Basic Principles and Mechanical Considerations of Satellites: A Short Review

The skin-frame structural concept mounts outside skin panels with fasteners or rivets, utilising an interior skeletal network of axial and lateral frames. Bending, torsional and axial forces are supported by these frames, while the skin strengthens the structure by sustaining the shear stresses induced by the connections between the inner elements. Despite the thin skin’s tendency to cause some structural instabilities, the skin’s thickness is occasionally decreased to minimise the total mass. All extra shear loading is transferred to in-plane tension forces at 45° when the skin buckles by shear, and the connections are then required to support these stresses. Excessive deformations occur in the skin through buckling modes, rendering it inappropriate for mounting outside components like solar cells. Adding intermediate elements usually improves the assembly’s buckling strength.

For stability, truss members are often fabricated individually and organised in triangle arrays. Extruded tubes composed of metallic or composite materials are used to fabricate the truss members. A stable truss contains no extra members that could generate additional load paths and is clearly recognisable. Trusses are often mass-efficient when the members are arranged into cross-sectional assemblies with a rectangular or triangular cross-section. However, if the cross-section is more circular or hexagonal, they become less effective. Inherent stress concentrations are also produced by the structure’s design at interface mounting locations, such as separation systems. Because there are no shear panels, components can be installed both internally and externally, and a payload is easily accessible. However, satellites that use body-mounted solar cells will not benefit from this absence of shear panels.

Monocoque cylinders are composed of axially symmetric shells without frames or stiffeners. Sandwich or metallic panels with rolled, curved sections are used to make the shells. Usually, the cylindrical form is created by fabricating and joining two or three curved parts. Monocoque cylinders’ buckling strength typically sets a limit on their strength. When the loads are evenly spread throughout the structure, the shells perform more efficiently. Fasteners are generally used to mount components to the walls, but it is important to take care not to overload the shell and lead to local failures. A satellite with body-mounted solar cells and reasonably light components can use the monocoque cylinder design.

Axial and lateral frame components attached to an outer skin are used in the construction of cylindrical skin-stringer systems. These designs resemble skin-frame structures, although this category of structures only includes circular cylinder arrangements. Despite the thin skin’s tendency to cause some structural instabilities, the skin is occasionally thinned to reduce mass. Through the diagonal tension phenomenon, the post-buckling behaviour of the skin converts the additional applied shear stresses to torsion. To allow the assembly to operate as a uniform surface, the skin and components must attach consistently. Commonly used connecting techniques include rivets and/or fasteners. Along the stringer assembly, interior parts are often fastened. When applying local loads, this approach is more effective than mounting monocoque cylinder parts. The skin must be sufficiently rigid to allow for the secure installation of exterior components such as solar cells on the body.

Isogrid improves a structure’s stiffness per weight by utilising an array of equilateral triangular cutouts. The geometric parameters of the isogrid structure are presented in Figure 5 [28]. A metallic panel may be machined to create the isogrid configuration, or fibre composite materials may be used to build it. Since the idea was first developed when NASA released a report titled ‘Isogrid design handbook’ [29], isogrid structures have been considered. This handbook examines the isogrid structure manufacturing process and mechanical testing. When these structures are subjected to an axial compressive load, the failure may be caused by global buckling if the panel is slenderer than one of the ribs or local buckling if the ribs are slenderer than the panel [29]. The development is still ongoing nowadays, with the majority of research going into composite applications.

Figure 5.

Since composite materials have high specific characteristics, their employment in the aircraft sector is a well-established practice [30]. The structures may be subjected to significant axial compressive loads in some applications, such as propellant tanks for rockets, fuselages and boosters, which could result in the composite panels failing due to buckling processes [31]. Composite stiffening ribs can be utilised to add stiffness and strength to prevent this failure. These stiffeners are often created as isogrid structures or equilateral triangles and attached to the laminates [32].

A few research studies have focused on the design optimisation of isogrid structures to enhance their particular strength and stiffness. Finite element simulation was used by Zheng et al. [30] to optimise a stiffened cylinder with a height of 6 m and a diameter of 4 m. They stated that due to the quadratic relationship between critical force and thickness, the rib thickness should be greater than the rib width.

Finite element simulation was also used by Dawood et al. [33] to investigate how the isogrid technique changed the structure’s natural frequencies. Due to the lower mass of the isogrid structure, the analysis results showed that the natural frequencies estimated for the isogrid case were greater than those for the non-isogrid case.

