4. Structures and Mechanisms

4.4 General Arrangement and Design Drivers

This section will review the design drivers that affect the structural design of the spacecraft: geometry, mass, structural loads, materials, and processes. These design drivers affect each other, sometimes beneficially, but also in ways that oppose each other and cause problems. The design process is always iterative and must consider all of these design drivers.

“The appropriate design and test factors for a given mechanical or structural flight hardware element depend on several parameters, such as the materials used, attachment methods (e.g., bonding), and the verification approach (prototype or proto flight). In addition to the minimum factors of safety specified in this NASA Technical Standard, some structural and mechanical members may be required to meet other more stringent and restrictive performance requirements, such as dimensional stability, pointing accuracy, stiffness/frequency constraints, or safety requirements (e.g., fracture control)” [NASA-STD-5001B].

Geometry

The structure of the EST CubeSat. Notice the placement of all the parts relative to each other. Which components are on the outside? Which is facing out? CC SA 3.0. Image by the University of Tartu.

For the primary structure, the design drivers are the first and foremost requirements, derived from external constraints and internal needs. The primary structure is constrained to the launch vehicle fairing or deployer enclosure and must enclose all the spacecraft bus components. Geometric considerations that affect every subsystem could include:

Subsystem Consideration
Payload
Structure and Mechanisms
  • Spatial organization to ensure components do not intersect
  • Vibration isolation
Thermal Control
  • Facilitate thermal management through conduction or radiation
  • Regulating thermally sensitive components through placement
  • Enabling thermal isolation
Power (including harness)
  • Securing harnessing in empty spaces
  • Insulating battery (which is typically thermally sensitive)
Telemetry and Control
  • Unoccluded field of view for antenna
Command and Data Handling
  • Radiation shielding by placing the computer behind other components
Attitude Determination and Control
  • Sensor and actuator mounting in defined orientations
  • Regulating moments of inertia (emphasize one principal axis or uniform across axes)
  • Minimize off-diagonal moments of inertia
Propulsion
  • Direction of thrusters
  • Placement of propellant exhaust exit with respect to payload optics

Artemis CubeSat Kit Arrangement

Artemis Kit Specific
The Artemis CubeSat kit is straightforward in its geometric arrangement as the subsystem components are rather homogenous in their geometry; all the subsystems are generally mounted on PCB boards and stacked on a threaded rod as seen in the figure. The antenna and deployer on the exterior of the satellite guarantee an unoccluded field of view. The exterior skin has a cut-out to allow a camera to peer through; the first board in the stack has the payload and payload-supporting electronics. The power distribution board follows. The onboard computer sits nearly in the center of the CubeSat, providing radiation shielding. Between boards starting from payload to onboard computer, 104-pin Cubesat kit bus headers are used to reduce clutter. Pycubed boards were not designed to be compatible with the CubeSat kit bus header and thus integrated through external connections. The low-level control computer called the PyCubed, is stacked underneath. Finally, we have the PyCubed battery board, which runs warm and needs exposure to the space environment to radiate its heat.

Mass

Labeled Printed Circuit Boards of Artemis 1U CubeSat with No Solar Panel Boards; A: Payload Board; B: Mock Payload; C: Onboard Computer, D: Power Distribution Unit; E: Battery Board; F: Antenna Board; G: Antenna

The structure and mechanisms subsystem specialist is an important player in generating and managing the mass budget, with the assistance of the systems engineer and the other subsystem specialists. For a 1U CubeSat, the total mass of the spacecraft must not exceed 2 kg [CubeSat Design Specification Rev. 14]. Any additional mass may be negotiated with the launch provider with an immense amount of paperwork and persistence but it’s not impossible. Typically, 1U CubeSats are between 1 kg and 1.33 kg. To reiterate, a suggested mass budget and specific 1U CubeSat project mass budgets are as follows:

Subsystem (% of Dry Mass) SMAD suggestion Hermes CubeSat Artemis CubeSat
Payload 41% Allocated in T&C 2%
Structure and Mechanisms 20% 32.3% 20%
Thermal Control 2% 0%
Power (including harness) 19% 13.5%
Telemetry and Control 2% 22.5%
Command and Data Handling 5% 3.6% 5%
Attitude Determination and Control 8% 2.4% 8%
Other (balance launch) 3% 25.7% –
Total 100% 100% 100%

The mass budget typically carries a margin in the preliminary design phase. The design margin decreases over time as the design converges to the final assembly. Refining the design toward spaceflight reveals additional interfacing and detailing that inevitably adds mass to the system [Hayhurst et al.]. Each subsystem’s mass growth by design gate is shown in the figure. Note, that the spacecraft studied are traditional in size and mass, which means that the study was not geared toward cube satellite design.

