Orbital transport systems and devices

Abstract

The present invention is embodied in systems and devices for transporting assets (small satellites, goods, or other payloads) in outer space. A preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. Since the orbits are complementary to each other, the relay stations can transport mass objects from one station to another. In this way, the system can support various logistics networks connecting space between orbits such as low Earth orbit (LEO), geostationary orbit (GEO), and low lunar orbit (LLO), as well as space between Earth and Mars.

Description

Technical Field

[0001] (Field of the Invention) The present invention relates to the field of transportation, and more specifically, to the field of mass transportation in outer space.

Background Art

[0002] At present, the space economy is worth $469 billion and is expected to experience even more explosive growth in the next decade. Over the past 20 years, the trend towards reusable launch vehicles has reduced the costs associated with going into space, making new industries centered around in-space services, assembly, and manufacturing (ISAM) increasingly viable. The growth of small and medium-sized enterprises, startups, and large established companies in the space economy is focused on making these new markets profitable and sustainable. Currently, market analysts predict that the in-orbit fuel replenishment market, a major component of ISAM, will grow at a CAGR of over 100% until 2032. The realization of ISAM also required the development of adjacent markets, including the Space Domain Awareness (SDA) sector. In an increasingly crowded space domain, SDA is strategically important as new capabilities are needed to characterize assets in orbit.

[0003] Importantly, this trend is not exclusive to the private sector. At the time of writing this article, NASA launched the Artemis program as part of its efforts to establish a sustainable presence in deep space. Similarly, the United States Space Force established a program execution unit specialized in Space Mobility and Logistics (SML), and the Air Force Research Laboratory installed an ISAM unit within its spacecraft directorate. Ultimately, the combination of private and government interest in the space economy is a powerful vector for generating new opportunities and economic growth.

[0004] Achieving a highly advanced and sustainable space economy requires addressing numerous complex challenges, one of the most fundamental being logistics and transportation. Currently, delivering goods via space (beyond LEO) is extremely expensive. Even with reusable launch platforms, the cost beyond LEO remains a major obstacle to the space economy. The primary reason for the high cost of space transportation is that existing space logistics solutions (rockets, orbital transporters, etc.) are massive and require large amounts of consumable propellant to deliver payloads to their orbital destinations. Furthermore, moving payloads in space requires accelerating those payloads, fuel, and chassis using existing logistics means such as rockets and orbital transporters. In other words, they are constrained by the rocket equation.

[0005] There are many approaches to converting electrical energy into linear motion, including linear induction motors and railgun (Lorentz force) actuators. Concepts developed for this type of system so far have utilized tethers (which are prone to entanglement and difficult to deploy) or rotated the entire satellite (which results in low rotational speeds and a high risk of losing control of the satellite). Railguns are highly wear-and-tear and have large masses.

[0006] Ultimately, a robust ISAM sector will require low-cost logistics solutions for transporting essential inputs such as propellants, raw materials, replacement parts, and structural components (like antennas) in space. [Overview of the Initiative]

[0007] The subject matter of this invention is embodied in systems and methods for transporting assets (small satellites, goods, or other payloads) in outer space. A preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. Since the orbits are complementary to each other, the relay stations can transport mass objects from one station to another. Simply put, a “relay station” is a device that deploys or redeploys capsules launched from other relay stations. Launched capsules can carry various materials, sensors, and other cargo. As will be described in detail below, the system supports various logistics networks connecting space between orbits such as Low Earth Orbit (LEO), Geostationary Orbit (GEO), and Low Lunar Orbit (LLO), as well as space between Earth and Mars.

[0008] A preferred system comprises a first relay station in a first orbit around a celestial body and a second relay station in a second orbit around the celestial body. The preferred relay station includes a rigid rotary arm configured to rotate around a central shaft. In a launch operation, a motor rotates the rotary arm to a desired rotational speed, and the relay station releases the payload from the distal end of the rotary arm at a specific speed and angle. A pair of momentum wheels connected to the central shaft of the first relay station counteract and absorb the rotational inertia caused by the rotation of the rotary arm and the release of the payload. The preferred relay station also includes a receiver capable of collecting or "catching" the payload as it flies through space.

[0009] This feature allows the first relay station to launch a payload onto an orbit that intersects with the second relay station in the second orbit. The receiver and rotary arm of the second relay station allow it to receive payloads from the first (or adjacent) relay station and launch the same payload to the next station (above or below) in the constellation. Placing multiple relay stations in complementary orbits between two points in space creates a constellation of relay stations capable of moving mass objects anywhere between those two points in space.

[0010] The key to this invention is converting solar energy and / or nuclear energy into kinetic energy to move a mass in space, thereby minimizing the use of consumable propellants to move the payload in space. [Brief explanation of the drawing]

[0011] [Figure 1] A perspective view of a preferred relay station mounted on a satellite bus is shown.

[0012] [Figure 2] An embodiment of a small relay station is shown.

[0013] [Figure 3] A perspective view of one embodiment of the relay station 10 is shown.

[0014] [Figure 4] Figure 3 shows a front view of the embodiment.

[0015] [Figure 5] Figure 3 shows a side view of the embodiment.

[0016] [Figure 6] Figure 3 shows a perspective view of the embodiment with the counterbalance assemblies and momentum wheels removed.

[0017] [Figure 7] An exploded view of an embodiment of the rotary arm is shown.

[0018] [Figure 8] Figure 7 shows a cross-sectional view.

[0019] [Figure 9]Perspective view of the rotary arm and the counterbalance assembly, showing the counterbalance assembly in position 1. Note that button 58 is pushed in.

[0020] [Figure 10] Perspective view of the rotary arm and the counterbalance assembly, showing the counterbalance assembly in position 3. Note that button 58 is pulled out.

[0021] [Figure 11] Perspective view of the counterbalance assembly disassembled from the rotary arm.

[0022] [Figure 12] Shows the cross-sectional view of FIG. 15.

[0023] [Figure 13] Perspective view of the counterbalance assembly.

[0024] [Figure 14] Shows the cross-sectional view of FIG. 17.

[0025] [Figure 15] Exploded view of FIG. 17.

[0026] [Figure 16] Outer perspective view of an embodiment of the rotor hub.

[0027] [Figure 17] Inner perspective view of an embodiment of the counterweight assembly.

[0028] [Figure 18] Outer perspective view of an embodiment of the counterweight assembly with the outer hub removed.

[0029] [Figure 19]This is an internal perspective view of an embodiment of the counterweight outer hub.

[0030] [Figure 20] This is a perspective view of an embodiment of a rotary arm, which has a release mechanism on the first spoke and a catch mechanism on the second spoke.

[0031] [Figure 21] Figure 29 is an exploded view.

[0032] [Figure 22] This is a rear perspective view of an embodiment of the catch mechanism 45.

[0033] [Figure 23] This is a diagram illustrating an embodiment of a motor / stator mounted on the main shaft.

[0034] [Figure 24] This is a diagram of a motor sleeve.

[0035] [Figure 25] This is a diagram of the stator hub.

[0036] [Figure 26] This is a perspective view of a preferred embodiment of the relay station 10.

[0037] [Figure 27] Figure 26 is an exploded view of the embodiment.

[0038] [Figure 28] This is an exploded perspective view of a preferred embodiment of the momentum wheel 94.

[0039] [Figure 29] This is an exploded perspective view of a preferred embodiment of the momentum wheel 94, with the rotating arrows indicating which part rotates.

[0040] [Figure 30] This is a second diagram of the momentum wheel 94 in Figure 26.

[0041] [Figure 31] Figure 30 is a cross-sectional view of the momentum wheel, with each component shown separated from the others for clarity.

[0042] [Figure 32] This shows the node architecture of the relay station connecting LEO and GEO.

[0043] [Figure 33] This shows a circular-elliptic-circular constellation with three nodes in complementary orbits.

[0044] [Figure 34] This shows a circular constellation with multiple nodes.

[0045] [Figure 35] This shows a circular constellation with three nodes.

[0046] [Figure 36] This shows a constellation that transitions from a circular to an elliptical shape with a change in the angle of inclination.

[0047] [Figure 37] The launch and recovery embodiments are shown.

[0048] [Figure 38] The launch and catch mechanisms are shown.

[0049] [Figure 39] Embodiments of launch and capped catch are shown.

[0050] [Figure 40]Figure 40 illustrates an embodiment in which a single relay station launches the capsule toward itself. Figure 40 shows what happens if the capsule is not recovered by a receiving relay station, and the capsule eventually returns to its original launch site.

[0051] [Figure 41] An embodiment of a relay station equipped with a retractable rotary arm 154 in a folded state is shown. [Figure 42] An embodiment of a relay station equipped with a retractable rotary arm 154 in a folded state is shown.

[0052] [Figure 43] The embodiment shown in Figure 41 is depicted with the arm extended. [Figure 44] The embodiment shown in Figure 41 is depicted with the arm extended. [Figure 45] The embodiment shown in Figure 41 is depicted with the arm extended.

[0053] [Figure 46] An embodiment of a relay station using Shroud 156 to protect the rotary system from orbital debris and cosmic dust is shown.

[0054] [Figure 47] An embodiment of a relay station comprising a main momentum wheel pair and a secondary momentum wheel pair is shown.

[0055] [Figure 48] This shows how the operating region changes depending on delta-v and orbital magnitude.

