Spacecraft Heat Shield

Abstract

A deployable spacecraft atmospheric entry heat shield (30, 40, 50, 60) is configured in a flasher pattern having a deformable polyhedral surface having a plurality of sectors (2), each sector (2) having a plurality of mountain folds (4, 5), a plurality of valley folds (6, 7), and a plurality of facets (8) between the folds (4, 5, 6, 7) and an outer edge (12). The heat shield (30, 40, 50, 60) is configured to deploy from a stowed configuration to a deployed configuration.

Description

[Technical field]

[0001] This invention relates to spacecraft heat shields, and more particularly to deployable spacecraft atmospheric entry heat shields. A heat shield is a device for thermal protection and deceleration of a spacecraft as it enters the atmosphere from space. [Background technology]

[0002] Spacecraft heat shields are necessary to prevent excessive heat transfer to a spacecraft during re-entry from Earth orbit or entry into another planetary atmosphere, such as those of Mars or Venus (hereafter referred to as atmospheric entry for simplicity). There are two main methods of heat dissipation used during atmospheric entry: ablation and radiation.

[0003] Ablative heat shields rely on ablation: the evaporation of the heat shield's surface material to allow heat to be carried away in the outflowing gases and not build up within the spacecraft. All currently operational or soon-to-be operational manned entry spacecraft (e.g., Soyuz, Shenzhou, Crew Dragon, Starliner) use ablative-dominant heat shields because of their long tradition and the well-understood physics involved.

[0004] Radiation heat shields rely on electromagnetic radiation, especially in the infrared range (i.e. thermal radiation), to remove heat from the heat shield, and the amount of energy released depends on the temperature of the surface and its ability to release energy. Potential materials for radiant heat shields in shuttle-type configurations are limited due to the high temperatures reached during atmospheric entry, typically exceeding 2000k. The temperature reached by the heat shield is governed by its area-to-mass ratio. Heat flux (MW / m 2The incoming rate of the heat shield is determined primarily by its ballistic coefficient at entry, which is a measure of the heat shield's ability to overcome air resistance during flight and is equal to the mass divided by the drag coefficient per cross-sectional area of ​​the heat shield (kg / m). For simple geometries, the ballistic coefficient is a measure of the aerodynamic drag area-to-total mass ratio.

[0005] For example, one currently operational commercial re-entry vehicle has a ballistic coefficient during re-entry of ~570 kg / m 2 The US Space Shuttle has a large wing area, so the maximum load is ~350kg / m 2 Both of these vehicles typically deliver single-digit megawatts per square meter (MW / m) for tens of seconds during re-entry from low Earth orbit (LEO). 2 ) per kilogram of vehicle. Furthermore, one currently operating commercial reentry vehicle has a heat flux of ~0.002 square meters (m 2 / kg), the US Space Shuttle has a heat shield radiation area of ​​~0.003 m 2 had equivalent values ​​of 0.01 mg / kg.

[0006] Spacecraft may also require deceleration means such as parachutes, umbrellas, and / or heat shields that extend outward from two dimensions of the spacecraft. However, such deceleration means may be bulky. This may be a disadvantage due to limited volume in the spacecraft's launch vehicle or cargo bay. Summary of the Invention

[0007] In view of the above, there is a need for a compact heat shield that can provide protection and deceleration for a spacecraft.

[0008] A first aspect of the invention provides a deployable spacecraft entry heat shield having a deformable polyhedral surface, the spacecraft entry heat shield having: a plurality of sectors, each sector having a plurality of mountain fold lines, a plurality of valley fold lines, a plurality of facets between the fold lines, and an outer edge. The heat shield is configured to deploy from a stowed configuration to a deployed configuration.

[0009] The heat shield can be a radiant heat shield.

[0010] The deformable polyhedral surface may have a polygon central region of order m from which fold lines radiate. The mountain fold lines may include major mountain fold lines and minor mountain fold lines, and the valley fold lines may include major valley fold lines and minor valley fold lines. The major mountain fold lines and major valley fold lines may extend from the polygon central region. The minor mountain fold lines and minor valley fold lines may extend between the major mountain fold lines and between the major valley fold lines.

[0011] The deformable polyhedral surface may have m-th order radial symmetry about a polygonal central region. The sectors may also be radially symmetrical.

[0012] The surface of the heat shield may be formed as a unitary sheet of material.

[0013] The plurality of fold lines in each sector may be configured to substantially limit angular rotation about the longitudinal axis of the heat shield during use.

[0014] The ratio of the polygonal central region diameter to the outer flat-diameter of the unfolded configuration of the deformable polyhedral surface can be 1:11. The outer flat-diameter is the diameter if the deformable polyhedral surface were forced into a single plane, i.e., flat.

[0015] The deformable polyhedral surface can be configured to adopt a swept back form in a deployed configuration such that the center of mass of the heat shield (or spacecraft having the heat shield) is behind the polygon center relative to a longitudinal normal of the heat shield.

[0016] The deployed configuration may be a stable state and a retention means may be configured to maintain the heat shield in the stowed configuration. The retention means may be configured to release when an activation threshold is exceeded.

[0017] The outer edge may be rounded.

[0018] A second aspect of the invention provides a deployable spacecraft atmospheric entry heat shield configured in a flasher pattern. All features of the first aspect may be combined with the second aspect.