Historical Mass, Power, Schedule & Cost Growth for NASA Instruments & Spacecraft by Marc Hayhurst, Robert Bitten, Daniel Judnick, Ingrid Hallgrimson, Megan Youngs The Aerospace Corporation. Stephen Shinn NASA Goddard Space Flight Center
Historical Mass, Power, Schedule & Cost Growth for NASA Instruments & Spacecraft by Marc Hayhurst, Robert Bitten, Daniel Judnick, Ingrid Hallgrimson, Megan Youngs The Aerospace Corporation. Stephen Shinn NASA Goddard Space Flight Center

Structural Loads

This section will provide an overview of typical structural loads and how they drive the spacecraft’s structural design. Loads are generated by forces, deformations, or accelerations that cause stresses, deformations, and displacements in structures. There are two types of structural loads: static and dynamic. Static loads are steady-state loadings, like loads imparted on spring-loaded deployers, launch acceleration, or pressurized vessels. Think of these loads as built-up loads that are ready to burst or buckle. Dynamic loads are loads from vibrations generated by natural frequencies, like launch vehicles, pyrotechnic separation, or deployment events. Think of these loads as shocking events. Structural engineers are concerned with mitigating the effects of the critical load: the load that the spacecraft most intensely feels and is most likely to break the spacecraft. Critical loads could be launch loads for an assembled spacecraft, pressurization loads for a rocket casing, thermal loads for a propulsion system, centrifugal forces from rapid rotation, or on-orbit collisions.

Three 1U CubeSats beside a 3U (Poly Picosatellite Orbital Deployer (PPOD) developed at CalPoly. The spring mechanism used by P-PODs to deploy CubeSats can be seen within the main housing, prior to loading. Image Credit: California Polytechnic State University Source publication
The Fox-1A CubeSat satellite has been integrated into the Poly-PicoSatellite Orbital Deployer rig (P-POD) with two other CubeSats. The red arrows show the static load on the CubeSats generated from the compressed spring’s force. Image by Dan Passaro.
Apollo 15 Launch and Reaching Earth Orbit. Courtesy of NASA.
Typical vibration spectrum of a launch. How to test satellites and not destroy them by Ben Sampson Courtesy of Aerospace Testing International.

Just as we reviewed every phase of the spacecraft lifecycle in the Typical Requirements section, we will revisit these phases to identify all loads and estimate the load. Load quantification may be obtained through measurements, tests, references, and asking the relevant engineers. Critical load estimation is not always straightforward and may need to be indirectly quantified or estimated.

  • During manufacturing and assembly, stresses could include welding, joint stressing due to tightening bolts
  • During transport and handling, requirements may include loads from transferring the spacecraft to a shipping container (especially for large spacecraft) and shock during transportation through land/sea/air transport environments (like on freight boats or trucks). The Artemis CubeSat Kit expects to be handled delicately by human hands, which yield gentle loads. The kit is not expected to survive being dropped. The kit will arrive fitted snugly in fitted foam within a Pelican case, which mitigates the shock loads during transportation. These loads will not be critical loads.
  • During thermal and vibration testing, critical loads could be the stress from misaligned thermal expansion, launch shock, acceleration loads, and random vibration environments. These tests replicate launch loads and thermal stressing from the mission operations environment. The difference between thermal and vibe testing in this phase with respect to the real environments could be the disparity in replicating the same loads in which case the spacecraft must survive two different loading profiles. Ideally, the test matches the launch and space environment conditions so we will discuss critical loads in those phases. A summary of tests is seen below:
    Launch Services Program. Courtesy of NASA.
  • During pre-launch, the spacecraft must be handled and packaged into the launch vehicle. For large spacecraft, handling loads could include static loads at hoisting interfaces. CubeSats experience the compression of the P-POD deployer. Drawing from the Nanoracks External CubeSat Deployer Document, “The CubeSat shall be capable of withstanding a force 1320N across all load points equally in the Z direction”. The Artemis CubeSat kit was analyzed to withstand this integrated load. This could be a critical load for which we should do structural analysis.
  • During launch and ascent, the structure must withstand steady-state booster accelerations, vibroacoustic noise during launch and transonic phase, propulsion system engine vibrations, pyrotechnic shock from separation events, transient loads during stage separations, etc. Generally, the critical loads to launch loads as these loads are the most intense out of any phase.  The Artemis CubeSat Kit has been tested on a vibration table to withstand these loads per the Launch Services Program Level Dispenser and CubeSat Requirements Document [NASA LSP-REQ-317.01] and NanoRacks External CubeSat Deployer (NRCSD-E) Interface Definition Document (IDD) [NR-NRCSD-S0004]. The CubeSat may be soft-stowed on a resupply mission to the ISS or hard-stowed as a secondary payload. We’ve tested both profiles. For your convenience, we have listed all relevant loads taken word for word from the NanoRacks document.
  • Acceleration loads: Payload safety-critical structures shall (and other payload structures should) provide positive margins of safety when exposed to the accelerations documented in Table 4.3.1-1 at the CG of the item, with all six degrees of freedom acting simultaneously.