[0056] [Figure 49] This provides examples of various rotary arm lengths and the mechanical energy required to rotate the payload. [Modes for carrying out the invention]

[0057] (Description of preferred embodiments) Overview of preferred relay stations for capsule delivery This portion of this specification relates to relay stations 10 for delivering capsules 25 into space. The relay stations described below are typically part of a system of relay stations delivered across multiple strategic orbits. Typically, relay stations are mounted on satellite buses 150. However, relay stations can also be mounted on other spacecraft (shuttle, ship, craft, etc.). Furthermore, relay stations can be modified to operate independently by equipping them with their own thruster sets.

[0058] In this specification, the term "relay station" means a device that launches capsules launched from other relay stations and, if necessary, catches (or otherwise recovers) them. In this specification, the term "capsule" broadly refers to a device capable of containing supplies, sensors, and other cargo as it travels between relay stations. As will be explained in detail below, various logistics networks can be constructed connecting space to low Earth orbit (LEO), geostationary orbit (GEO), low lunar orbit (LLO), and even to Mars.

[0059] This specification describes launching assets at a linear velocity of 200 m / s in many cases. However, those skilled in the art will understand that the present invention is expandable. For example, see the table in Figure 48, which shows how the operating range changes depending on the delta-v and orbital magnitude. As shown, a delta-v of 75 m / s may be suitable for small relay stations, while a delta-v of 350 m / s may be suitable for large relay stations. The key is that the size of the rotary arm and the RPM are appropriate for the size of the asset being accelerated. Expanding the system can reduce costs or allow larger assets to be moved faster and over longer distances.

[0060] The following procedure outlines how the preferred relay station 10 operates in space orbit. The main elements of the preferred relay station are described in detail in the following sections. The 200 m / s launch and catch can generally be described as follows: 1. Capsule 25 is loaded into the relay station 10. See Figure 26 for a preferred relay station 10. 2. The motor accelerates the rotary arm to 34-640 RPM. A pair of momentum storage wheels 94 counteract the angular acceleration of the rotary arm 34, preventing the relay station 10 from rotating. 3. Once stable, the Momentum Storage Wheel 94 ceases to accelerate in order to counteract the movement of the Rotary Arm. This allows the precise Momentum Storage mechanism to fine-tune the orientation of the spacecraft, enabling it to be precisely pointed at the desired target. 4. The relay station 10 releases the capsule 25 at the desired angle and speed. (See the launch operation section below.) 5. Releasing the capsule changes the inertia of the rotary arm. To counteract this change in inertia, a dynamic counterbalancing system is preferably employed. (See the counterbalancing section below.) 6. The Momentum Storage Wheel 94 responds to changes in angular momentum and prevents the spacecraft from entering a spin state. 7. The receiving relay station 10 can be moved to a position to catch the capsule using the propulsion of the satellite bus. 8. Before catching, the rotary arm 34 of the receiving relay station 10 rotates so that the tangential velocity at the tip of the arm matches the linear velocity of the inbound capsule. Ideally, the additional net force is generated only by converting the capsule's velocity to angular velocity. (See the catching operation below.) 9. When the receiving relay station catches the capsule, the receiver's rotary arm 34 decelerates, allowing the receiver to lower the capsule. Through a regenerative breaking process, the receiver can recover some of the kinetic energy of the returning capsule, improving the system's energy efficiency. Similarly, the relay station can recover this energy when the launch arm / counterbalance / momentum storage wheel decelerates to 0 rpm.

[0061] Steps 8 and 9 can be replaced by the receiving relay station 10 crossing paths with the return capsule released from the relay station. In this case, the receiving relay station can then capture the capsule using a more conventional docking method or a robotic arm.

[0062] The relay station 10 has several key components, including a rotary arm, counterbalance, release mechanism, bridge, and momentum storage mechanism (reaction wheel and / or control moment gyroscope). The relay station is typically mounted on a satellite bus. The rotary arm, counterbalance, and release mechanism are disclosed in a published notification entitled "Low Gravity Power Distribution System and Device" (WO / 2023 / 212682 (February 11, 2023)), and are incorporated by reference in their entirety. Based on this overview, each of the key components of the preferred relay station 10 is described below. Framework / Optional Bridge

[0063] The relay station 10 is preferably mounted to the satellite bus 150 (or other spacecraft) via a base plate 152. As shown in the preferred embodiment in Figure 26, a pair of frames 31 are firmly connected to the base plate 152. The main shaft 28, from which the rotary arm 34 rotates, is connected between each frame 31.

[0064] Optionally, a rotating bridge (or "bridge") 17 can be used to rotate around the main shaft 28 via actuator motors, as shown in Figures 3-5. The primary purpose of the bridge is to create angle adjustment capabilities for payload launch and reception. However, bridges are not a preferred option in space. It is preferable to rotate the entire relay satellite via satellite bus attitude control rather than using a bridge to adjust the launch and reception angles.

[0065] The payload release magnets 62 and 63 can be mounted to the bridge via the bridge frame 20 and are used to activate the release mechanism (described later). In short, the payload release magnet 62 is preferably a single electromagnet that activates a cam collar for releasing the payload. The counterbalance magnets (60 and 61) are also mounted to the bridge via the bridge frame 20. The release magnets (62 and 63) and counterbalance magnets (60 and 61) are preferably positioned at different radii from the central shaft 28 so as not to interfere with each other. See Figure 4. Rotary arm

[0066] A preferred rotary arm 34 includes four main components: a rotor hub 66, a pair of spokes (40, 47), a release mechanism 43, and a catch mechanism 44. As shown in Figure 7, the rotor hub 66 incorporates a pair of roller bearings 37, a motor 65, and a rotary encoder 70. The spokes (40, 47) are connected to the distal end of the rotor hub 66. The spokes are preferably identical, as shown in Figure 20, and have a release mechanism at one end and a catch mechanism at the other. Optionally, a relay station may have multiple spokes. The release mechanism 43 is connected to the first spoke 40, and the catch mechanism 44 is connected to the second spoke 47. A shaft 28 passes through the center of the rotor hub 66. The release mechanism 43 houses the capsule during rotation and then acts to release (launch) the capsule from the tip of the rotary arm.

[0067] Motor 65 rotates the rotary arm. The motor is preferably a brushless DC direct drive motor (BLDC DDM) including a stator 72. The stator 72 can be epoxy bonded to a DDM sleeve 74 that is press-fitted into a hub 82. See Figures 23-25. The hub 82 can be press-fitted into a stationary shaft 28. See Figure 35. In this configuration, the motor can drive the rotary arm around the main shaft via bearings 37. Motor 65 is preferably positioned near the shaft and can rotate the rotary arm 34 to a desired speed.

[0068] Figures 26 and 27 show a preferred embodiment. As shown, the main shaft 28 is positioned along a central axis and connected at both ends to a frame 31. The frame provides structural support to the shaft and transmits forces to a satellite bus or other body (not shown). The main shaft 28 is preferably fixed and the rotary arm 34 is connected to the shaft via a pair of roller bearings 37 (see Figure 7). It is also preferable that the shaft 28 is tubular so that wiring and other elements (e.g., a rod 89) can pass through its interior.

[0069] When using a rotary launch mechanism, the release angle and velocity should be precisely controlled. Typically, the payload launch window is less than 1.5 milliseconds. The precise position and velocity of the rotary arm 34 can be actively monitored using a high-pulse-rate rotary encoder 70.

[0070] Optionally, the spokes (40, 47) of the rotary arm 34 can be curved. 5. By being curved, a deviated payload (i.e., one that is not directly caught at the tip) can be caught and then rolled to the tip of the arm, where a release mechanism holds it in place. Furthermore, the rotary arm 34 can be sized to withstand impacts reaching one-third of the distance to the center of the arm.

[0071] Similarly, both spokes (40, 47) can be positioned to catch the payload. If the payload is caught near the center of the arm, the kinetic energy imparted by the velocity difference may damage the rotary mechanism. There are two catch windows for each rotation (one for each spoke). The catch windows are larger than the launch windows. Optionally, the rotary arm can be dimensioned to three times the width of the payload to increase the tolerance for error during catching.

[0072] As shown in Figure 21, the catch mechanism 44 is connected to spokes (40, 47). The catch mechanism preferably includes a housing 45 with an open front. A funnel / guide 46 is attached to the housing to accommodate slight deviations in the incoming payload. The funnel / guide 46 guides the incoming object to a concave surface 42 attached to the inside of the housing 45. The catcher mechanism (or catcher's mitt) is used to catch the incoming capsule. The preferred catch surface is at least twice the size of the incoming capsule. release mechanism

[0073] In some cases, the rotary arm 34 houses the payload in the launch chamber 35. The launch chamber 35 includes a collar 68, which, when rotated, acts on four helical tightening cams (via teeth or gears). The cams move rollers (in and out) to lock and release the payload. Spoke covers are connected to spokes (40, 47) to allow rotation while fixing the cams in place.

[0074] During operation, the payload is initially locked in a predetermined position (closed position), in which state four cam-operated rollers hold the payload within the chamber 35. When the collar is activated, the cams rotate, releasing the capsule 25 within the launch chamber 35 ("open position"). Preferably, only the clamping cams hold the payload in place so that the capsule 25 does not come into contact with the side walls. Since the clamping cams are located near the widest diameter of the payload, only a small amount of activation is required when releasing.

