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Spacecraft Mechanical Sub-system Design & Manufacturing

Mechanical engineering is an engineering branch that combines engineering physics and mathematics principles with materials science, to design, analyze, manufacture, and maintain mechanical systems It is one of the oldest and broadest of the engineering branches.


“Mechanical Engineering takes in all built structures and moving parts flown in space, which includes automation and robotics, instruments for scientific missions as well as assessing the effects of the space environment on materials,” According to the European Space Agency


The mechanical subsystem is responsible for designing and manufacturing the satellite structure while ensuring that the satellite can bear all the structural and thermal loads.
This makes three major subdivisions in the mechanical subsystem:

  • Structures: Ensures the structural stability of the satellite while launch and during the operations of the satellite.
  • Thermals: Ensures that the satellite functions in varied thermal cycles presented by the harsh space conditions in the orbit.
  • Integration: Manufacturing and the integration of the various components of the satellite to ensure proper functioning of the satellite within the mass budget.


Keeping in mind the end goal of the fully functional satellite for the whole period of mission life, the mechanical subsystem contributes by making sure that the journey to the orbit and the harsh environment of space doesn’t pose a problem to the other subsystems. The Structures division provides the basic structure of the satellite known as the bus of the satellite. Also, the various mechanisms such as the deployable solar panels are designed by the structures team.


The Thermals team makes sure that all the components of the satellite are within their operational temperature throughout the mission life. Finally, the Integration team gets to do the coolest part of actually ‘building’ the satellite. They put together the whole satellite ensuring that there are no components on board which interfere with each other.


Requirements on Structures

Launch Vehicle Placement Requirements

The satellite is to be launched on a launch vehicle. Each launch vehicle has some specifications for different classes of satellites. The satellite structure should be able to interface with the launch vehicle and hence meet the specifications of the launch vehicle.

Launch Loading Requirements

The satellite is carried to its orbit by a launch vehicle in a flight lasting about 17 minutes for Low Earth Orbits. During this period, the vehicle experiences high levels of acceleration, vibrations and shocks which are transmitted to the payloads attached to the flight decks of the vehicle. Launch loads experienced include static loads, vibration loads, acoustic loads and shocks and impose certain strict requirements on the structure of the satellite. Satellite structure should be able to withstand these loads during launch. All the components should be safe and working after the launch. The loading specification for which the launch vehicle interface is tested is assumed to be the loading data for the satellite during launch.

Deployment Requirements

Many satellites have various deployable components such as antennas. The structures team has to ensure that the components are deployed when required.

Transportation and Handling Requirements

The satellite, once integrated, needs to be handled and transported to the launch site. For this purpose the structures team needs to design a handle for the satellite and a satellite box for transportation. Many factors such as electrostatic charge, vibration, contamination from humidity, pressure management, thermal control needs to be looked at.
Apart from this transportation boxes for individual components to be sent for integration at the clean room may be required.



The atmosphere above the height of 500 km above Earth comprises of too thin to moderate ambient thermal conditions and consequently, the environment above such a height is prone to extremes of temperature, being directly exposed to solar radiation and deep space. So, the Thermals subsystem needs to ensure all the electrical components of the satellite are well within their operating range.
Various different methods are employed to ensure the thermal stability. The major two divisions being Active and Passive thermal control. By employing these methods iteratively and analysing the results for different initial conditions, the Thermals team tries to come up with the best possible solution, with minimum increase in mass or power requirements.

Requirements on Thermals

  • Ensure that the temperature within the satellite is within the operating range of different electrical components.
  • Remove excessive heat from heat producing components.


The team works for the goal of launch of the satellite. The final Flight Model (FM) is the one which is launched into the space and undergoes the mission of the satellite. Integration team’s work comes into the picture once all the subsystems are ready with their designs and the individual components are manufactured. Then the critical part of integrating the components comes up. Integration team makes an Integration Sequence, making sure than integration can be done in a fast and effective manner.

Apart from the Flight Model (FM), the team also integrates a Qualification Model (QM) which is used for qualifying the satellite design, before the flight model is made. It is generally a replica of the flight model, which is passed through more rigorous tests than the flight model. It is generally mandated by space agencies to ensure that the satellite design is fit to go into their launch vehicle, along with giving a confidence to the team. The Flight Model is also tested in similar conditions as Qualification Model but the loads applied in qualified model are higher.