  • Random Vibration Environment: The CubeSat shall be capable of withstanding the dynamic flight environment for the mission-applicable launch vehicle (shown in Table 4.3.2.1-1 through Table 4.3.2.1-4). Nominally, NRCSDE missions are launched on the Antares rocket; however, Atlas V rockets have been utilized in the past.

  • Launch shock environment: The CubeSat shall be capable of withstanding the shock environment shown in Table 4.3.3-1. Any mechanical or electrical components on the spacecraft that are highly sensitive to shock should be identified and assessed on a case-by-case basis as defined in the unique payload ICA.

  • During mission operations, loads include thruster acceleration, transient loads from pointing maneuvers, docking events, pyrotechnic shock from separation or deployment, and loads from thermal expansion. The Artemis CubeSat Kit experiences antenna deployment and thermal expansion for which both the thermal vacuum chamber and antenna deployment tests verified survival. For your convenience, we have reiterated the thermal environment taken word for word from the NanoRacks document.

  • The CubeSat shall be capable of withstanding the expected thermal environments for all mission phases, which are enveloped by the on-orbit EVR phase prior to deployment. The expected thermal environments for all phases of the mission leading up to deployment are below in Table 4.3.5-1.
  • In the final phase of reentry and landing, spacecraft may experience aerodynamic heating and pressure, transient winds, or landing loads. The Artemis CubeSat Kit need not survive reentry as it is designed to burn up upon reentry.

Upon identifying the various loads, we may conduct structural analysis to determine which of these loads is the critical load. From experience and intuition, a good guess is to design the structural components of the spacecraft to the launch conditions if your spacecraft will remain in orbit. If your spacecraft will re-enter Mars’s atmosphere, for example, the entry, descent, and landing phases may be more critical. The structural analysis will be described in the last section of this chapter but the consequence of the structural analysis is the iterative design and redesign of the structural components to fulfill sufficient margins of safety.

Structural analysis plays into the initial design of structures by conducting back-of-the-envelope (simplified) calculations as to the sizing or thickness of a structural component, like the primary structure wall or supporting bracket. The structural analysis enters the redesign phase by showing that some structural components fail at the critical load and need reinforcement to achieve mission success. Structural analysis may also show some structural components more than sufficiently carry that piece’s critical load and could be trimmed in mass to allocate elsewhere. Finally, structural analysis in the way of finite element analysis is a critical method of verifying that structural designs will survive tests or survive conditions that would otherwise be infeasible to test.

Materials

The selection of the structural material affects the survivable structural load, mass, geometry, and concerns around outgassing. Material properties include density, stiffness, strength, weight, ductility, coefficient of thermal expansion, fatigue, and outgassing:

  • Density is the mass per unit volume of a material. As space missions are proportional in cost to the mass launched into space, lower-density materials are preferred.
  • The precise term for material stiffness is Young’s modulus, which “defines the relationship between stress (force per unit area) and strain (proportional deformation) in a material in the linear elasticity regime of uniaxial deformation. Young’s modulus enables the calculation of the change in the dimension of a bar made of an isotropic elastic material under tensile or compressive loads” [Wikipedia]. This value is commonly represented by the letter E or Y.
    • E = \tfrac{\sigma}{\epsilon}
    • E is Young’s modulus
    • \sigma  is the uniaxial stress or uniaxial force per unit surface
    • \varepsilon  is the strain, or proportional deformation (change in length divided by original length); it is dimensionless
    • Both E and \sigma have units of pressure, while {\varepsilon is dimensionless. Young’s moduli are typically so large that they are expressed not in pascals but in megapascals (MPa or N/mm2) or gigapascals (GPa or kN/mm2).
  • There are two types of material strengths that we care about yield strength and ultimate strength.
  • Ductility “is a measure of a material’s ability to undergo significant plastic deformation before rupture or breaking, which may be expressed as percent elongation or percent area reduction from a tensile test” [Wikipedia].

% EL=  \tfrac{\text{final gage length}-\text{initial gage length}}{\text{initial gage length}}  =  \tfrac{l_{f}-l_{0}}{l_0}}*100

Malleability is the compressive counterpart of ductility. Malleability “is a material’s ability to deform under compressive stress”.