[0075] The operation of activating the collar to release the payload (i.e., moving the release mechanism from the closed position to the open position) is preferably performed via the release magnet 62. (In this specification, “magnet” means a magnetic member; that is, “magnet” may be an actual magnet. However, it is preferable that the actual magnet be placed on the bridge frame 20 and transmit the magnetic field to the collar via a ferromagnetic material. The release magnet 62 is preferably an electromagnet. The collar preferably contains a ferromagnetic material. Thus, when the release magnet 62 is activated, it creates a magnetic couple with the collar and rotates it. Activating the collar using a magnet avoids the need to place an actuator that must withstand the G load at the end of the arm (essentially holding the payload from the opposite side to prevent it from flying out). Optionally, the tip of the release magnet 62 can be angled to increase the magnetic force on the collar. During release, a second collar magnet 62 acts to move the collar in the reverse direction and return it to the locked position.)

[0076] Alternatively, instead of using a magnet, the payload could be launched by extending a roller just once. When the roller extends sufficiently, it engages with the collar, causing the collar to rotate slightly relative to the entire release mechanism. As a result, the cam releases their hold on the payload, and the payload is launched.

[0077] The logic circuits necessary to ensure high precision within the system can be implemented on a microprocessor. The microprocessor enables the collection and interpretation of data acquired from connected devices and sensors. Furthermore, the microprocessor ensures that the motor operates under specified conditions, resulting in consistent firing and catching. The logic circuits within the microprocessor collect environmental feedback from the devices, ensuring that all components are protected and maintained within the desired operating conditions. Counterbalance

[0078] It is crucial for the relay station to maintain balance as much as possible during operation. Otherwise, vibrations induced by unbalance can lead to reduced launch accuracy, shortened relay station lifespan, or loss of control of the relay station's payload. Normally, the rotary arm 34 is balanced when the chamber 35 is empty. However, when a payload is placed in the chamber 35, the rotary arm 34 becomes unbalanced. As a result, without a counterbalancing system, the relay station becomes unbalanced as it rotates to the desired angular launch velocity. Alternatively, if the rotary arm 34 was balanced before the payload was released, it will become unbalanced after the release.

[0079] The counterbalancing assemblies (48, 49) address this issue by offsetting the payload load applied to the rotary arm 34 before launch and then balancing the load on the rotary arm after launch. The rotational inertia of each of the two counterweight assemblies is preferably equal. Furthermore, the total rotational inertia of the counterweight assemblies is preferably equal to the rotational inertia of the payload. Thus, the system is balanced when both counterweights 50, 52 are positioned on opposite sides of the payload. Similarly, when the chamber 35 is empty, the system is also balanced when both counterweights 50, 52 are positioned on opposite sides of the main shaft (i.e., 180° apart).

[0080] A preferred counterbalance assembly (48, 49) is connected to the main shaft on the opposite side of the rotary arm 34, as shown in Figure 15. The first counterweight 50 is adjustable to the first counterweight arm 54. Similarly, the second counterweight 52 is adjustable to the second counterweight arm 56. Both counterweights (50, 52) are preferably adjustable along the length of their respective counterweight arms, thereby allowing the total inertia of each arm to be changed depending on the situation (e.g., differences in payload weight). Adjustability can be achieved in various ways, such as screw connections.

[0081] During rotation and firing, the counterweight assemblies (48, 49) are preferably oriented on the same side relative to the main shaft, but away from the loading chamber 35. See Figure 13. Furthermore, the counterweight assemblies (48, 49) are preferably connected ("locked") to the rotary arm 34. In this state, the entire assembly (rotary arm, capsule, and the two counterweight assemblies (48, 49)) rotates on the rotary arm bearing 37. This is referred to as the first position.

[0082] When the capsule 25 is released from the chamber 35, the rotary arm 34 loses balance. This is because the capsule is no longer present in the release mechanism 43, and both counterweight assemblies remain connected to one side. To balance this state, both counterweight assemblies (48, 49) are first detached from the rotary arm 34. Once both counterweight assemblies are detached, the rotary arm 34 itself is in a balanced state (however, the entire system is not yet balanced; see the next paragraph). The rotary arm then rotates on the rotary arm bearing 37. The two counterweight assemblies (48, 49) can rotate freely on their respective corresponding counterweight bearings at different rpms than the rotary arm. This is called the second position.

[0083] The rotary arm 34 is balanced in the second position, but the counterweight assemblies (48, 49) are not balanced. This is because both counterweights (48, 49) remain on the same side relative to the shaft 28. To balance the counterweight assemblies (48, 49), one counterweight assembly needs to be rotated 180° relative to the other. When this happens, the counterweight assemblies (48, 49) are also balanced. See, for example, Figures 13 and 14.

[0084] To reiterate, position 1 refers to the position where the counterbalance assemblies 48, 49 are directly locked to the rotary arm 34. See Figure 13. Position 2 is the position where the counterbalance assemblies 48, 49 are neutral (detached from the rotary arm) and can rotate around the shaft 28 on their own bearings. In position 2, the counterbalance assemblies 48, 49 can change each other's RPM (and position) and their RPM relative to the rotary arm 34. Position 3 refers to the position where the two counterbalance assemblies (48, 49) are relocked to the rotary arm 34 via the outer hub 57, and the rpm and angle of the counterbalance assemblies are fixed to the rotary arm. Position 3 occurs at least two positions: 0° or 180°. In this specification, positions 1 and 3 are referred to as locked modes (or locked positions), and position 2 is referred to as floating mode (or free mode).

[0085] Broadly speaking, the desirable balancing process proceeds as follows: Before payload release: Both counterbalances are at position 1. After the payload is released: Both counterbalances are pulled to position 2.

[0086] One counterbalance assembly is pulled into position 3. The magnet of the other counterbalance assembly is activated, creating an angular velocity difference with the rotary arm, causing the counterbalance assembly to begin moving 180° (relative to the rotary arm).

[0087] When the other counterbalance assembly reaches 180°, it is retracted to position 3.

[0088] Those skilled in the art are aware of various ways to move between locked and free modes. A preferred method for moving between modes is to use a male / female key system. For example, as shown in Figures 16 and 17, an inner key 64 can be placed on the counterweight assembly and engaged with an inner keyway 67 located on the rotor hub 66. Thus, when the inner key 64 engages with the inner keyway 67, the counterbalance assembly is "locked" to the rotary arm 34 (position 1). Similarly, as shown in Figures 18 and 19, an outer key 69 can be placed on the counterweight assembly and engaged with an outer keyway 73 located on the inner surface of the outer hub 57. Thus, when the outer key 69 engages with the outer keyway 69, the counterbalance assembly is "locked" to the outer hub 57 (position 3). Between position 1 and position 3 is position 2, in which the counterweight assembly is not engaged with either the rotary arm 34 or the outer hub 57.

[0089] As shown in Figure 15, a preferred counterweight assembly comprises the following components: an outer hub 57, a button 58, a thrust flange 59, a bearing housing 78 (including bearings not shown), and an inner hub 79. The inner hub 79 is fixed to the rotor hub 66, and the outer hub 57 is fixed to the inner hub 79. The outer hub 57 preferably has a notch so as to form a slot 55 when the outer hub 57 is connected to the inner hub 66. The bearing housing 68 is slidably and rotatably fitted inside the inner hub 79 / outer hub 57 combination. Counterweight arms 54, 56 are connected to the bearing housing.

[0090] As shown in Figure 17, slot 55 restricts the movement of counterweight arms 54 and 56. This means that when the counterweight assembly is in position 2, the counterweight can rotate freely relative to the rotary arm, but its degree of rotational freedom is restricted by slot 55.

[0091] A stepping motor 88 moves the counterweight assembly (48, 49) between first, second, and third positions via a rod. The rod connects the stepping motor 88 to the counterweight assembly (48, 49). More specifically, the rod is connected to a button 58. The button 58 is connected to a thrust flange 59, which engages with a bearing flange 78. In this configuration, the stepping motor 88 acts linearly on the rod, pushing the inner key 64 into the inner keyway 67 and then out. Similarly, the stepping motor 88 acts linearly on the rod, pushing the outer key 69 into the outer keyway 73 and then out.

[0092] To achieve position 1, the stepping motor 88 pushes the inner key 64 into the inner keyway 67, thereby locking the counterweight assembly to the rotary arm 34. This applies to both counterweight assemblies (48, 49).

[0093] To achieve position 2, the stepper motor 88 pulls the inner key 64 out of the inner slot 67, thereby allowing the bearing flange 78 / counterweight arm (54, 56) to rotate within slot 55. This applies to both counterweight assemblies (48, 49).

[0094] To move from position 2 to position 3, one of the two counterbalance electromagnets 61, 62 is activated. When a magnetic field is generated, the rotation of the counterbalance assembly begins to slow down relative to the rotary arm 34. To facilitate the formation of a magnetic couple, it is preferable to add a ferromagnetic ring 75 to the counterbalance assembly. The magnetic coupling slows down the rotation of the counterweight assembly as it passes. This allows the rotary arm 34 to "catch up" with the counterbalance, enabling a 180° adjustment. When the counterbalance assembly reaches a position 180° relative to the rotary arm (opposite the slot 55), it stops. The stepping motor 88 pulls the outer key 69 into the outer keyway 73 and locks it into the outer hub 57. This outer hub 57 locks into the rotary arm 34. This is position 3.

[0095] Under normal circumstances, the other counterweight does not rotate relative to the rotary arm 34. Instead, after the payload is released, the stepping motor 88 pulls the bearing hub 68 directly into position 3. Once both counterweight assemblies 48, 49 are in position 3 (but on opposite sides of each other), the system is again in equilibrium. See, for example, Figure 14. The loads from the two counterweights are now borne on their respective bearings because the counterweight assemblies have moved away from post position 1 of the rotary arm 34.