[0019] A third aspect of the present invention provides a spacecraft having a heat shield according to any one of the preceding claims. All the features of the first aspect may be combined with the third aspect.

[0020] Advantageously, the heat shield of the above embodiment is deployable in a variety of situations to perform a variety of functions, including acting as a heat shield. The heat shield can be of compact size in a stowed configuration, which is desirable for space applications due to launch volume limitations. The heat shield is also deployable in a deployed configuration to greatly increase its effective area. This allows the heat shield to not only efficiently protect the spacecraft from excess heat, but also to reduce the speed of the spacecraft during atmospheric re-entry. This combination of properties allows a wider range of materials to be effectively used in the manufacture of the heat shield. Other advantages will be apparent to those skilled in the art and are also highlighted herein. [Brief description of the drawings]

[0021] An embodiment of the present invention will now be described with reference to the accompanying drawings.

[0022] [Figure 1A] This is an origami pattern known as the Flasher pattern, which assumes negligible thickness defined by the parameters m=5, r=2, and h=2. [Figure 1B] 1B is a sector of the flasher pattern of FIG. 1A. [Figure 2A] 1B is a thickness accommodating flasher pattern produced by optimizing the flasher pattern of FIG. 1A. [Figure 2B] FIG. 2B is a thickness-corresponding flasher pattern of FIG. 2A, with minor fold lines omitted. [Figure 3A] FIG. 2B is an isometric view of a deployable heat shield based on the flasher pattern of FIG. 2A in a stowed configuration. [Figure 3B] FIG. 3B is an isometric view of the deployable heat shield of FIG. 3A in a partially deployed configuration. [Figure 3C] FIG. 3B is an isometric view of the deployable heat shield of FIG. 3A in a fully deployed configuration. [Figure 4A] FIG. 3B is a top view of the deployable heat shield of FIG. 3A in a stowed configuration. [Figure 4B] FIG. 3B is a top view of the deployable heat shield of FIG. 3A in a partially deployed configuration. [Figure 4C] FIG. 3B is a top view of the deployable heat shield of FIG. 3A in a fully deployed configuration. [Figure 5A] FIG. 3B is a side view of the deployable heat shield of FIG. 3A in a stowed configuration. [Figure 5B] FIG. 3B is a side view of the deployable heat shield of FIG. 3A in a partially deployed configuration. [Figure 5C] FIG. 3B is a side view of the deployable heat shield of FIG. 3A in a fully deployed configuration. [Figure 6A] FIG. 3B is an isometric view of the deployable heat shield of FIG. 3A in a stowed configuration installed on a spacecraft. [Figure 6B] FIG. 3B is an isometric view of the deployable heat shield of FIG. 3A in a partially deployed configuration mounted on a spacecraft. [Figure 6C] FIG. 3B is an isometric view of the deployable heat shield of FIG. 3A in a fully deployed configuration mounted on a spacecraft. [Figure 7A] 13 is a further thickness-enabled flasher pattern having minor fold lines defined and illustrated by parameters m=5, r=1, and h=2. [Figure 7B] This is the thickness-corresponding flasher pattern of FIG. 7A with the minor fold lines omitted. [Figure 8A] 13 is a further thickness enabled flasher pattern having minor fold lines defined and illustrated by parameters m=4, r=3, and h=1. [Figure 8B] This is the thickness-corresponding flasher pattern of FIG. 8A with the minor fold lines omitted. [Figure 9A] 13 is a further thickness-enabled flasher pattern having minor fold lines defined and shown with parameters m=10, r=2, and h=5. [Figure 9B] This is the thickness-corresponding flasher pattern of FIG. 9A with the minor fold lines omitted. [Figure 10A] 13 is a further thickness-enabled flasher pattern with an overlaid dashed circle template and omitting the minor fold lines defined by parameters m=6, r=3, h=2. [Figure 10B] 10B is a thickness-corresponding flasher pattern of FIG. 10A cut into the dashed circle template of FIG. 10A. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0023] Generally, this description describes a deployable spacecraft atmospheric entry heat shield (herein heat shield) that enables protection and deceleration of the spacecraft during atmospheric re-entry. The spacecraft can be any type of space vehicle or vehicle that enters or re-enters the atmosphere of Earth or another planet or planet's moon at high speed, where thermal protection (and deceleration) of the spacecraft is required.

[0024] Origami-like patterns are used to allow the heat shield to be deployed from a compact stowed configuration to a deployed configuration prior to (and / or during) atmospheric entry. Origami and similar folding techniques can efficiently pack large deployable space structures to a size suitable for launch into space. In traditional origami, the thickness of the folded material is usually assumed to be negligible because the material is usually paper and treating the thickness of the paper as negligible is a valid assumption. However, in structural origami, the thickness of the material must be considered to avoid undesirable material deformation and allow for stronger structures.