The schematic appearance of round metal bars after tensile testing. (a) Brittle fracture (b) Ductile fracture (c) Completely ductile fracture. Image by Sigmund.
  • The coefficient of thermal expansion (CTE) “describes how the size of an object changes with a change in temperature. Specifically, it measures the fractional change in size per degree change in temperature at constant pressure” [Wikipedia]. You’ll notice bridges or parking lot structures have expansion joints that fill gaps within the structure and act as a flexible, variable filler that helps the structure adapt to temperature changes without distorting [Science Clarified].
  • Fatigue occurs when a material is cyclically loaded and unloaded at mean stress. Fatigue limit “is the stress level below which an infinite number of loading cycles can be applied to a material without causing fatigue failure” [Wikipedia]. Interestingly, aluminum seemingly has no fatigue limit. “Fatigue failures, both for a high and low cycle, all follow the same basic steps process of crack initiation, stage I crack growth, stage II crack growth, and finally ultimate failure” [Wikipedia]. Characteristics of fatigue include randomness in the location of the failure, usual association with tensile stresses, inverse relationship between applied stress and life, and irreversible damage. “Fatigue life is influenced by a variety of factors, such as temperature, surface finish, metallurgical microstructure, presence of oxidizing or inert chemicals, residual stresses, scuffing contact (fretting), etc.”, which is why attention to manufacturing processes is important to preserve the structural integrity of components likely to fatigue.
    Representative curves of applied stress vs a number of cycles for steel (showing an endurance limit) and aluminum (showing no such limit). Image by Andrew Dressel.

    Representative curves of applied stress vs a number of cycles for steel (showing an endurance limit) and aluminum (showing no such limit). Image by Andrew Dressel.
  • Outgassing or off-gassing is the “release of gas that was dissolved, trapped, frozen, or absorbed in some material” [Wikipedia]. Outgassing commonly occurs when the spacecraft is exposed to a high-vacuum environment. NASA keeps a database of outgassing data of materials intended for spacecraft use and promotes the use of materials with low-outgassing properties. “Outgassing products can condense onto optical elements, thermal radiators, or solar cells and obscure them. For most solid materials, the method of manufacture and preparation can reduce the level of outgassing significantly. Cleaning of surfaces or heating of individual components or the entire assembly (a process called “bake-out“) can drive off volatiles” [Wikipedia].

Common choices for spacecraft structures include aluminum, steel, titanium, and composites. Aluminum is incredibly common due to its high material strength with relatively low density to save on mass and low cost. Steel is stronger and generally cheaper but heavier. Titanium is stronger and lighter but much more expensive. Composite materials are higher in strength and lower in density, also making them attractive candidates, but have less space heritage or historical use. For a more quantitative comparison, refer to the table below:

Material Aluminum 6061-T6 Stainless Steel 316 Titanium Ti-6Al-4V Carbon-carbon composite
Density 2.7 g/cc 8 g/cc 4.43 g/cc 1.6 g/cc
Young’s Modulus 68.9 GPa 193 GPa 113.8 GPa 80 GPa
Tensile Yield Strength 276 MPa 290 MPa 880 MPa 260 MPa
Tensile Ultimate Strength 310 MPa 580 MPa 950 MPa
Ductility 12 – 17 % 50 % 14 %
CTE 23.6 – 25.2 µm/m-°C 16 – 17.5 µm/m-°C 8.6 – 9.7 µm/m-°C 0.2 – 5.7 µm/m-°C
Fatigue Strength 96.5 MPa 270 MPa-N/mm2 510 MPa
Outgassing rate 33 x 10^-9 \tfrac{Pa - m^3}{s -m^2} 5.1 × 10−9 \tfrac{Pa - m^3}{s -m^2} 10 x 10^−9 \tfrac{Pa - m^3}{s -m^2} \tfrac{Pa - m^3}{s -m^2}
Thermal Conductivity 167 W/m-K 16.3 W/m-K 6.7 W/m-K 4 – 27 W/m-K
Cost for ¼” x 1” x 1’ bar 3.46 USD 23.31 USD 98.46 USD 20.50 USD

Process

Although the process may seem like an afterthought, we must consider the manufacturing, integration, assembly, and testing process. The design may be geometrically elegant or structurally strong but the design is not feasible if the structure components are impossible to manufacture or assemble. The most straightforward way to gauge if a design is possible to fabricate, assemble, and test is to attempt to fabricate, assemble, and test. Infeasible plans may be revealed through preliminary plans, like consulting with machinists on part drawings or generating an integration procedure. The best practice is to fabricate and test prototypes prior to the actual deadline to iron out any hiccups in the implementation progress.

ALTEN engineers during assembly, validation tests, and final integration (before launch) of a satellite. Image by ALTEN.

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A Guide to CubeSat Mission and Bus Design Copyright © by Frances Zhu is licensed under a Creative Commons Attribution 4.0 International License, except where otherwise noted.

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