[0096] Ultimately, three elements are involved in the release of each payload: the payload and two counterweights. Before launch, both counterweights react to the payload. After launch, the counterbalance assembly is detached from the rotary arm 34, and one counterweight is dragged to the other to balance the system. The counterweights re-engage the rotary arm 34 when a new payload is loaded into the relay station. Example: Launching operation

[0097] First, when the rotary arm 34 is in the vertical, lowered position, the payload is loaded into the relay chamber 35 of the first spoke 40. When the payload is inside the chamber, both counterweight assemblies are preferably connected to the second spoke 47 (i.e., the two counterweights are preferably connected to the spoke opposite the payload). Keys 64 mounted inside the counterbalance assemblies maintain the connection of the counterweight assemblies to the rotary arm. The inertia of the first spoke (with the payload) is preferably balanced with the inertia of the second spoke (with the two opposing sets of counterweight assemblies).

[0098] Next, the launch angle is selected. To do this, the satellite bus or momentum wheel rotates the device 10 to the predetermined launch angle. The motor 65 moves the rotary arm 34 to the desired angular velocity, and the payload is ready for launch. The control system issues a launch signal and simultaneously turns on the color magnet 62. The rotary arm 34 continues to rotate until the color 68 is activated. Next, three actions coincide with each other: (1) the payload is released, (2) the first counterweight arm is released from the rotary arm (position 2), and (3) the second counterweight arm is released from the rotary arm (position 2). This keeps the rotary arm balanced. At this point, the first and second counterbalance arms have their respective counterweight forces relative to their respective bearings. Next, one of the counterweight magnets 60, 61 is activated, and one of the counterweight assemblies begins to decelerate. The other counterweight assembly moves to position 3. When the first counterweight rotates 180°, it also moves to position 3. Once one of the counterweight arms switches to the opposite side, the counterbalance system reaches equilibrium, and the rotary arm can begin to decelerate using the regenerative braking system. Item: Catch operation

[0099] Before catching, the rotary arm 34 is preferably balanced to minimize vibration at the start of rotation. A preferred way to achieve this is to "lock" the first counterweight assembly onto the first spoke and the second counterweight assembly onto the second spoke (position 3). In this configuration, both counterweight assemblies are connected to the main rotor via the rotary arm and are located on opposite sides of the rotary arm. In this configuration, the rotary arm is in equilibrium before catching.

[0100] Next, the motor 65 rotates the rotary arm 34 until its tip speed is parallel to the speed of the incoming payload. Preferably, the tip of the rotary arm rotates at a speed approximately 1% slower than the incoming payload, so that the payload can maintain contact with the rotary arm 34 after catching. This slight speed difference mitigates destructive impact and allows for a smoother catch.

[0101] When an incoming payload arrives, it makes contact with the rotary arm 34 and pushes it in. Once contact is established, one counterweight assembly preferably rotates 180° to the other side of the rotary arm to maintain balance. This is preferably achieved by a magnet. This causes one counterweight to move to the opposite side as the payload pushes the rotary arm 34 in a circular motion. After the catch, both counterweights are positioned in parallel, balancing the rotary arm 34 directly opposite the payload. Capture of kinetic energy

[0102] When the payload is caught, its kinetic energy is converted into electrical energy via regenerative braking. To achieve this, the rotary arm 34 is directly mounted to the DDMB LDC motor. Preferably, the motor 65 moves along a fixed shaft that is held in place by the receiving module. The motor itself rotates, but the output shaft does not. This reduces the number of bearings required in the system, improves efficiency, and reduces wear. This also makes it possible to mount a momentum storage mechanism 94 that rotates in the opposite direction to the rotary shaft. See Figure 26.

[0103] As the arm rotates due to the payload's energy, the motor rotates. The motor recovers its kinetic energy while decelerating the arm. For a 12kg payload (when recovering kinetic energy from the launch arm and payload), the required input energy is 697kJ, but the net input energy is only 325.7kJ. The system's efficiency depends on the specific electrical components used and the amount of inertia required for spin-up and spin-down relative to the payload weight.

[0104] To transport the mass, the system uses electrical energy to rotate the DDM and launch the payload. The launch device can utilize regenerative braking to recover kinetic energy from the arm using the DDM. The launch rotary arm may eventually come to a complete stop. If the constellation is designed so that multiple relay stations sequentially reach the same rendezvous point, the first relay station can transport the payload to the second relay station, which can then change direction while keeping its arm rotating to target the third relay station and release the payload, thereby minimizing the amount of energy required to transport the payload and allowing it to continue until it reaches the desired position.

[0105] The receiving device uses stored electrical energy to rotate its catching arm. Once the payload is caught, the kinetic energy of the payload and the receiving device is converted into electrical energy using regenerative braking and stored inside the aircraft.

[0106] The table in Figure 49 provides examples of various rotary arm lengths and the pure mechanical energy required to rotate the payload. Rigid folding / telescopic rotary arm

[0107] An optional embodiment is the use of a foldable (or retractable) rotary arm. Figures 41 and 42 show an embodiment of a relay station with a retractable rotary arm 154 in a folded state. Figures 43–45 show the same embodiment with the arm extended. Telescopic arms reduce the size of the satellite and make it easier to pack. While foldable or telescopic arms can save space, it is preferable for the arm to maintain rigidity after deployment to mitigate the risks associated with non-rigid structures (i.e., non-rigid arms like tethers can become entangled or immobile). It also allows for more aggressive RPM changes, enabling faster launch or catch and higher RPMs. By rotating the rotary arm on its own bearings, the rotary arm can rotate freely, while a momentum storage mechanism keeps the spacecraft under control and prevents it from rotating. Because a preferred relay station has its own momentum storage system and a rigid rotary arm, the relay station has built-in flexibility regarding the attachment of the relay station to a satellite or other space platform. For example, a relay station can be mounted as a payload attachment on a standard satellite or space station / platform. Furthermore, it can be installed to allow for quick ejection in the event of a system anomaly. Because the relay station is inherently independent, it reduces risks during operation, as the spacecraft does not need to be designed around the relay station for successful operation. Optional Shroud

[0108] Figure 46 shows an embodiment of a relay station with an optional shroud 156 to protect the rotary system from orbital debris and cosmic dust. When deployed in space, the shroud can cover the launch arm. The port can be opened during launch and catch operations, while the shroud remains closed at all other times to prevent debris from entering the relay station. This extends the system's service life and prevents the accumulation of residue that could adversely affect bearing performance. Momentum Wheel

[0109] Figure 26 shows a preferred embodiment of the momentum storage wheel 94 ("Momentum Wheel"). The purpose of the momentum wheel 94 is to prevent the relay station from rotating in space in response to the acceleration and deceleration of the rotary arm. Typically, the momentum wheel rotates in the opposite direction to the rotary arm 34 to prevent the relay station / relay satellite from rotating uncontrollably.

[0110] A preferred momentum wheel, like a rotary arm, rotates around a main shaft 28. The main shaft 28 is preferably fixed between a pair of frames 31. See Figure 26. As shown in Figures 28-31, the momentum wheel mainly comprises two elements: a hub 206 and a weight 222. Both the hub 206 and the weight 222 rotate around the main shaft 28. A preferred hub 206 comprises an outer hub 215, an inner hub 221, a bearing 216, a motor 217, a collar 218, a rotary encoder 219, and a rotor sleeve 220. The weight 222 is attached to the hub 206.

[0111] The outer hub 215 and the inner hub 221 form a single housing, which accommodates the elements of the hub 206. The innermost element of the hub 206 is the bearing 216. The bearing 216 is housed inside the collar 218. The collar 218 engages with the rotor sleeve 220. The motor 217 drives the momentum wheel 94 and is mounted between the rotor sleeve 220 and the inner surface of the inner hub 221.

[0112] The motor 217 is preferably a direct-drive, frameless BLDC motor capable of rotating the momentum wheel in both clockwise and counterclockwise directions. The outer hub 215 holds the motor 217, and the inner hub 221 holds the rotary encoder 219. The rotary encoder 219 monitors the position and rpm of the momentum wheel. The stator 231 of the motor 217 is rigidly mounted to the main shaft 28 via a collar 218 and does not rotate. The rotor 230 of the motor 217 rotates the hub 206. The rotor sleeve 220 holds the rotor in place and is fixed to the inner hub 221. As the rotor 230 rotates, the hub 206 rotates. The hub 206 is mounted to a bearing 216. The bearing 216 is mounted to the main shaft 28 and supports the load while the motor 217 drives the rotation. A weight 222 is mounted to the hub 206 and rotates with the hub 206. Weight 222 increases the overall inertia of the wheel, increasing the momentum per RPM. Weight 222 is often heavy relative to the hub 206 and is mounted outside the hub 206. Typically, there are often at least two momentum wheels, meaning that each momentum wheel preferably has half the combined inertia of the rotary arm and the payload. Preferably, the momentum wheels have at least three times the combined inertia of the rotary arm and the maximum assumed payload, so that the momentum wheels do not rotate at the same speed as the rotary arm, which is to match the angular momentum and cancel it out in real time, so that the aircraft is affected to a net-zero effect.

