 430c(<NUM>)(O) (i.e., segment (<NUM>), outer portion (O)); and 430c(<NUM>)(I) (i.e. segment (<NUM>), inner portion (I)). The outer portions (O) of each of the segments (<NUM>), (<NUM>), (<NUM>), (<NUM>), (<NUM>), and (<NUM>) are attached to their corresponding inner portions (I) via at least one fastener (e.g., bolt) <NUM>. For example, with regard to segment (<NUM>) of interstitial ring 430c, 430c(<NUM> )(O) is attached to 430c(<NUM>)(I) via fasteners (e.g., bolts) <NUM>.

<FIG> is a diagram showing a perspective view of the disclosed space vehicle payload adapter <NUM> of <FIG>, where electronic components 510a, 510b, 510c, 510d, 510e are mounted onto the interstitial rings 430a, 430b, 430c, in accordance with at least one embodiment of the present disclosure. In one or more embodiments, the interstitial rings 430a, 430b, 430c are configured to accommodate electrical equipment. In this figure, various different electrical components 510a, 510b, 510c, 510d, 510e are shown to be mounted to the top (forward) side or the bottom (aft) side of the interstitial rings 430a, 430b, 430c. Various different types of electrical equipment may be mounted onto the interstitial rings 430a, 430b, 430c of the disclosed space vehicle payload adapter <NUM> including, but not limited to, a power source (e.g., a battery), a propulsion and mechanism module (PAM), and an embedded relay module (ERM).

It should be noted that in one or more embodiments, the interstitial rings 430a, 430b, 430c comprise wiring holes (e.g., refer to <NUM> in <FIG>, <FIG>, and <FIG>) to provide pathways for routing wiring from the electrical components 510a, 510b, 510c, 510d, 510e between the struts <NUM>. As such, wiring connected to the electrical components 510a, 510b, 510c, 510d, 510e may be fed through the wiring holes on the interstitial rings 430a, 430b, 430c to their corresponding payloads.

<FIG> is a diagram showing a perspective view of secondary satellites (e.g., secondary payloads) 610a, 610b, 610c, 610d, 610e, 610f mounted onto the disclosed space vehicle payload adapter <NUM> of <FIG>, in accordance with at least one embodiment of the present disclosure. In this figure, the secondary payloads 610a, 610b, 610c, 610d, 610e, 610f radially mounted onto the space vehicle payload adapter <NUM> are of unequal size (volume) and mass. In particular, secondary payloads 610a, 610b, 610d, 610e are shown to be smaller in size (volume) than secondary payloads 610c, 610f. It should be noted that the configuration shown in <FIG> is merely exemplary in nature and that any number and composition of secondary payloads 610a, 610b, 610c, 610d, 610e, 610f may be mounted to the disclosed space vehicle payload adapter <NUM>. In addition, it should be noted that although <FIG> shows the secondary payloads 610a, 610b, 610c, 610d, 610e, 610f space equidistantly apart from one another, the locations of the secondary payloads 610a, 610b, 610c, 610d, 610e, 610f mounted onto the space vehicle payload adapter <NUM> may vary.

<FIG> is a diagram showing a perspective view of the disclosed space vehicle payload adapter <NUM> of <FIG> comprising circumferential exterior multi-layer insulation (MLI) <NUM>, in accordance with at least one embodiment of the present disclosure. In this figure, the exterior circumference <NUM> (excluding the secondary payload adapters 720a, 720b, 720c, 720d, 720e, 720f) of the space vehicle payload adapter <NUM> is covered (e.g., wrapped) with a MLI <NUM> material, such a MLI <NUM> blanket. MLI <NUM> is a thermal insulation comprising multiple layers (e.g., six layers) of thin sheets. Wrapping the space vehicle payload adapter <NUM> with MLI <NUM> ensures that the space vehicle payload adapter <NUM> maintains a constant interior temperature despite the extreme temperature fluctuations in space.

<FIG> is a diagram showing a perspective view of secondary satellites (e.g., secondary payloads) 610a, 610b, 610c, 610d, 610e, 610f mounted onto the disclosed space vehicle payload adapter <NUM> of <FIG> comprising circumferential exterior MLI <NUM>, in accordance with at least one embodiment of the present disclosure. In this figure, the MLI <NUM> blanket is shown to be covering the exterior surface around the circumference of the space vehicle payload adapter <NUM>, except for the portions of the exterior surface comprising the secondary payload adapters.

Claim 1:
A method for reacting loads into a space vehicle payload adapter (<NUM>), the method comprising:
reacting (<NUM>), by more than two interstitial rings (430a, 430b, 430c) of the space vehicle payload adapter (<NUM>), the loads created by secondary payloads (220a, 220b, 220c) mounted onto the space vehicle payload adapter (<NUM>), into a truss structure of the space vehicle payload adapter (<NUM>); and
reacting (<NUM>), by struts (<NUM>) of the truss structure, the loads to a forward ring (450a) and an aft ring (450b) of the space vehicle payload adapter (<NUM>);
wherein the struts (<NUM>) connect the forward ring (450a) to the aft ring (450b).