By using additive manufacturing, Forcellese et al. [34] created isogrid panels in short carbon fibre-reinforced polyamide with varying rib thickness and rib height values. Through compression experiments performed at room temperature, they examined how geometric parameters such as rib thickness and rib height affected the isogrid panels’ compressive strength and buckling behaviour. They found that increasing rib width causes a rise in the peak load and specific maximum load, while increasing rib height causes a decrease in the maximum load and specific maximum load. They also observed that the global buckling failure mode caused the isogrid panel to fail during compression testing, indicating that the structure’s slenderness is higher than that of one of the ribs.

The buckling and environmental performances of composite structures produced by 3D printing were also studied by Forcellese et al. [35]. In their comparison, solid, dried solid, isogrid with 3 mm rib width and dried isogrid with 5 mm rib width were all employed. They found that the lowest environmental loads and lowest buckling resistance are exhibited by isogrid with 3 mm rib width. The results for the solid and the isogrid with dried panels that have 5 mm rib width in terms of buckling resistance are quite close. The latter, however, has worse environmental burdens than the former.

The effects of rib width and rib thickness on the strength and specific strength of polyamide reinforced with 20% of the weight of short carbon fibre were examined by Ciccarelli et al. [28]. They found that an increase in rib width causes a drop in specific strength and an increase in overall strength, while an increase in rib thickness causes an increase in the overall strength and specific strength, though this impact is less pronounced when specific strength is considered. They also noticed that depending on the geometric parameters, isogrid structures can fail under local or global buckling; the configurations that result in the highest strength are different from those that result in the highest specific strength, and isogrid structures that experience a local buckling failure mode exhibit lower specific strength than anticipated.

A sandwich structure is made up of two thin face sheets connected to both sides of a lightweight honeycomb core. Sandwich structures are designed so that the core can withstand normal flexural shears, while the outer face sheets can support axial loads, bending moments and in-plane shears [27]. Due to the heterogeneous nature of the core/face sheet assembly, sandwich structures are prone to failure due to high local stress concentrations. The point loads from connections must consequently be distributed during component mounting using potted inserts. Aluminium or composite panels made of graphite and epoxy are frequently used to manufacture sandwich panel face sheets. Typically, an aluminium honeycomb structure is used to manufacture the core. The lightest solution for situations involving compressive or bending loading is a honeycomb sandwich panel (Fig. 6) [36]. The geometric parameters of the honeycomb sandwich structure are shown in Figure 7 [37]. The honeycomb geometry is anisotropic, and the longitudinal direction exhibits increased stiffness [38]. However, when put together in a sandwich arrangement, the core behaves almost isotopically for in-plane loads. The mounting requirements for potted inserts and the thermal inefficiency are drawbacks of using honeycomb cores. These inefficiencies are due to the limited thermal conductivity of the adhesive layers utilised in fabrication, which makes the honeycomb structure unsuitable for optical and mirror aircraft applications.

Figure 6.

Figure 7.

Thin strips are shaped into honeycomb cells to create honeycomb sandwich cores. It is quite challenging to mill honeycomb core due to several of the required characteristics that make it appropriate for a wide range of applications. The intricate procedures required by the honeycomb composites’ unique shape make them different from other CNC machining processes. In addition, honeycomb cutting frequently necessitates substantial postprocessing to eliminate partially released flags along the machined edge of the core walls. Specialised cutting tools that are more expensive than regular CNC machine parts are frequently needed for honeycomb machining. Additionally, the complex honeycomb machining process with fewer errors can be achieved with a skilled machinist partner [39]. Applying the right and properly maintained tools is a crucial aspect of successful honeycomb machining. To create the intricate hexagonal structure required in the honeycomb core panels, all cutting tools for honeycomb CNC machining must be manufactured using hard materials and kept at a high level of sharpness. For honeycomb machining, polycrystalline diamond (PCD)-tipped milling cutters, PCD-tipped turning tools, flute cutters and 5-axis CNC machines are the best options.

The crushing strength of the kraft paper honeycomb with varying cell wall thickness, cell size and density values under compression loading was numerically examined by Abd Kadir et al. [40]. They found that crushing strength increases as cell size decreases, with 10-mm cells showing the highest strength. Meanwhile, the honeycomb has better crushing strength, thanks to its 0.4-mm-thick cell walls. Additionally, denser materials have greater crushing strength.