[0113] In a typical launch operation, it is preferable that both the rotary arm and the momentum wheel begin rotating simultaneously, but in opposite directions. The intention is to ensure that the angular momentum of the wheel cancels out the angular momentum of the rotary arm in real time. During operation, a difference in momentum occurs, and spin is generated as the system approaches a steady state. This is acceptable as long as the spin imparted to the spacecraft remains within acceptable limits. When the capsule is released from the rotary arm, both angular and linear momentum are introduced into the system. The momentum wheel (or multiple wheels) absorbs the added angular momentum, stabilizing the relay station. As mentioned earlier, the rotary arm is stabilized by the operation of a counterbalance.

[0114] While a single momentum wheel 94 may function independently, it is preferable to position a main pair 202 of momentum wheels on either side of the rotary arm, as shown in Figure 26. The main pair is preferably configured to absorb most of the angular momentum generated during firing and receiving operations. As a result, the main pair 202 preferably has significantly greater inertia than the rotary arm so that any deviation in motor control has only a minor effect.

[0115] Optionally, a secondary pair 204 of momentum wheels can be added to both sides of the rotary arm. See, for example, Figure 47. The secondary pair of momentum wheels is preferably a smaller version of the momentum wheel 94. In the case of a small relay satellite, the primary momentum wheel can provide 600 Nms of momentum storage, and the secondary momentum wheel can provide 12 Nms of momentum storage. The secondary pair 204 works in conjunction with the primary pair to maintain high pointing accuracy of the relay station during operation. Optionally, the secondary pair may include a high-precision command moment gyroscope (CMG). This type of CMG can control the rotation of the satellite with an accuracy of 0.8 arcseconds.

[0116] Once the rotary arm reaches a stable state, the need for minute acceleration changes in the momentum wheel increases. However, the rotary arm's motor RPM can fluctuate. While these fluctuations are detected and corrected, the spacecraft may enter a rotational state before the corrective operation can begin. To counteract this, the momentum wheel can have significantly large inertia, thereby making motor fluctuations incomparably negligible.

[0117] Subsequently, the rotary arm can be decelerated. The momentum wheel can also be decelerated. However, compared to launch (with the capsule inside the chamber), the inertia of the rotary arm is significantly reduced (without the capsule inside the chamber), so even when the rotary arm approaches 0 RPM, the momentum storage mechanism does not approach 0 RPM.

[0118] If there are concerns about the momentum storage mechanism continuously switching the direction of rotation, a primary pair of more than one momentum wheel can be used. For example, primary pair 1 (control moment gyroscope 1) can operate only in the range of 100 RPM to 1,000 RPM, and primary pair 2 can operate only in the range of -100 RPM to (-1,000 RPM). In the first launch, primary pair 1 starts at 100 RPM and ends at 300 RPM. At the start of the second launch, primary pair 1 operates until it reaches 100 RPM, and then primary pair 2 takes over. This can extend the lifespan of the momentum storage mechanism.

[0119] Command moment gyroscopes may only have a maximum storage capacity of 12 Nms, but small relay stations require more than 60 times that amount of storage. As shown in Figure 47, the angular momentum during spin-up can be offset by placing two rotary wheels on either side of the rotary arm. Alternating firing rotation

[0120] When a relay station rotates a launch arm in one direction to continuously launch capsules, momentum accumulates over time, eventually requiring a deceleration operation. A typical desaturation operation requires the system to burn propellant in the thrusters. One option to prevent this is to alternate the rotation direction of the rotary arm with each successive launch. For example, clockwise for one launch and counterclockwise for the next. When launch rotations are alternated, the accumulated momentum is removed every two launches. For example, if the capsule mass and release RPM are exactly the same, the rotary arm and momentum storage mechanism will decelerate and both come to a complete stop at 0 RPM. Nevertheless, over time, the accumulation of launch inaccuracies is likely to necessitate a conventional desaturation operation.

[0121] Another embodiment involves adding a second rotary arm on which the capsule is mounted. If the second rotary arm rotates in the opposite direction to the first rotary arm, the second rotary arm can cancel out the momentum of the first rotary arm. In this case, the second rotary arm functions as a primary momentum storage wheel (i.e., canceling out the momentum induced by the first rotary arm). Because the second rotary arm cancels out the momentum generated by the first rotary arm, the primary pair 202 of the momentum wheel 94 can freely function as a secondary pair 204 of the momentum wheel, allowing for more precise control of the orientation of the relay station at launch. This alternative embodiment reduces the overall mass of the system while providing redundancy in the release mechanism.

[0122] By integrating a regenerative braking system into the momentum storage mechanism, energy generated from the rotation of the reaction wheels can be recovered. When not actively used in launch and catch operations, the momentum wheels can be rotated to function as kinetic energy batteries, and by using regenerative braking to decelerate them as needed, they can provide auxiliary power during launch in power-constrained environments. For example, consider a scenario where our rotary arm is at 0 RPM, our momentum storage wheels are rotating clockwise at 400 RPM, and our spacecraft itself is not rotating in a steady state. The rotary arm can be rotated clockwise, and as the angular velocity of the rotary arm increases, the angular velocity of the momentum storage wheels can be proportionally decreased so that the spacecraft does not rotate. When the rotary arm is at 200 RPM, the momentum storage wheels may have decelerated to 200 RPM. This allows for the generation of the power necessary for the rotation of the rotary arm while maintaining angular momentum balance. Satellite bus

[0123] The relay station should be able to move in space to catch a drifting or failed launch capsule. One embodiment involves placing thrusters on all sides of the relay station to allow the relay station to shift horizontally and vertically to catch an approaching capsule. The thrusters may be located directly on the relay station or on the satellite bus to which the relay station is mounted. The satellite bus should have its own attitude control system that can be used to provide the directional precision required for the operation of the relay station. The attitude control system works in conjunction with momentum storage wheels to provide the precise directional precision required for normal operation. Since the satellite bus is exposed to high vibration spikes during capsule release or catching, the specially designed mounts that interface between the device and the host platform should be designed to withstand these spikes throughout their entire service life.

[0124] It is crucial that the relay station's position can be changed by even 1 to 25 centimeters. The satellite bus's ability to compensate for these positional discrepancies depends on the specific bus selected and its response time. A feasibility study was conducted to evaluate this capability in a low Earth orbit (LEO) scenario, highlighting the acceleration (0.4 meters per second) required for the bus to effectively respond to the control system's commands. In geostationary orbit (GEO) catches, the repositioning challenges were found to be somewhat mitigated compared to LEO catches due to the relatively lower inertial velocities involved. Furthermore, the simulations considered out-of-plane scenarios as the launcher approaches the relay station, exploring various approach angles to ensure broad applicability of the system. Node operation / Unit-level startup

[0125] One of the advantages of the relay station is that it only needs to accelerate the capsule via the relay station, eliminating the need to accelerate the chassis of a large transport vehicle or the large amount of chemical propellant it carries. The relay station does this by rotating a rotary arm with a payload attached to the end of the arm. This invention significantly reduces the cost of rail logistics operations because the rotary arm only launches the payload (no large transport vehicle is required). As a result, relatively little propellant is needed to compensate for the linear momentum transfer. In other words, when the capsule is accelerated and released using the rotary arm, the momentum transferred to the relay station upon release is equivalent to that of the capsule being launched (the entire mass of the relay station is not accelerated).

[0126] One or more momentum wheels provide electrical energy to resist the rotation of the rotary arm, enabling control of the relay station. The launch process of the relay station capsule can be driven by a solar array or nuclear energy, eliminating reliance on chemical propellants. However, if correction is desired, chemical propulsion may still be required for linear momentum transfer acting on the relay station in a single capsule release in a particular direction.

[0127] By operating at the unit level, a relay station can launch a capsule that deploys anywhere within its operational volume within the maximum delta-v range it provides. Figure 48 shows the reach of a 100 m / s relay station compared to a 350 m / s relay station. Because relay stations do not need to launch at their maximum DV, customers can place their payloads and relay stations on their own separate orbits as needed. For orbits that do not immediately fall within the relay station's range due to their orbital shape, the relay station can wait for the payload to arrive until a node crossing occurs that brings the desired orbit within range. Relay stations can also adjust their orbits to reach an acceptable launch site earlier and place the payload on the desired final orbit. Architecture Overview

[0128] Considering the required mass and capital requirements, a single pair of relay stations to directly propel a capsule, for example, from an LEO to a GEO, or from an LEO to the Moon, is not ideal. A preferred system would use multiple relay stations ("nodes") to accelerate the capsule over vast distances, with the relay stations progressively moving the capsule to different orbits. This type of system enables a distributed node architecture that allows assets to be continuously moved up and down a gravity well between LEOs and GEOs. Furthermore, this system would allow assets to be moved from Earth orbit to Lunar orbit. Those skilled in the art will understand that this could be used to move around or between any celestial body. This architectural design minimizes the individual size of each relay station, increasing mission flexibility.

[0129] The node architecture of the relay station connecting the LEO and GEO can be constructed as shown in Figure 32. This embodiment can be used to support operations around or between other celestial bodies. The circular-elliptical constellation 350 is a preferred embodiment for launch and retrieval operations, and launch and limited catch operations. This embodiment extends from the LEO to the GEO, allowing goods to be transported to higher orbits. Using multiple nodes allows for a greater overall delta-v separation from the original launch position than using a single node alone. The payload can be transported from the bottom node to the top node, or to any point within the entire constellation.

[0130] One of the keys to this system is creating "complementary orbits." In this specification, the term "complementary orbit" refers to the relationship between orbits between different nodes in a constellation. In this specification, "node" refers to a single relay station, which is preferably mounted on a satellite bus. The shape of the orbits does not determine whether two orbits are complementary. The determining factor is the "delta-v separation" between two particular nodes. In this specification, "delta-v separation" refers to the change in velocity required for a capsule launched from the first node 300 to reach (cross) the second node 302 and for the relative velocity between the second node and the capsule to become zero.