[0025] FIG. 1A shows an origami pattern known as the Flasher pattern 1, which assumes negligible thickness. This style of origami pattern was originally coined as a "Flasher" by origami artist Jeremy Shafer in his book "Origami to Astonish and Amuse" (St. Martin's Griffin, 2001, ISBN 0-312-25404-0). However, the term "Flasher" is now well understood and used in the art. A Flasher pattern is a specific type of deformable polyhedral surface having multiple sectors. A single sector 2 is shown in FIG. 1B. Each sector has multiple mountain folds 4, 5, multiple valley folds 6, 7, and multiple rigid facets 8 between the folds. The Flasher pattern 1 is shown as having a polygonal central region 10 with mountain folds 4 and valley folds 6 extending outwardly to the periphery. In the case of Figures 1A and 1B, the polygonal central region 10 is a pentagon.

[0026] A Flasher pattern typically has major and minor fold lines, such that there are major mountain fold lines 4, major valley fold lines 6, minor mountain fold lines 5, and minor valley fold lines 7. Major mountain fold lines 4 are shown in bold dashed lines in Figs. 1A-2B. Major valley fold lines 6 are shown in bold dotted lines in Figs. 1A-2B. Minor mountain fold lines 5 are shown in light dashed lines in Figs. 1A and 2A. Minor valley fold lines 7 are shown in light dotted lines in Figs. 1A and 2A. Facets 8 that border the major fold lines substantially abut when the Flasher pattern is in a stowed configuration. In contrast, minor fold lines can be considered "minor" because they fold to a lesser extent, as their purpose is to allow the Flasher pattern to be rolled up in a stowed configuration, as is evident from the variations shown in Figs. 3A-6C. As a result, major fold lines 4, 6 extend from polygonal central region 10 to the outer edges 12 of the Flasher pattern. Minor fold lines 5,7 branch off from and pass between the major fold lines 4,6 to form a complex fold pattern.

[0027] The outer edge 12 has a shape that generally corresponds to the shape of the polygonal central region 10. However, if there are vertices in the polygonal central region 10, there are at least two vertices 13, 14 in a stepped configuration in the outer edge 12. Each sector 2 of the Flasher pattern 1 includes at least two vertices 13, 14 such that one vertex 13 defines an angle of 90 degrees or less and the other vertex 14 defines an obtuse angle (greater than 90 degrees). Within each sector 2, the major fold lines 4, 6 do not extend to at least one vertex 13 of the outer edge 12.

[0028] Flasher patterns are generally designed to move from a stowed configuration to an deployed configuration and vice versa with one theoretical degree of freedom. When a flasher pattern moves from an deployed configuration to a stowed configuration, the sectors wrap around the longitudinal axis of the flasher pattern (for flasher pattern 1, this axis is in the page of FIG. 1A, shown as reference numeral 32 in FIGS. 3C and 5C). Flasher patterns typically have an deployed configuration that is positioned substantially perpendicular to the axis of the polygonal central region. Although the flasher pattern's one theoretical degree of freedom is due to its intended fold line structure, with the application of sufficient force, the flasher may deform in an unintended manner due to material limitations. Stated another way, the fold line pattern of the flasher pattern limits expansion to the deployed configuration such that it can only occur in a single way / method (i.e., unfolded from polygonal central region 10), and deviation / deformation from this single way / method of expansion can occur if a strong enough external force is applied to the flasher pattern (i.e., applying a force that bends the flasher pattern in the deployed configuration across its diameter - this is a deviation / deformation because it is not intended by the fold pattern). The flasher pattern can have a large difference in diameter between the stowed and deployed configurations, typically 1:5.

[0029] As an advantage, flasher patterns are mathematically definable, and therefore can be scaled in size and complexity. A flasher pattern can be defined by parameters m, r and h: "m" is the rotational order of the Flasher pattern, i.e. the number of polygonal sides of the polygonal central region 10 (which is also equal to the number of sectors 2). "r" is the number of rings of the Flasher pattern, i.e., the number of layers of each section that wrap around the polygonal central region 10, or equivalently, the number of times the diagonals (the fold line(s) that run from the polygonal central region 10 through the center of each sector 2) move from bottom to top and vice versa. For example, as shown in Flasher pattern 1, the diagonals run from the polygonal central region 10 first as major mountain fold lines 4 and then as major valley fold lines 6. The bottom and top of the Flasher pattern are visible from the stored configuration (see Figures 3A, 5A, 6A) and are relative to the polygonal central region 10 such that the polygonal central region 10 can be at the top or bottom of the Flasher pattern in the stored configuration depending on the orientation. "h" is the height of each ring of the flasher pattern. The height is in units of facets 8 defined by the minor fold lines 5, 7 (i.e., before the minor fold lines 5, 7 are removed). For example, flasher pattern 1 in FIG. 1A is shown having two rings (i.e., r=2), with two facets 8 of unit height extending between the outer edge 12 and the adjacent ring. Thus, flasher pattern 1 has a height of 2. In another example, flasher pattern 50 in FIG. 8A is shown having three rings (i.e., r=3), with only one facet 8 of unit height extending between the outer edge 12 and the nearest adjacent ring. Thus, flasher pattern 50 has a height of 1.

[0030] For example, flasher pattern 1 in FIG. 1A has parameters m=5, r=2, and h=2.