The effect of honeycomb design parameters on the forced vibration behaviour of aircraft sandwiches with honeycomb cores was investigated experimentally and numerically by Sadiq et al. [37]. To examine how honeycomb structure parameters affect the transient response of sandwich structures, core height, cell size and cell wall thickness were considered. They found that raising the cell wall thickness from 0.8 mm to 1.5 mm leads to an increase in the maximum transient response, while increasing the core height from 5 mm to 25 mm leads to a decrease in the maximum transient response. Additionally, they found that increasing the cell size had a detrimental impact on the response’s maximum transient. According to their findings, variation in the core height has a larger influence than the other factors, although the cell wall thickness has a lesser influence.

Aluminium honeycombs’ mechanical responses to in-plane and out-of-plane compression were investigated by Li et al. [41]. They noticed that the walls of aluminium honeycombs were thickening. Some significant observations about the quasi-static and dynamic crushing responses of honeycomb structures were established by Thomas and Tiwari [42]. The face-sheet material and thickness have a significant impact on the core’s ability to absorb energy during dynamic projectile and impulsive impacts. The performance of the honeycomb core under static stresses applied in various directions was influenced by geometrical factors such as cell size, node length, cell wall thickness and cell configuration. Hexagon, triangle, square and circular honeycomb cell configurations were chosen based on minimal material requirements and maximum strength standards. The ideal configuration for maximum strength and the lightest material was hexagonal honeycomb.

Static loads, which include static external and static self-contained loads, can be produced from a variety of sources. The external loads that are quasi-static loads are due to the structural modules’ and the supported components’ inertia from steady acceleration or gravity. A steady-state acceleration remains constant as time changes, and that comes from gravity when handling the satellite. The mechanical preloads known as static self-contained loads include thermo-elastic loads produced by temperature variations and purposeful internal loads created during assembly [43]. Static load tests are intended to simulate the launcher, motor and spin-induced static loads imposed by static and dynamic accelerations. Due to launcher acceleration and aerodynamics, satellites are subjected to simultaneous static and dynamic loads throughout the launch process. A satellite must be designed to endure these loads using FEA. For a better understanding of the structural dynamics of the component, more transient analysis methods have been required but at the expense of runtime and large data files. So, a typical method in design is to combine static and dynamic loads into an equivalent static load known as a quasi-static load to validate the satellite’s structural design. Whiffle-tree tests, centrifuge tests and acceleration with vibration test systems are the three methods of static testing that are employed [44]. One of those tests involves applying quasi-static loads to the component, utilising vibration test devices to simulate the loads the launcher adaptor exerts on the structure.

Different sources such as dynamic external loads, dynamic self-contained loads and thermo-mechanical loads can generate dynamic loads on a satellite structure. Engine power, sound pressure and wind gusts are some examples of dynamic external loads that might occur during launch, as well as time-varying forces brought on by vibration while being transported to the launch site. Dynamic self-contained loads can be imposed by mass loading a vibrating satellite while it is being environmentally tested or while it is in space after the force that created the excitation has been withdrawn. On-orbit loads that result from thermal cycles as a result of the on-orbit environment’s experience are considered of the thermo-mechanical load type. Other dynamic loads are brought about by satellite manoeuvres during the orbital or attitude correction stages [43].

A single degree of freedom (SDOF) system’s reaction to a shock or any other transient acceleration is known as the shock response spectrum (SRS). A graphical representation is used to represent it. Shock loading also happens as a result of the launch vehicle’s acceleration, but to put it into perspective, imagine yourself sitting in a bus that suddenly accelerates, pushing you backwards. You experience only a very small portion of what occurs on the launch vehicle while you are on the bus. The launch vehicle experiences vibration and increased acceleration. The acceleration of the satellite is affected by vibration frequency. The same things that need to be analysed for static analysis also need to be analysed for shock analysis; it must make sure that the maximum stress exerted on the parts is not greater than the material’s yield strength considering a safety factor [47].

The satellite will be thermally stressed during the thermal tests to (1) reveal latent faults by exhibiting three cycles between hot and cold temperatures without any failures, (2) ensure that the satellite can operate successfully over a wide temperature range and during temperature changes and (3) collect temperature data about the steady-state temperature to evaluate part, board and assembly gradients.