[0131] There are three preferred methods for capsule transport: launch and recovery, launch and catch, and launch and limited catch. In the three methods listed above and described below, the orbital altitude of the first node 300 is lower than that of the second node 302.

[0132] Figure 33 shows a circular-elliptical-circular constellation with three nodes in complementary orbits. The first device (or “node”) 300 for launching and receiving the payload is in the first orbit 301. The second device (or “node”) 302 is in the second orbit 303. The third device (or “node”) 306 is in the third orbit 307. In this example, all three devices orbit the same celestial body 310. The first and third orbits are circular, while the second orbit is elliptical. Of note is the perigee 304 of the second orbit. This is where the first and second orbits intersect, and where we can perform orbital transfers between the first and second devices. For launch and recovery (described later), this transfer is performed when the first and second devices are adjacent to each other (depending on the specific orbits, which can be within 100 km of each other). Similarly, this transition may also occur at the second orbital apogee 305 (the point where the second and third orbits intersect). Preferably, each device (or “node”) in this section is a relay station 10 mounted on the satellite bus 150.

[0133] The supplementary configuration in Figure 33 can be used for launch and catch, launch and limited catch, and launch and recovery (without catch). The elliptical second orbit acts as a bridge between the first and third orbits. In a Hohmann transfer, which is the most efficient way to ascend or descend altitude, two burn-outs are typically performed, starting from a circular orbit. The first burn-out changes the orbit from a circular to an elliptical orbit. After the aircraft moves 180° along its orbit, it performs a second burn-out, ascending until its perigee coincides with its apogee, thereby making the orbit circular again. Here, the second orbit functions as the first burn-out because it is formed in an ellipse, and therefore, when it makes a transition, either ascending to the third orbit or descending to the first orbit, it performs a second burn-out and the payload's orbit becomes circular.

[0134] Figure 34 shows a circular constellation 355. This embodiment functions in launch and catch operations. In this embodiment, the trajectory of each node is circular. To perform the transfer, the system is configured to launch and catch at equal delta-v, so as to circularize the payload between two nodes rather than the three nodes required for the circular-elliptical-circular constellation described in the previous paragraph. This example extends from Leo to GEO, allowing the goods to flow to higher (or lower) trajectories. Multiple nodes increase the overall delta-V separation from the original launch position. The payload can flow from the lower node to the upper node, or anywhere in between.

[0135] Figure 35 shows another circular constellation. This embodiment has a first device (or “node”) for launching and receiving the payload in a first orbit 316. A second device (or “node”) 317 is in a second orbit 318. A third device (or “node”) 319 is in a third orbit 320. All three orbits (316, 318, and 320) are circular. And all three of these devices orbit the celestial body 310. In this embodiment, the capsule's orbit is circularized by only two relay stations, as it is desirable for the capsule's launch velocity to match the catch velocity. Once the capsule is deployed from the launch relay station, it is placed into an elliptical orbit so that its orbit intersects the receive orbit. When the capsule is caught, the act of catching the capsule imparts a delta-v to the capsule, making the capsule's orbit circular. This constellation shows complementary orbits in cases where the launch capsule's velocity and release angle cause its orbit to intersect a higher (or lower) circular orbit.

[0136] For example, if a relay station can launch a payload at a maximum speed of 200 m / s, the receiving relay station cannot be more than 200 m / s away. Otherwise, the receiving relay station cannot cross paths with the payload and retrieve it. Alternatively, if the receiving relay station can "catch" the payload at 200 m / s, the delta-v separation could be 400 m / s. The amount of available delta-v is not the only limiting factor. The catch speed determines the relative velocity between the capsule and the node at the time of crossing. The launch geometry must be such that the capsule can physically cross paths with the receiving node.

[0137] Figure 36 shows a circular to elliptical constellation 364 that enables launch and recovery, as well as launch and limited catch. This takes things a step further by showing that the inclination angle changes once the capsule reaches its orbit. In this case, the delta-v separation is based on the plane change the capsule undergoes as it moves from an Earth polar orbit to an Earth equatorial orbit. If the launch and receiving relay stations are close together, the capsule can be transported between the relay stations. In this case, since this final transport results only in a plane change and not a change in altitude of the capsule, the periods of both relay stations are the same, meaning there is an opportunity for transport between stations every half rotation.

[0138] The final polar orbit 360 is the last orbit before initiating our aircraft modification. This orbit is circular and at a polar inclination. In the next launch along the chain, the capsule will move to an inclined orbit. The first inclined orbit 361 is a circular orbit that is the maximum delta-v separation possible from the final polar orbit 360 in the chain. The first inclined orbit 361 and the subsequent planar change orbit are equiperiodic circles, which allows for the maximum number of node crossings. In this specification, “node crossing” refers to a location where two nodes have an orbital shape that allows for the transfer of the capsule between the two orbits. These arrangements are based on the maximum delta-v separation allowed from the relay system, which allows for the largest possible change in orbital inclination. Changes in orbital inclination can be made by launch and catch, launch and limited catch, and launch and recovery.

[0139] The intersection of all inclined tracks 362 is achieved by arranging all inclined tracks to pass through precise times and positions, thereby enabling rendezvous points for multiple relay stations 362. To prevent all capsules from arriving simultaneously, it is preferable to have a separation of at least 50 km between each node. Alternatively, true anomalies of the nodes can be spaced so that each node arrives at the rendezvous point sequentially without the risk of the nodes being in the same place and time during the operation. Because the distance / time change is small, the payload velocity vector is hardly affected, and the risk of relay station collisions is also minimized.

[0140] In all these examples, the crucial step is that the orbit of the capsule launched by the first relay station intersects with the orbit of the second relay station. Thus, the system operates in two ways: (A) when the two nodes intersect (e.g., in the launch and retrieval embodiments described below), and (B) when the two nodes do not intersect but are separated by a suitable delta-v. The key is that the orbit of the launched capsule can intersect with the orbit of the receiving satellite. Launch and recovery

[0141] Figure 37 illustrates the launch and retrieval operation. As shown, the first node 300 is on the circular first orbit 301. The second node 302 is on the elliptical second orbit 303. The first node 300 preferably launches the capsule at the capsule launch position 335, which is the perigee of the second orbit 303. In this case, it is preferable to place the two nodes as close to each other as possible to reduce the possibility of perturbations affecting the capsule orbit 330. The second node 302 intersects the capsule orbit 330, and the second node retrieves the capsule.

[0142] In a typical launch and retrieval operation, the first relay station on the first node 300 rotates to a desired angular velocity and then launches the capsule. The relative velocity between the first node 300 and the second node 302 is, for example, 200 m / s. The first relay station launches the capsule at a first velocity, for example 198 m / s, which is approximately the same as the velocity of the second node but slightly slower. Because the first orbit 301 and the second orbit 303 intersect, the second node slowly overtakes the capsule and retrieves it. During retrieval, the second node (or the capsule) can make slight course corrections.

[0143] For retrieval, a conventional robotic arm connected to a second node can be used to collect nearby capsules and store them at the second node. The second node may also have a cargo bay that opens to directly receive the assets when they intersect with the second node. At the second node, minimal operation may be required to allow the assets to be transported. Of course, this retrieval method can be used at any node.

[0144] This method allows the first node to launch capsules rapidly and continuously without confirmation of catch / receipt from the second node. The second node retrieves each capsule as it passes. Once the second node has passed the first node, the first node stops its launch operation and the second node retrieves the launched capsule. Alternatively, the first node can be launched now at 202 m / s so that its capsule catches up to the second node. The second node can then perform the reverse operation to transfer the capsule to the first node. This method allows for the transfer of more payloads during orbital coupling than with the launch and catch operation (described later).

[0145] To facilitate this type of transport, the nodes need to be in complementary orbits. As they approach Earth, these orbits must maintain the constellation's integrity over long periods (1-3 years) without excessive station maintenance requirements, so nodal and apsidal precession must be considered. To do this, the constellation can be placed in either equatorial or polar orbits. The first node is in a circular orbit, the second node in an elliptical orbit, with its perigee intersecting the first node and its apogee intersecting a third node in a circular orbit. This pattern continues until we reach the desired position. The elliptical orbit precesses around its phelps line, which is aligned with the circular orbit. This means ensuring that the nodes can be within close range for operation while maintaining the constellation's integrity. This makes it possible to perform launches with short standoffs, for example, from 100 km. This significantly reduces the amount of perturbation that could exert work on the capsule. Both nodes in a pair use their own onboard SDA payloads to measure each other's position and orientation in real time and track the asset(s) flying between them. This allows for the rapid launch of many desired capsules, as there is no need to wait for the payload to be captured before launching the next capsule. Launch and catch

[0146] Figure 38 illustrates the launch and catch operations. The first node 315 orbits the celestial body 310 in a first circular orbit 316. The second node 317 also orbits the celestial body 310 in a second circular orbit 318. The capsule is launched from the first node 315 at the capsule launch position 340 at a first velocity and first direction along the capsule orbit 342. In this embodiment, the capsule is launched from the circular first orbit 316 and is positioned on an elliptical orbit (capsule orbit 342) that intersects the second circular orbit 318, allowing the capsule to rendezvous with the second node 317. The capsule launch position 340 is the perigee of the capsule.