[0031] To create structural origami (i.e., Flasher patterns with a finite thickness of material), the Flasher patterns are optimized to account for the thickness of the material. The optimization can be performed using the known thickness correspondence theory of origami as follows, which is incorporated herein by reference in its entirety: Zirbel, SA, Lang, RJ, Thomson, MW, Sigel, DA, Walkemeyer, PE, Trease, BP, Magleby, SP, and Howell, LL (October 3, 2013). "Accommodating Thickness in Origami-Based Deployable Arrays." ASME. J. Mech. Des. November 2013; 135(11): 111005; Bowen, L., Trease, B., Frecker, M., & Simpson, T. (2016). “Dynamic modeling and analysis of an origami-inspired optical shield for the starshade spacecraft”. In Multifunctional Materials; Mechanics and Behavior of Active Materials; Integrated System Design and Implementation; Structural Health Monitoring [V001T01A012] (ASME 2016 Conference on Smart Materials, Adaptive Structures and Intelligent Systems, SMASIS 2016; Vol. 1). American Society of Mechanical Engineers; Hossain Bhuiyan, ME, Semer, D, & Trease, BP. "Parametric Studies of Geometric Design Factors on Static and Dynamic Loading of an Origami Flasher." Proceedings of the ASME 2017 Conference on Smart Materials, Adaptive Structures and Intelligent Systems. Volume 1: Development and Characterization of Multifunctional Materials; Mechanics and Behavior of Active Materials; Bioinspired Smart Materials and Systems; Energy Harvesting; Emerging Technologies. Snowbird, Utah, USA. September 18-20, 2017. V001T08A016. ASME.

[0032] Any material thickness can be used as long as the material can be folded without breaking. It is understood that a thickness-optimized flasher pattern still has all of the properties of a flasher pattern, and thus the term flasher pattern can also describe a thickness-optimized flasher pattern. However, to avoid confusion, the term "thickness-enabled flasher pattern" is used herein to describe a thickness-optimized flasher pattern.

[0033] 2A and 2B show a thickness-responsive flasher pattern 20 generated by optimizing the flasher pattern 1 of FIG. 1A (with parameters m=5, r=2, h=2) for use as a deployable spacecraft atmospheric entry heat shield. The thickness-responsive flasher pattern 20 maintains the same underlying characteristics as the flasher pattern 1 of FIG. 1A. Accordingly, corresponding features have been given the same reference numbers. The thickness-responsive flasher pattern 20 can be described as a deformable polyhedral surface. FIG. 2B shows the thickness-responsive flasher pattern 20 without the minor fold lines 5, 7. The configuration of the thickness-responsive flasher pattern 20 without the minor fold lines 5, 7 can be used when the multiple facets 8 are non-rigid / elastic. This is because the minor fold lines 5, 7 are not required to allow the facets 8 to flex as the thickness-responsive flasher pattern 20 moves between a stowed configuration and a deployed configuration.

[0034] The thickness-dependent flasher pattern 20 is also an example of a suitable folding pattern for the heat shield 30 shown in Figures 3A-6C. The properties of the thickness-dependent flasher pattern 20 allow the heat shield 30 to be deployed from the stowed configuration shown in Figures 3A, 4A, 5A, and 6A to the deployed configuration shown in Figures 3C, 4C, 5C, and 6C. The "stowed configuration" refers to when the heat shield 30 is folded such that the surface area / outer diameter is minimized. The "deployed configuration" is the arrangement in which the shield naturally rests in a semi-folded equilibrium position without the application of any external force. This condition is known as the "unloaded" configuration and is determined by several parameters. The nature of the central polygon and the ratio of the flat outer diameter to the inscribed central polygon diameter determine the sweep angle at which the unloaded configuration rests. The rigidity of the shield material and the stiffness of the joints also determine the angle between the plates. The deployed configuration may be the most stable configuration of the heat shield 30 or the stable configuration of the heat shield 30 .

[0035] 3A-3C show the deployable heat shield 30 configured in the form of a flasher pattern, specifically the thickness-dependent flasher pattern 20. FIGS. 3A-3C show the deployable heat shield 30 in isometric views. FIG. 3A shows the deployable heat shield 30 in a stowed configuration. FIG. 3B shows the deployable heat shield 30 in a partially deployed configuration, which is a snapshot of the deployable radiant heat shield 30 moving from a stowed configuration to a deployed configuration or vice versa. FIG. 3C shows the deployable heat shield 30 in a deployed configuration, showing a longitudinal axis 32 passing through the polygonal (pentagonal) central region 10 in the direction of the intended movement of the deployable spacecraft atmospheric entry heat shield 30.

[0036] The deployable radiant heat shield 30 can be described as having a deformable polyhedral surface having a plurality of sectors 2, each sector 2 having a plurality of mountain fold lines 4, a plurality of valley fold lines 6, and a plurality of facets 8 between the fold lines. As shown in the progression from FIG. 3A to 3C, the deformable polyhedral surface has only a single intended degree of freedom, and the heat shield 30 is configured to deploy from a stowed configuration to a deployed configuration with a single degree of freedom. The deformable polyhedral surface can also have a polygonal central region 10 from which the fold lines 4, 6 radiate outwardly or to an outer edge 12 of the heat shield 30, similar to the flasher pattern 20. The deformable polyhedral surface also has radial symmetry about the polygonal central region 10, just as the plurality of sectors are radially symmetric.

[0037] In the deployed configuration, the deployable heat shield 30 has an outer diameter of 1.3 m. For this diameter, the radiant heat shield 30 has a resistance of 5 kg / m 2 Ballistic coefficient of less than 1m 2 / kg mass-to-radiating area, with a peak heat flux of 300 kW / m 2 and the equilibrium temperature is reduced to ~1300K.