Importance of Integration

At first glance, Integration looks like a simple task, which may not require a separate subsystem team working on it. However, a faulty integration can blow up the entire mission. Not drafting a proper integration sequence and not testing it rigorously will often cause last minute panics if you are lucky, and indefinite delays in the mission if you are not. Also, as much as 10% of the mass can come from the various integration related components such as connectors, screws etc. So neglecting integration when preparing the mass budget is a bad idea. Also mass being a critical factor, the wire routing is to be done to ensure that minimum mass of wires is added, all the while ensuring no tangling between them. Also mass being a critical factor, the wire routing is to be done to ensure that minimum mass of wires is added, all the while ensuring no tangling between them.

Requirements on Integration

  • Complete all connections between electrical packages and route wires between them.
  • Assemble the satellite (both QM and FM).

Spacecraft Structure Features

Primary Features

✔ Accommodate payload, instruments, and compulsory on-board avionics
✔ Provide a mechanical interface between spacecraft & launcher
✔ Provide a sufficient stiffness to meet a natural frequency requirement of launcher and to minimize a dynamic amplification between launcher & satellite by carefully performing a modal design
✔ Provide a capability of S/C to withstand a launch environment (vibrations, high G-force, and shock event)
✔ Provide a sufficient mechanical stability to satisfy a pointing sensitive payload requirements i.e. build tolerance, alignment, micro-vibration, and thermo-elastic stability)


Secondary Features

✔ Protect payloads & avionics from harsh environment in outer-space (cosmic radiation & temperature
fluctuation in orbit)
✔ Protect payloads & avionics component from depressurization during launch and vacuum
environment in space
✔ Provide a proper electrical grounding design for electronic and power equipment
✔ Provide mechanical ground support equipment (MGSE) interface and harness routing accessibility to
accommodate AIT needs
✔ Provide an optimum assembly pattern to satisfy AIT requirement


Spacecraft Structure Classification

Primary Structure

– Carries a majority of mechanical loading during launch
– Know as “load path structure” that transmits loads from launcher interface to the rest of S/C
– Any failure of this structure can cause the S/C to collapse which means a major loss of mission functionality
– i.e. separation ring, centre tube, thrust tube, sep


Secondary Structure:

– Carry a significant mechanical loading
– This structure accommodate most of equipment inside/outside S/C
– Any failure of this structure can cause a partial loss of S/C mission functionality
– i.e. closure panels, deployable panel & hinges

Tertiary Structure / Non-Structural Mass:

– Any component which does not carry a significant load within the structure or contribute to structural behavior / performance
– i.e. small equipment, avionic modules, brackets


Requirements imposed by Mechanical Subsystem on other Subsystems

Volume and Mass requirements

The satellite’s mass and volume are constrained by the maximum payload the launch vehicle can carry. To ensure that these constraints set for the satellite are met, mechanical subsystem poses mass and volume requirements on all the subsystems.

Stiffness Requirements

To ensure the launch loading requirements listed earlier, mechanical subsystem poses stiffness requirements on the other subsystems. The other subsystems need to ensure that PCBs and other components comply with these requirements so that they have natural frequency well above natural frequency of the launch vehicle.

The heart of satellite mission is “a payload”, but the payload can’t work alone without support from “various avionic sub-systems”
⮚ We design “a spacecraft (S/C) structure” to accommodate both “payloads” and “avionics” by assemble them together into one piece
⮚ We also design “the S/C structure” to make it’s compatible with “a selected launcher”
⮚ Payload & launcher requirements mainly “drives” the S/C structural design
⮚ To verify a robustness & integrity of S/C structural design, structural analysis & test shall be performed


Satellite Shape & Dimension

Things to be considered…
All avionic and payload shall be fit in this structure
– Size of main payload (i.e. imager)
– Amount of avionic system required
– Power requirement
High power consumption 🡪 need more solar cell area
It must be compatible with launcher
– Will it fit inside a launcher fairing?
– Will it fit inside a dynamic envelope?
– Will it fit with available separation system?