[0147] As shown, the capsule trajectory 342 intersects with the second circular trajectory 318 and the capsule catch position 341. This embodiment shows two circular trajectories and a capsule moving between them. This is not an ideal embodiment because the capsule is not caught at 180° from the launcher, reducing the efficiency of transport. However, this demonstrates how the system can catch within a wider error band (e.g., 30° up and down) relative to the trajectory, allowing the system to perform further transport over time, as with the limited catch described below. The relative velocity of the catch will not be the same as the launch velocity, and the catch velocity vector will deviate from the prograde direction because the tangential velocity of the arm tip must have the same velocity vector as the arriving capsule to ensure a non-destructive catch.

[0148] In a typical launch and recovery operation, the first node rotates to launch the capsule. It is desirable to have sufficient flight time to allow the second node to make appropriate course corrections as needed. For example, the first node may have a relative velocity of 400 m / s to the second node. The first node can launch the capsule at an exit velocity of 200 m / s. The capsule has a relative velocity difference of 200 m / s between itself and the second node. The capsule is preferably launched at a first velocity and in a first direction such that its trajectory intersects the second trajectory. In this way, the second node can catch up to and capture the capsule. The node architecture may be similar to the circular / elliptical orbit approach described in launch and recovery. Alternatively, the constellation may consist solely of purely circular orbits. In this configuration, a node launches a capsule at a node intersection, and after half a revolution of the capsule's flight time, the capsule is caught by a receiving unit at a position 180° away from the launch point, resulting in two impulses similar to a Hohmann transition.

[0149] As catching is imminent, the receiving node catches the capsule, as described in the relay station overview, and the rotating arm is adjusted to match the speed of the arriving capsule. During catching, the capsule's momentum is transferred to the relay station. Angular momentum is released to the momentum storage wheel. Following the same principle as launch, the catching relay station can catch in both clockwise and counterclockwise directions, thereby counteracting saturation of the momentum storage mechanism.

[0150] This method has advantages over the aforementioned launch and recovery method because, since the launch speed matches the catch speed, one relay station pair can have a total delta-v separation of 400 m / s. In the launch and recovery system, three relay stations are required to achieve equivalent separation.

[0151] In the launch and catch embodiments, the catcher's operational readiness time is a critical factor in ensuring secure engagement with the capsule. The launch and catch embodiments are designed to be flexible in terms of power consumption and operating speed. Under conditions where maximum power is available, the catcher can reach operational readiness in as little as 17 seconds. Conversely, even in scenarios where power consumption needs to be limited, the embodiments still function effectively, achieving operational readiness within 2 minutes. A key part of this is the power limitation of the relay station. A key feature of this particular embodiment is its detection range capability. The design aims for an ideal detection range of 100 km, while specifying a minimum detection range of 50 km. This range specification is essential to ensure that the link budget between the capsule and the receiving unit is closed. Considering a relative velocity of 200 m / s and a standoff distance of 50 km, the embodiments are designed to provide a reaction time of approximately 250 seconds before catching. This reaction time is essential for aligning the catcher with the projectile's trajectory, enabling precise and successful interaction between the capsule and the relay station. Launch and limited catch

[0152] The launch and limited catch method is similar to the launch and catch method described above, but with one notable difference. The maximum catch speed of the second relay station is limited to a relative speed lower than the exit speed from the first relay station. For example, the first node may launch at 200 m / s, while the second node may catch at 50 m / s. The maximum delta-v separation between the pairs in this example is 250 m / s. This allows for a lower catch speed, thereby reducing the risk of the capsule colliding with or failing to receive the receiving unit.

[0153] Figure 39 illustrates the launch and limited catch operation. As shown, the first node 300 is on the circular first orbit 301. The second node 302 is on the elliptical second orbit 303. The first node 300 preferably launches the capsule at position 335, which is the perigee of the second orbit 303. This launch position allows the capsule to ascend most efficiently so that its orbit coincides with the elliptical second orbit 303, starting from the circular first orbit 301. Orbit 337 shows the trajectory of the capsule during its flight time until it reaches the catch position 336 at the second node 302. The catch position 336 is the first intersection of the second orbit 303 and the capsule orbit 337. Before the catch, the capsule's elliptical orbit is slightly larger than the second orbit 303, so that it can have a velocity difference that allows for a rendezvous. This embodiment generates three possible intersections. (1) at the perigee 335 of the second orbit 303, (2) at the catch position 336 where the orbits first intersect, and (3) at the second orbital intersection 338 after the capsule has passed its apogee.

[0154] A preferred embodiment of this operation uses a node architecture that transitions from a circular orbit to an elliptical orbit and back to a circular orbit, allowing for the application of equal, unbalanced burn times to create a circular orbit between the first node (launcher) and the second node (catcher). Optionally, the capsule can use onboard propulsion to compensate for the delta-v required to create a circular orbit before the second node "catches" it, thereby enabling a fully circular orbit node architecture.

[0155] In a typical launch and limited catch operation, the first relay station rotates to launch the capsule. In this operation, there are two transport positions: a catch position immediately adjacent to both nodes when the travel distance is short, or when the first relay station launches the capsule when the first and second relay stations are relatively far apart. In this case, the second relay station captures the capsule, preferably less than one rotation. This embodiment allows for the launch of more capsules but requires a relative velocity difference between the capsule and the second relay station. Otherwise, it may take a relatively long time for the capsule to reach and be captured by the second relay station. Preferably, the intersection position is approximately 160° from the launch point. Nodal Summary

[0156] As mentioned above, this system operates in circular and elliptical nodal orbits centered on a single celestial body. However, the system also supports transition orbits between two different celestial bodies (e.g., Earth to the Moon). For a transition to occur, the capsule must be able to rendezvous with the receiving node, and the relative velocity at the time of rendezvous must be within the acceptable window for catch or retrieval. This type of capture can occur between significantly different elliptical orbits, between flower constellations, or between injection orbits between celestial bodies. Crucially, the launched capsule must be able to intersect with the receiving satellite.

[0157] Before launch, retrieval, or capture occurs, both relay stations preferably know their precise position, velocity, and attitude orientation. Depending on the required accuracy, it is best to utilize GNSS (GPS) for the position and velocity of each relay station. If this is not possible, distance data may be obtained by S-band communication, laser communication, or laser ranging.

[0158] Relay stations can also be tracked by ground stations, and each relay station may have an onboard payload that can be used to track other relay stations within range. If a relay station's onboard space domain surveillance system is tracking other relay stations, ground tracking may not be necessary. Relay station constellations can be configured to function autonomously, with relay stations throughout the constellation (not necessarily having to be launch or receive units) maintaining control of the capsule and launch and receive relay stations and reporting their location information. In addition, additional sensors and RF transceivers can be placed in the capsule to provide telemetry to the associated relay stations.

[0159] The relay station can actively track the capsule after launch. Telemetry consisting of state estimates of the capsule and the relay station can be transmitted to the receiving unit and the capsule. The capsule may have an onboard propulsion system that allows for minor orbital adjustments. The receiving unit (or satellite bus) may have thrusters in all directions, which allows for the separation of position and attitude movement. When the capsule is close enough, the catch relay station's SDA system provides inertial position and velocity state estimates for final adjustments.

[0160] Depending on the embodiment, a conventional Hohmann transition can take the form of a capsule launch and recovery corresponding to one combustion, with the subsequent transition corresponding to a second combustion, and the three nodes coordinating to ascend from a circular orbit to a higher (or lower) circular orbit. The specific size and shape of the orbit are not important. Energy recovery

[0161] Each launch and potential catch requires a significant amount of energy. To conserve energy, kinetic energy can be recovered from the arm and momentum storage mechanism. When the capsule is caught, kinetic energy from both the capsule and the arm can be recovered and stored. This energy can then be used for subsequent launches, reducing the amount of power stored required for the launch cycle. Capsule Storage

[0162] After a capsule is caught, it must be quickly stored so as not to interfere with the next catch. The capsules can be stored in a storage space provided on the side of the relay station. The relay station may have magnetic dimples to secure the capsules. The relay station can gently push out a capsule after it has been caught, and then, once all capsules in a particular launch cycle have been caught, the relay station can retrieve them. This method is also applicable to more conventional rendezvous and capsule retrieval (the launch and retrieval described above). To save storage space, the relay station may also place the capsules in an extendable storage rack (extending beyond the relay station's pre-deployment footprint).

[0163] The number of capsules depends on the host platform. Capsules of different sizes can perform different types of operations, providing flexibility and long-term usefulness.

[0164] Optionally, a robotic arm attached to the relay station can be used to pick up and load the capsule like a muzzle-loading weapon. In this concept, the release mechanism of the relay station is retracted and held in place using a magnetic system and a BLDC motor, after which the capsule is inserted and the release mechanism closes. However, this embodiment means that the rotary arm must be completely stopped, loaded, and then rotated again after each capsule is launched.