[0038] The deployable radiant heat shield 30 may be made of a rigid or semi-rigid material, in which case the flasher pattern of the heat shield may define minor fold lines 5 and minor fold lines 7 to allow the heat shield to be deployable from a stowed configuration to a deployed configuration. Examples of such rigid materials may be certain metal alloys, such as pure elemental metals, such as steels, titanium alloys, nickel superalloys, and refractory groups (or other materials typical of radiant heat shields), or ceramics, such as silica-based ceramics and carbon-carbon (or other materials typical of ablative heat shields). Alternatively, the deployable radiant heat shield 30 may be made of a resilient material, in which case the deployable radiant heat shield 30 may not define any minor fold lines 5, 7, and the resilient properties of the material (of each of the facets 8) may allow the deployable radiant heat shield 30 to be deployable from a stowed configuration to a deployed configuration. The material of the deployable radiant heat shield 30 can be a resilient material capable of absorbing energy when elastically deformed into a stowed configuration without permanent distortion and releasing to a steady state upon unloading. Examples of such resilient materials can be metals and metal alloys.

[0039] The stowed configuration is shown in FIG. 3A. The outer edge 12 of the deployable radiant heat shield 30, evident in FIGS. 1A, 2A, and 3B, is wrapped around the hub (i.e., polygonal (pentagonal) central region 10) when the deployable radiant heat shield 30 is in the stowed configuration. By stowage of the deployable radiant heat shield 30 in the stowed configuration, elastic energy can be stored in the fold lines and / or material of the heat shield. The deployable radiant heat shield 30 can be held in the stowed position by a retention means. The heat shield can be configured to deploy from the stowed configuration to the deployed configuration (the stable state of the deployable radiant heat shield 30) under elastic energy of at least the fold lines and / or material of the heat shield. By using a helical winding of elastic material (e.g., metal / metal alloy) in the stowed configuration of the flasher pattern, the deployable radiant heat shield 30 can be preloaded, which simplifies the deployment to the unloaded shape (i.e., deployed configuration / stable state) without the need for an active deployment system. In this manner, the deployable radiant heat shield 30 is positioned to spring outwardly once released from its deployed position.

[0040] The retention means is used to hold the deployable radiant heat shield 30 in the stowed configuration. The retention means may be any suitable high tensile string, cord, chain, binding, sheet, etc. The retention means may be released when an activation threshold, such as temperature or altitude, is exceeded. The release process may be automatic under certain conditions, or may be microprocessor controlled or pre-programmed based on the output of sensors. The retention means may be released using a thermal knife system, a pyro release system, or other suitable means. Once released, the internal strain of the rolled deployable radiant heat shield 30 causes rapid self-deployment to the deployed configuration.

[0041] Figure 4A shows the deployable heat shield 30 from a top view in a stowed configuration, Figure 5A shows the deployable heat shield 30 from a side view in a stowed configuration, and Figure 6A shows the deployable heat shield 30 in an isometric view in a stowed configuration as it would be attached to a spacecraft via the polygonal central region 10.

[0042] Figure 4B illustrates the deployable heat shield 30 from a top view in a partially deployed configuration, Figure 5B illustrates the deployable heat shield 30 from a side view in a partially deployed configuration, and Figure 6B illustrates the deployable heat shield 30 installed on a spacecraft in a deployed configuration.

[0043] Figure 4C illustrates the deployable radiant heat shield 30 from a top view in a deployed configuration, Figure 5C illustrates the deployable radiant heat shield 30 from a side view in a partially deployed configuration, and Figure 6C illustrates the deployable radiant heat shield 30 mounted on a spacecraft in a fully deployed configuration.

[0044] Advantageously, the flasher pattern can be designed to have a large difference between the stowed and deployed diameters (typically 1:5), i.e., the diameter of the heat shield in the deployed configuration is larger than the diameter of the heat shield in the stowed configuration.

[0045] For the deployable radiant heat shield 30, the ratio between the polygonal central area 10 and the outer flat diameter of the flasher shape is 1:11. This allows the pre-deployed deployable radiant heat shield 30 to be compatible with spacecraft 36, such as cube-shaped spacecraft (i.e., cubesats), nanosats, or other small or large satellites that are typically subject to limited launch volumes. As mentioned above, the ballistic coefficient of a body is a measure of its ability to overcome air resistance during flight and is equal to its mass divided by the drag coefficient per cross-sectional area. Thus, a lower ballistic coefficient can be achieved by using a deployable radiant heat shield 30 with a relatively large deployed diameter. The inventors have determined that a lower ballistic coefficient reduces the peak heating of the heat shield to less than 100 kW / m 2 It has been found that the radiation surface area can be reduced to a level where the equilibrium temperature during peak heating of a low orbital re-entry vehicle is less than the melting point of many metals / metal alloys that can withstand the aerodynamic and deployment stresses during flight. Examples of such metals / metal alloys are pure elemental metals such as steels, titanium alloys, nickel superalloys, and those of the refractory group.