Structure Elements

Every S/C structures, whatever how much of complexity is, they usually comprise of a following structure elements
Monocoque structure: thinwalled tubes (round or rectangular cross-section) made of composite skins with stiffener or sandwich panels
Beam, Bar, and Strut: Long component that supported axial and lateral loadings
Trusses: a set of beams or struts connected together at angles which is stronger that a single beam
Panels: thin walls designed to enclose volumes or to provide a mounting surface for spacecraft equipment. Honeycomb sandwich panel is the most popular usage


Mechanical Design & Manufacturing Tools

In order to properly design and construct a CubeSat, the analysis must be performed on CubeSat models. Examples of such “virtual tests” can include a manufacturability test, stress analysis test, and dynamic response analysis test, among others. Performing such studies on the models helps to optimize parts for improved performance in the intended environment and provides a low-cost solution to testing, in which the computer-based model is tested rather than machining the actual CubeSat and testing it multiple times, essentially eliminating multiple field tests. Furthermore, parts can be optimized for mass by performing stress analysis tests on the models to determine the minimum mass needed to have adequate structural strength.

The aim of offshore structural engineering is to produce structures that are safe, functional, economical, and able to resist the forces included by man and the environment or required period of time. Finite Element Analysis (FEA) is an engineering method of calculating stresses and strains in all materials.

The greatest stress occurs during launch hence the force likely to be experienced by the CubeSat during launch is to
be modelled and analysed using Solidworks simulation.

Strain and Deformation Analysis

The next area of concern was the deformation that occurs during launch from random vibrations and static loads. If
the loads are too great, the structure could deform and cause massive damage to the internal components. SolidWorks was able to produce values for the worst-case scenario.

CAD: Solidworks, CATIA, NX
CAD file control: Solidworks PDM, Enovia
CAM*: NC Solution, Cura


Material Selection

Materials for spacecraft structure are selected based primarily on the specific strength (strength/density) and the specific rigidity (elastic modulus/density). Other properties for consideration are ductility, fracture toughness, thermal conductivity, thermal expansion, corrosion resistance, volatility, fabrication ease, and procurement ease. The use of large amounts of magnetic materials is often undesirable from the attitude control stability consideration and interference with the environment during space physics measurements. Aluminum alloys are widely used in any part of the structure, but graphite-epoxy composite materials are also increasingly utilized for both the primary and the secondary structures to take advantage of the superior mechanical properties.

Among the aluminum alloys, A 7075 and A 2024 find wide application areas. Honeycomb sandwich panels and shells are composed of face sheets of A 7075-T6 or A 2024-T3, and honeycomb core of A 5052 or A 2024. Composite materials are also used for the core, but the aluminum core is selected when the higher thermal conductivity between the face sheets is needed. Machined elements, like ring frames, flanges, fittings, and brackets are made from A 7075-T7351 and A 7075-T7352. A 6061-T6 can be used for elements which do not require high strength. Stainless steels (A286CRES, 302CRES, 305CRES) and titanium alloys (Ti-6Al-4 V) are nonmagnetic materials and they are used for small mechanical elements and bolts. Magnesium is superior with its low-density and good vibration damping property, but special care must be taken against corrosion. Beryllium has very high specific rigidity and good thermal properties, but its use is limited or sometimes not allowed because of its toxicity.

• Yield strength • Stiffness • Specific mass • Thermal expansion coefficient • Manufacturability • Shelf life • Outgassing • Non-magnetic* • Raw material cost • Order lead time

MaterialUniverse – Ansys Granta


Manufacturing Technique

• CNC Machining
• CNC Turning
• Wire EDM (Wire cut)
• Laser cut
• 3D printing
• Honeycomb sandwich panel
• Injection molding


Design for Manufacturing

✔ Reduce the total number of parts
✔ Develop a modular design
✔ Use of standard components
✔ Design parts to be multi-functional
✔ Design parts for multi-use
✔ Design for ease of fabrication
✔ Avoid separate fasteners
✔ Minimize assembly directions
✔ Maximize compliance
✔ Minimize handling



References and Resources also include:



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