[0165] Another option is to load the capsules through a hollow shaft. In this embodiment, the rotary arm of the relay station can be loaded while it is in motion, enabling high-speed launch of the capsules and reducing the energy required for the subsequent release of each capsule, thereby reducing the power storage requirement. In this case, a system such as a conveyor can be used to pull the capsules towards the center of the rotary arm. Nodal Offload

[0166] After the capsule is transported to its desired orbit via the node network, it can be offloaded. In some cases, the capsule may be launched directly from a relay station and accelerated to its own desired orbit. For resources that need to be transported to a station or relay station, an interorbital transport vehicle (OTV) may dock with the relay station, and the relay station may transport the resources to the OTV. The OTV can then transport the payload to its final destination. Track maintenance / Self-catch

[0167] Figure 40 shows how launching a capsule from a relay station affects the trajectories of the host vehicle and the capsule itself. The original launch position 370 indicates the point from which the relay station launched the capsule. The relay station and capsule return to this point in each orbit, but not necessarily simultaneously. The relay station trajectory 371 is lower (compared to the relay station's original trajectory 373) due to the momentum transfer given by releasing the capsule. In other words, the momentum transfer of the capsule has a proportional effect on the relay station, pushing it down to a lower trajectory. The capsule trajectory 372, on the other hand, is higher than the original relay station trajectory 373 because it was accelerated when it was launched.

[0168] The act of launching the capsule from the relay station alters the trajectories of both the relay station and the capsule, but both return to their original launch position 370 in each orbit. By utilizing orbital resonance, it is possible to calculate the time it takes for the capsule to reach the rendezvous point (original launch position 370) simultaneously with the launcher. This allows for control over the capsule's mission duration.

[0169] Figure 40 also shows an embodiment in which a single relay station launches a capsule toward itself. When a single relay station launches multiple capsules toward itself, the capsules follow the same trajectory shown here. However, each time a capsule is launched, the relay station's original trajectory and rendezvous position change. This can be easily corrected by adjusting the relay station's trajectory with propulsion when necessary for rendezvous. Another option is to use the momentum transferred at the launch of the payload to adjust the relay satellite's trajectory so that it can respond more appropriately when recovering another capsule in the future.

[0170] As shown in Figure 40, if the capsule is not recovered by the receiving relay station, it will eventually return to its original launch position 370 after each orbit. This is because the payload has been given instantaneous velocity, and its trajectory is tied to the position from which it was launched. Of course, perturbations can cause its trajectory to deviate, and it may not reach the exact same position over time. Ultimately, the launch relay station can rendezvous with and recover the payload at a later point without any intervention. The receiving unit may only need to make a slight maneuver to recover the payload if it is deemed a preferred course of action.

[0171] After the capsule is launched into its desired orbit, it needs to maintain that orbit for the duration of its planned mission. This can be achieved by an on-board propulsion system and propellant. A relay station system can supplement the capsule's required orbital maintenance by performing orbital maintenance tasks. As shown in Figure 40, the capsule and relay station periodically reach a conjunction state (the timing is determined by the relationship between the two orbits). Once coupled, the relay station can catch the capsule and then relaunch it into the appropriate orbit. If the capsule drifts or requires forced coupling, the relay station may need to operate to recover the capsule, but this electromechanical orbital maintenance function reduces the overall mass of the payload. If only one capsule is maintained, if either the payload or the relay station is disturbed and deviates from their original orbits, the entire relay station needs to be accelerated and coupled to the capsule, which is costly.

[0172] This is ideal for long-term operation with mini constellations. For example, in one embodiment, four sensor capsules can be placed "below" it and four capsules "above" it. Each time a relay station joins a capsule, it can pick it up and redeploy it. Node Architecture

[0173] With each launch and catch, the linear momentum transfer element alters the relay station's orbit, creating unique opportunities. Instead of immediately correcting the orbital change, the relay station can utilize this for momentum transfer, enabling faster connection with the next relay station. Using Lambert transfer, the relay station can calculate the most advantageous burn window, improving capsule throughput by increasing the number of orbital rendezvous opportunities. In an ideal scenario, if the up-mass and down-mass passing through the architecture are equal (e.g., moving 10,000 kg of material to the GEO and 10,000 kg to the LEO), the net linear momentum transfer cancels out. This allows for mass transfer without any propellant in space. Operationally, when a relay station launches 200 kg in one direction, it is functionally out of position. The relay station reserves a delta-v budget for the net launch in one direction before using its thrusters to make orbital corrections. In a "Launch & Catch" operation, both the launch and receive units move either above or below the gravity well by the same delta-v in the same relative direction. In a "Launch & Recovery" operation, only the relay station is moved, although the receive unit may move the relay station by a similar amount in a subsequent launch. Relay stations do not necessarily benefit from correcting their trajectory immediately after a launch or catch scenario. A relay station may now be in an orbit that allows it to rendezvous more quickly with the next relay station in the chain, and will later inject fuel to correct its trajectory. This reduces propellant costs while increasing the number of potential rendezvous opportunities between nodes.

[0174] The spacing between nodes in a network can be reduced by using more efficient transports that require less delta-v. However, in this type of transport network, the overall mass that can be transported through the network will be smaller because aggregation takes longer. Similarly, the spacing between nodes can be increased by using less efficient transports that require a higher delta-v, thus reducing the number of nodes required for the constellation. Capsule design:

[0175] The preferred capsule is designed as spherical, which provides a symmetrical shape for catching, regardless of the type of spin applied to the capsule during launch. However, the capsule can be any shape (e.g., disc, egg, cylinder), and only the release mechanism of the relay station needs to be modified to accommodate the different shapes. Since rotation must be considered for catching, and rotation will change the shape of the catch to an irregular shape, a less symmetrical shape is more suitable for operations that do not require catching.

[0176] Some assets transported at relay stations require spherical protective capsules, but many do not. For example, when recycling decommissioned spacecraft, the recovered aluminum can be cast into the desired shape and volume before being loaded into the relay station.

[0177] By utilizing custom-made capsules with thrusters, we can compensate for any launch inaccuracies of the relay stations by allowing for course corrections during flight. This enables each capsule, equipped with its own thruster, to function as a miniature relay station, with its initial acceleration provided by the relay station, but after that, it can operate independently.

[0178] Although the present invention has been described in detail with reference to one or more specific preferred embodiments, it will be understood by those who are ordinary skill in the art to which the invention pertains that various modifications and improvements are possible without departing from the spirit and scope of the claims.

Claims

1. A device for transporting a mass in outer space, wherein the device is A shaft positioned along the central axis and connected between a pair of frames, A first rotary arm connected to the shaft, wherein the first rotary arm is a rigid body and is configured to rotate around the shaft, A first momentum wheel connected to the shaft, wherein the first momentum wheel is configured to rotate around the shaft, A second momentum wheel connected to the shaft, the second momentum wheel being configured to rotate around the shaft, The device is configured to release a mass from the first rotary arm.

2. The device according to claim 1, further comprising third and fourth momentum wheels.

3. The device according to claim 2, wherein the first and second momentum wheels comprise a main pair, and the third and fourth momentum wheels comprise a secondary pair.

4. The device according to claim 3, wherein the moment of inertia of each momentum wheel of the main pair is at least three times the moment of inertia of each momentum wheel of the sub-pair.

5. The device according to claim 4, wherein the main pair is configured to counteract the momentum induced by the first rotary arm, and the sub-pair is configured to rotate the device.

6. The device according to claim 1, wherein the rotary arm is configured to rotate in opposite directions to the first and second momentum wheels.

7. The device according to claim 1, further comprising a second rotary arm, wherein the second rotary arm is configured to rotate in the opposite direction to the first rotary arm.

8. The device according to claim 1, wherein the first rotary arm is configured to rotate both clockwise and counterclockwise, and the first and second momentum wheels are configured to rotate both clockwise and counterclockwise.

9. A device for receiving and launching a mass in outer space, wherein the device is Equipped with a receiver, shaft, and rotary arm, The receiver is configured to receive a mass from a first position in outer space, A device in which the rotary arm is connected to the shaft between a pair of frames, the rotary arm is a rigid body, and is configured to launch the mass to a second position in outer space.

10. The device according to claim 9, wherein the first position is a relay station in a first orbit around a first celestial body.

11. The device according to claim 10, wherein the device is mounted on a satellite bus, and the satellite bus orbits the first celestial body in a second orbit.

12. The device according to claim 11, wherein the second location is a relay station in a third orbit around the first celestial body.

13. The device according to claim 11, wherein the first orbit revolves around a first celestial body, and the second orbit revolves around a second celestial body.

14. The device according to claim 9, wherein the receiver is a mechanical arm configured to grasp the mass.

15. The device according to claim 9, wherein the receiver has a concave surface for catching the mass.

16. The device according to claim 9, wherein the receiver is connected to the rotary arm.

17. The device according to claim 9, wherein the first position is the device itself, and the second position is the device itself.

18. A system for transporting a mass in outer space, wherein the system is A first device for launching and receiving a payload, wherein the first device is in a first spatial orbit, and the first device comprises a rigid arm for launching a first payload in the first orbit, A system comprising: a second device for launching and receiving a payload, the second device being in a second spatial orbit, the second device comprising a first receiver configured to receive the first payload, and the second spatial orbit intersecting the first orbit with the second device.

19. The second device includes a rigid arm for launching the first payload into a second orbit, The system according to claim 18, further comprising: a third device for launching and receiving a payload, the third device being in a third spatial orbit, the third device comprising a second receiver configured to receive the first payload, and the third spatial orbit intersecting the second orbit of the third device.

20. The system according to claim 18, wherein the first spatial orbit revolves around a first celestial body, and the second spatial orbit revolves around the first celestial body.

21. The system according to claim 18, wherein the first spatial orbit revolves around a first celestial body, and the second spatial orbit revolves around a second celestial body.

22. The device according to claim 18, wherein the first receiver is a mechanical arm configured to grasp the mass.

23. The device according to claim 18, wherein the first receiver comprises a concave surface for catching the mass.

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