[0046] As shown in Figures 6A-6C, when deployed, the deployable radiant heat shield 30 is much larger in two dimensions than the spacecraft 36 to which it is attached, and therefore extends beyond the main volume of the spacecraft. This allows the deployable radiant heat shield 30 to radiate heat both forward (in the direction of flight) and aft, which greatly increases the effective radiating area per kilogram of heat shield material. Also, the aft-facing sections of the deployable radiant heat shield 30 radiate to a cooler environment (during atmospheric re-entry) than the forward-facing sections, which further increases the heat radiation efficiency.

[0047] Advantageously, the deployed configuration of the deployable radiant heat shield 30 leaves a non-flat pattern with pronounced peaks 4 and valleys 6 protuberances that increase the total radiating surface area of ​​the deployable radiant heat shield 30 without affecting the total drag area, which allows these two parameters, i.e., total radiating surface area and drag area, to be controlled separately.

[0048] Thus, the deployable radiant heat shield 30 benefits from an increased radiating surface area to mass ratio and has the dual advantages of lowering peak heat flux (related to the ballistic coefficient) and increasing radiating area, both of which result in lower equilibrium temperatures that allow for the use of less exotic materials (e.g., ablative materials such as silica ceramics or carbon-carbon that are difficult to manufacture and machine) for use in the deployable radiant heat shield 30. The low equilibrium temperatures during peak heating of low Earth orbit (LEO) re-entry of the radiant heat shield 30 can be lower than the melting points of many metals and metal alloys.

[0049] Advantageously, the low ballistic coefficient reduces terminal velocity to a level in the atmosphere where a separate fabric parachute is not required. This is a major advantage since the deployment of fabric parachutes for atmospheric entry spacecraft typically requires a mortar system and carries significant risk of parachute deployment failure. Fabric parachutes also reduce the accuracy of landing zone estimates due to transient expansion and drag characteristics.

[0050] The deformable polyhedral surface (e.g., flasher pattern) of the deployable radiant heat shield 30 can be formed as a unitary sheet material or of materials with substantially similar thermal conductivity. This allows thermal energy to be easily distributed around the deployable radiant heat shield 30, smoothing out temperature hot spots and reducing the risk of uneven heating damaging the deployable radiant heat shield 30. For example, the mountain folds 4 naturally absorb more thermal energy upon atmospheric entry, but this is distributed through the sheet material(s) to the valley folds 6 and areas that absorb less thermal energy. Metals / metal alloys are generally suitable materials for the deployable radiant heat shield 30 because they have high thermal conductivity. The overall shape of the deployable radiant heat shield 30 can be etched as a single monolithic piece using computer-controlled metal machining tools. This single monolithic piece can then be coated with a high emissivity (>0.8) and low absorption coating (<0.1) to further reduce the peak heating temperature.

[0051] Radiation heat shields are generally more easily reused because they do not rely on surface evaporation, as opposed to ablative heat shields, which rely on surface evaporation. Advantageously, the deployable radiant heat shield 30 has a large deployed diameter state, and therefore a relatively large radiating area, as compared to typical fixed size (i.e., non-deployable) radiant heat shields. The larger radiating area allows a greater proportion of heat to be radiated from the system, and thus the balance between inflow and outflow heat occurs at a lower equilibrium temperature.

[0052] The deformable polyhedral surface (e.g., flasher pattern) of the deployable radiant heat shield 30 adopts a swept back shape in the deployed configuration, as shown in Figure 5C. The swept back shape is a result of the mrh parameters of the flasher pattern. The swept back shape forms the deployable radiant heat shield 30 into a generally frusto-pyramidal shape, as shown in Figure 5C.

[0053] This sweepback angle is used to vary the center of pressure along the longitudinal axis 32 (i.e., the flight axis), so that the center of pressure is located aft of the polygonal central region 10 and / or aft of the center of mass of the spacecraft 36. This significantly improves the aerodynamic stability of the spacecraft 36 during all stages of atmospheric entry. For example, a ratio of 1:11 between the polygonal central region 10 and the unloaded outer flat diameter of the deformable polyhedral surface gives a natural sweepback angle of substantially 30 degrees, as shown in FIG. 5C. FIG. 5C shows a sweptback shape of the deployable radiant heat shield 30 with an angle between the longitudinal axis normal and the mountain ridge of the outer edge of the heat shield of 30 degrees. This further enables passive aerodynamic stability that reduces the requirements of the attitude control system of the spacecraft 36 to maintain a safe pointing (i.e., against the direction of flow, which best protects the platform from heat) during atmospheric entry. Advantageously, the non-flat pattern of the deployable radiant heat shield 30 in its deployed configuration, having pronounced peaks 4 and valleys 6 protuberances (formed via peak and valley fold lines), also enables improved passive aerodynamic stability.

[0054] The deformable polyhedral surface (e.g., flasher pattern) of the deployable radiant heat shield 30 is configured such that the multiple fold lines 4, 5, 6, 7 of each sector substantially limit the angular rotation of the deployable radiant heat shield 30 about the longitudinal axis 32. This prevents the deployable radiant heat shield 30 (and the spacecraft 36) from undergoing rotation during atmospheric re-entry, which leads to high aerodynamic stability in the longitudinal axis and easier aerodynamic analysis. In other words, the deployed configuration shape of the deployable radiant heat shield 30 is approximately axisymmetric. This provides a flight stabilization function and also resists collapsing back to the stowed position because the ridges of the peaks 4 and valleys 6 distribute air pressure to the opposite sides of the deployable radiant heat shield 30. This contributes to the deployable radiant heat shield 30 being able to retain the deployed configuration shape without additional retainers.