Both the ground station (or electrical ground support equipment, EGSE) and the dynamic simulator are connected to the satellite that is within the thermal chamber during the thermal test. Electronic stimuli required for functional testing are provided by the ground station.

A thermal balance test is performed first. The thermal chamber should be capable of regulating temperatures between 40°C and +55°C at a rate of at least 3°C/min. During the test, chamber temperature gradients should not increase over 4°C/s. Temperatures at chamber and satellite test points should be monitored and recorded by a data acquisition system. The ground station should gather data from the satellite telemetry. When the chamber is within 2°C of the required setpoint, environmental stability is attained.

The satellite is normally qualified for the launch environment using vibration tests. These tests generally consist of sine, sine-burst and random test cycles. The choice of the facility where the test will be conducted must be made beforehand. The capabilities of the vibration table and the number of instrumentation channels available are the two main factors to be taken into account when selecting a test facility. To fasten the satellite to the vibration table, an interface fitting must be made. The interface fitting should have a very high resonance frequency (> 1,000 Hz). The bolt pattern for bolting to the table would be provided by the testing facility. Theoretically, a test may be carried out with no instrumentation at all; as long as the satellite survives the test environment without malfunctioning, the test is a success. Even though there would be a big risk, it would be acceptable, especially during random vibration testing.

The satellite is normally not powered during tests, and it is in the launch configuration without thermal blankets. The satellite is covered by a lightweight bag that enables accelerometer attachment. Lifting equipment that is not necessary for flight configuration is taken out. The stowed configuration of solar panels is installed and locked.

When the vibration test is performed on multiple axes, there will be one that is most crucial. It is possible to learn more about the structure before applying the critical test loads by beginning with the non-crucial axis. For each axis, the first and last test cycles are signature cycles, either sine (5–2,000 Hz) or random (20–2,000 Hz). In each axis, the sine, sine-burst and finally random vibrations are tested. It is simple to compare the sine and sine-burst tests to the predictions from the pretest FEM study. This enables the engineer to confirm that the instrumentation is operating properly. All satellite modes across the whole frequency range are stimulated during the random test. The test results give the first indication of the model’s accuracy. Certain modes might couple, resulting in a higher response than expected, while the results of the sine and sine-burst tests are often what is expected.

The launch vehicle acts as a flexible constraint when the satellite is attached to it during launch. When one of the satellite modes takes off, the interface loads with the launch vehicle rise, although some of these loads may be reduced by the flexibility of the satellite. Depending on measured loads (from prior launches) or coupled load analysis, static design loads are determined. These account for this dampening. Therefore, from a strength perspective, the satellite just has to be able to withstand 1.25 times the design loads. The sine-burst test is considered a strength test, while random testing is considered a workmanship test, not a structural or strength test. This is important to take into account when performing the random vibration test. On the other side, the satellite is fastened to a vibration table that offers no damping during vibration testing. Therefore, the response may be substantially greater than what would be observed during flight if one of the satellite modes takes off during testing. This phenomenon allows for the limitation of input levels during random testing.

This optional bakeout test’s goal is to outgas satellite components to reduce the possibility of condensable volatile contamination upon their flight. Before bakeout, the satellite undergoes thermal cycling at ambient pressure to reduce the possibility of condensable volatile contamination upon its flight. Before baking, the satellite undergoes thermal cycling at ambient pressure. Integration of the satellite into the thermal vacuum chamber (TVAC) requires confirmation of the chamber’s cleanliness. The thermal vacuum bakeout chamber is where the spacecraft and its TVAC ground support equipment (GSE) are installed. During the test, the chamber is kept at +40°C and less than 1 × 10⁻⁵ torr of pressure. The data acquisition system continuously monitors satellite temperatures, compares them to alerts and records temperatures in real time until the satellite has reached bakeout.

Satellites or their components are commonly tested in a TVAC in a space environment simulation. The TVAC has a controlled radiative thermal environment. Typically, the thermal environment is created by applying thermal lamps for high temperatures or by flowing liquids or fluids through thermal shrouds for cold temperatures. Temperatures are recorded real time by the data acquisition system, which continuously checks and compares them with satellite temperature alarms. The purpose of TVAC tests is to validate the thermal design by exposing the satellite and payload to thermal test environments that conservatively simulate the hot and cold conditions of flight, gathering steady-state and transient data to correlate the thermal models, and operating the satellite at temperatures above those anticipated in the orbit.