[0055] As mentioned above, the materials of the deployable radiant heat shield 30 can include metals and metal alloys (here metals). These have many advantages, for example, metals can be much easier to process and form into customized shapes than brittle silica tiles. Metals are not brittle, allowing for thermal expansion without cracking. Plastic deformation of metals can also be used to define the desired deployed configuration (i.e. unloaded shape) without additional retainers or additional secondary fold lines 5, 7. Metals can be coated with a wider range of materials to improve corrosion and radiation properties. Metals are much less expensive to procure and can be machined using more ubiquitous tools, leading to a significantly reduced cost per square meter. Metals have predictable and very linear stress-strain characteristics with temperature, which aids in strength analysis.

[0056] The deployable radiant heat shield 30 may be mounted to the spacecraft 36 using a polygonal central region 10 on a custom shield mount connected to a forward-facing panel of the spacecraft 36. Other mounting methods are possible and will be apparent to those skilled in the art.

[0057] Turning now to alternative heat shields, the heat shield can be an ablative heat shield with many of the features, configurations, and advantages of the deployable radiant heat shield 30, as will be apparent to those skilled in the art. Ablation requires highly specialized heat shield materials with low vaporization thresholds and high heat capacity. Ablative dominated heat shields are typically not reusable due to the significant erosion of material that occurs upon atmospheric re-entry. The ablative heat shield can be made from a rigid material, in which case the flasher pattern of the heat shield can define minor fold lines and minor fold lines to enable the heat shield to be deployable from a stowed configuration to a deployed configuration. Examples of such rigid materials can be ceramics, or other materials typical of ablative heat shields. Retracting the heat shield into a stowed position can store elastic energy in the fold line material, which can be different from the ablative material. The ablative heat shield can also be held in the stowed position by a retention means. The ablative heat shield can be configured to deploy from a stowed configuration to a deployed configuration under elastic energy at least at the fold lines of the heat shield, although additional retainers may be used.

[0058] The fold lines of the deployable heat shield (radiation or ablation) can be achieved by proxy folds, hinges, or any other suitable folding means, for example, two layer panels with a thin flexible material between them.

[0059] The heat shields disclosed herein may additionally have retainers that may be used to support the deployed configuration of the deployable heat shield (radiation or aerial) during atmospheric re-entry. The retainers may be reinforcing ribs that extend along the fold lines.

[0060] The deployable heat shields disclosed herein may alternatively have a deployed configuration with a swept back angle between a normal to the longitudinal axis and a crest of the outer edge of the deployable heat shield of substantially between 0-45 degrees, 20-40 degrees, 22-38 degrees, 25-35 degrees, 27-32 degrees, or 30 degrees. Alternatively, the swept back angle may be substantially 30 degrees.

[0061] Alternatively, the diameter ratio of the deployable heat shield to the spacecraft may be any ratio, in particular greater than 2, such as when there is a 1:11 ratio between the polygonal central area and the outer flat diameter of the deformable polyhedral surface.

[0062] Although the deformable polyhedral surfaces of the deployable heat shields disclosed herein are shown to be the same as mathematically defined (thickness-accommodated) flasher patterns, the heat shields disclosed herein may vary from the strict mathematical definition, i.e., they are substantially mathematically defined, while still performing substantially the same functions. For example, the deployable heat shields may not have a polygonal central region from which fold lines radiate outward. The deployable heat shields may not have radial symmetry about a polygonal central region such that multiple sectors are radially symmetric. For example, there may be intentional modifications that result in the deployable heat shields not having radial symmetry.

[0063] The deployable heat shield technology disclosed herein is scalable from pico (<1 kg) to large (>1000 kg) spacecraft and can be adapted by further tailoring the area-to-mass ratio for higher orbits and re-entry from other planetary bodies, including the Moon and Mars.

[0064] The deployable heat shields disclosed herein can also be used as an additional barrier while in space, thus acting as a layer of protection against these threats. Unlike other space deployable structures, the deployable heat shields disclosed herein can be made from elastic materials that can potentially protect the spacecraft from incoming particles, especially in low Earth orbit / space. For example, the deployable heat shields disclosed herein can be used to protect the spacecraft from tail debris when used in comet exploration missions.

[0065] The deployable heat shields disclosed herein can also be used as radiation shields.

[0066] The deployable heat shield disclosed herein can also be used as a wake shield to reduce ambient pressure and atomic oxygen ingress to improve vacuum in low Earth orbit. Simulation results also show that a deployable heat shield suitable for surviving re-entry would reduce the ambient pressure behind the shield by two orders of magnitude when at altitudes between 400 and 800 km. This is important for on-orbit manufacturing processes and scientific research that require ultra-high vacuum levels.

[0067] The deployable heat shields disclosed herein can also be used as drag sails to passively reduce altitude without the use of propellants. A deployable heat shield suitable for surviving re-entry would provide a passive reduction in altitude of a LEO satellite of ~10km / month (depending on solar activity). This is useful for re-entry missions as it saves propellant needed to move into a low orbit in preparation for re-entry (low starting point leads to high accuracy of the landing zone).

[0068] The deployable heat shields disclosed herein can also be used as flotation devices, with the large relative displacement and high center of volume of the retreating shield ensuring static stability in light to moderate seas, allowing the heat shield to replace the need for separate flotation device deployment systems that require additional resources and involve deployment risks.

[0069] The deployable heat shield features disclosed herein can be created based on any mathematically definable thickness-enabled flasher pattern / mathematically definable deformable polyhedral surface. The parameters m=5, r=2, h=2 used in Figures 1-6C were selected for illustrative purposes, and it is understood that any type of flasher pattern / deformable polyhedral surface that can be substantially defined by mrh parameters can be used to create the deployable heat shields disclosed herein. Examples of other flasher patterns with alternative mrh parameters are shown with reference to Figures 7A-9B.

[0070] Figures 7A and 7B show alternative thickness-dependent flasher patterns 40 having parameters of m = 5, r = 1, and h = 2. Figure 7A shows thickness-dependent flasher pattern 40 with minor fold lines shown, and Figure 7B shows thickness-dependent flasher pattern 40 without minor fold lines shown.

[0071] Figures 8A and 8B show alternative thickness-dependent flasher patterns 50 having parameters of m = 4, r = 3, and h = 1. Figure 8A shows thickness-dependent flasher pattern 50 with minor fold lines shown, and Figure 8B shows thickness-dependent flasher pattern 50 without minor fold lines shown.

[0072] Figures 9A and 9B show an alternative thickness-dependent flasher pattern 60 having parameters of m = 10, r = 2, and h = 5. Figure 9A shows thickness-dependent flasher pattern 60 with minor fold lines shown, and Figure 9B shows thickness-dependent flasher pattern 60 without minor fold lines shown.

[0073] 10A and 10B show an alternative thickness-responsive flasher pattern 70 in which the minor fold lines are omitted and defined by parameters m=6, r=3, and h=2. FIG. 10A shows the flasher pattern 70 with a superimposed dashed circle 80. FIG. 10B shows a thickness-responsive flasher pattern 90, which is the thickness-responsive flasher pattern 70 cut into the dashed circle 80. That is, the thickness-responsive flasher pattern 90 is a "circularized" thickness-responsive flasher pattern. The outer edge 12 of the thickness-responsive flasher pattern 90 corresponds to the dashed circle 80. Thus, the outer edge 12 of the thickness-responsive flasher pattern 90 corresponds to a circle (i.e., the outer edge 12 is circular). Specifically, if the deformable polyhedral surface were forced into a single plane, i.e., flat, the outer edge 12 would be circular. This is preferably done at the design stage. The thickness-responsive flasher pattern 90 is dimensioned based on its radius such that the outer flat diameter can be defined by the radius of the dashed circle 80. The thickness-varying flasher pattern 90 may provide improved aerodynamic stability over the thickness-varying flasher pattern 70. Any thickness-varying flasher pattern may be modified to provide an outer edge 12 that corresponds to a substantial circle and may benefit from improved aerodynamic stability.

[0074] Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications can be made without departing from the scope of the invention as defined in the appended claims.

Claims

1. 1. A deployable spacecraft atmospheric entry heat shield having a deformable polyhedral surface comprising: mth order polygon center area; and a plurality of sectors disposed about the polygonal central region, each sector having a plurality of mountain fold lines, a plurality of valley fold lines, a plurality of facets between the fold lines, and an outer edge; the plurality of mountain fold lines includes a plurality of major mountain fold lines and a plurality of minor mountain fold lines; The plurality of valley fold lines include a plurality of major valley fold lines and minor valley fold lines; the major mountain fold lines and the major valley fold lines extend from the polygonal central region to the outer edges; the heat shield is configured to deploy from a stowed configuration to a deployed configuration; the deformable polyhedral surface is configured to adopt a retracted shape in the deployed configuration; the major mountain fold lines have a sweepback angle that is substantially constant from the polygonal center region to the outer edges in the unfolded configuration; Heat shield.

2. The heat shield is a radiant heat shield. The heat shield of claim 1 .

3. The secondary mountain fold line and the secondary valley fold line extend between the main mountain fold line and the main valley fold line, The heat shield of claim 1 .

4. the deformable polyhedral surface has m-th order radial symmetry about the polygonal central region, and the sectors are radially symmetric; The heat shield of claim 1 .

5. the heat shield surface being formed as a unitary sheet of material; The heat shield of claim 1 .

6. the plurality of fold lines in each sector are configured to substantially limit angular rotation of the heat shield about a longitudinal axis in use. The heat shield of claim 1 .

7. a ratio between a diameter of the polygonal central region of the deformable polyhedral surface and an outer flat diameter of the unfolded configuration is 1:11; The heat shield of claim 1 .

8. The deployed configuration is configured such that the center of mass of the heat shield is behind the polygonal central region relative to a normal to a longitudinal axis of the heat shield. The heat shield of claim 1 .

9. the deployed configuration being stable, and a retention means configured to maintain the heat shield in the stowed configuration. The heat shield of claim 1 .

10. The retaining means is configured to release when an actuation threshold is exceeded. The heat shield of claim 9.

11. A spacecraft comprising a heat shield according to any one of claims 1 to 10.

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