Spaceborne Experiments
18.4 Space Mission Design .1 The Qualification Program
Since repairing in space is extremely expensive, if at all possible, designing and building spacecraft and payloads which maximal reliability is a must in the field of space engineering. It follows that quality control is an essential part during the various phases of the program. TheQualif ication P rogramadds to the cost of the space hardware construction, sometime very significantly, but it makes sure that the program is not headed for failure.
Qualification tests must be designed and implemented to check that the space- craft/payload can withstand the challenges of launch, deployment and operation in space. Subsystems and components environmental tests include vibration, shock and thermal vacuum, electromagnetic compatibility and radiation hardness.
Although the goal is the same, testing strategies are not unique. There are indeed various testing methods:
• dedicated qualif ication hardware(QM): a set of qualification components is built and tested at qualification levels. A set of flight components (F M) is then built and launched after passing a qualification test a lower levels;
• proto−f light approach: a set of flight components is tested at qualification level then assembled into a subsystem or payload which is tested at qualification levels and then launched;
• similarityapproach: demonstrate that the components and the environment are identical to previously qualified hardware.
A typical test sequence includes a series of functional tests preceeding/following each environmental test, for example:
• functional test;
• vibration test (levels depending on the mission);
• functional test;
• shock test (levels depending on the mission);
• functional test;
• thermal-vacuum tests, including some functional tests during exposure;
• Electro Magnetic Compatibility (EMC) tests (if required);
• flash X rays with functional tests during exposure (if required);
18.4.2 Vibration and Shock Test
A payload must withstand vibrations caused by the launch vehicle and transmitted through its structural mount. During launch, payload components may experience shocks due to the explosives used for the separation of the various stages. In case reentry is foreseen, they do experience shocks when entering the atmosphere as
well as during the landing phase. In order to understand the dynamical behavior of the payload and of its mounting under these circumstances, Finite Element Analysis (F EA) dynamic and numerical analysis together with Computer Aided Design (CAD) simulation should be performed. In this way it is possible to search for resonances of the mechanical structures, identifying conditions where the material could be stressed or damaged. Following an iterative process the mechanical design of the payload and of its mountings can be improved until all negative margins are eliminated. Dynamic and vibration tests are then performed on a qualification model, using for example an electro-dynamical shaker operating at frequencies between 5 and 3000 Hz, with a spectrum which depends on the mission characteristics. Table18.2shows a typical acceleration spectrum expected for payload launched using the Shuttle transportation system. Qualification levels are typically higher by factor 2 to 4. Shock tests are performed using a similar strategy. For example Fig.18.6shows shock levels used to simulate the launch of an Alpha-Centaur rocket.
Table 18.2 Maximum expected flight levels for a shuttle mission
Frequency range Frequency dependence Xaxis 20–58 Hz 0.0025g2/Hz
58–125 Hz +9 dB/Octave 125–300 Hz 0.025g2/Hz 300–900 Hz −9 dB/Octave 900–2000 Hz 0.001g2/Hz Overall=3.1 Grms
Yaxis 20–90 Hz 0.008g2/Hz 90–100 Hz +9 dB/Octave 100–300 Hz 0.01g2/Hz 300–650 Hz −9 dB/Octave 850–2000 Hz 0.001g2/Hz Overall=3.1 Grms
Zaxis 20–45 Hz 0.009g2/Hz 45–125 Hz +3 dB/Octave 125–300 Hz 0.025g2/Hz 300–900 Hz −9 dB/Octave 900–2000 Hz 0.001g2/Hz Overall=3.1 Grms
18.4.3 Environmental Tests
The environmental qualification campaign of a space component can be divided into three main steps. The first step consists in the development of requirements and constraints related to the payload and to the mission. The second step is
Peak density [g2 /Hz]
10,000
1,000
100
10100 1,000 10,000
Frequency [Hz]
Qualification level
Expected level
Fig. 18.6 Shock levels simulating the launch environment of an Alpha-Centaur rocket: qualifica- tion levels are designed to be greater than the expected design values [12]
to determine and define the space environment (in terms of temperatures, heat transfer ways, worst hot and cold case, etc.) that will characterise thermal conditions throughout the entire life of the component. An important part of the process of qualification is then the thermal analysis which can be conducted usingF EA techniques. Once an acceptable thermal model has been developed, test predictions can be calculated to correlate thermal verification tests with the test results. If this correlation is found acceptable, the thermal model is then used to perform flight predictions. If, instead, the correlation is poor the thermal analysis and the hardware configuration need to be carefully checked to understand whether the actual configuration (hardware) requires modifications or the thermal model needs to be updated. Payload temperature requirements derive from the spacecrafts thermal design and the orbital environment and attitude. The purpose of these tests is to demonstrate that the subsystems comply with the specification and perform satisfactorily in the intended thermal environment with sufficient margins. The test environment should be based either on previous flight data, often scaled for differences in mission parameters, or, if more reliable, on analytical prediction or by a combination of analysis and flight data. A margin can include an increase in level or range, an increase in duration or cycles of exposure, as well as any other appropriate increase in severity of the test. Humidity and thermal qualification tests in climatic rooms are performed to test the behaviour of the electronic components and mechanical structures under thermal and humidity changes. The tests are conducted using climatic chambers, with temperature ranges depending on the mission parameters: for a LEO mission typical range lays within −80◦C and +120◦C for a planetary mission wider intervals are required. Components should be switched on and work both at temperature extremes or during transition, following the mission specifications (see Fig.18.7).
Hot storage [Q: 85, F:80]
Hot operating [Q: 55, F:80]
Cold operating [Q: –25, F:–20]
Cold storage [Q: –45, F:–40]
Ambient
1H
1H
1H 1H
1H
1H
1H
1H
1H 1H
1H
1H 2H cold soak
2H cold soak 2H
hot soak 2H
hot soak
Cold balance
Trancient cool down Hot
balance Power on Power of
A B C D E F G H I J K L M N O P Q
Return to ambient Thermal balance
phases Thermal cycles phases
Pump down
Fig. 18.7 Typical Thermal Vacuum Cycling test profile for a payload to be operated on a LEO orbit on the ISS [54]
18.4.4 EMC Tests
Another area where space payloads are submitted to extensive testing is the compatibility to Electro Magnetic fields, either radiated or received. During the test of radiated EM the device under test is powered and operated in standard operating condition in an EM anechoic chamber. Through suitable antennas and filters read by receivers, the intensity of the emitted radiation is measured as a function of the frequency. The results are compared with the limits requested by specific standards or design rules. If the limits are exceeded, then the electrical grounding or design of the device should be modified. During the received EM test, the device is operated within an EM anechoic chamber while EM radiation, monochromatic or with a specific spectral structure, is generated at a predetermined intensity using special antennas located nearby. The purpose of the test is to check that the item under test does not exhibits anomalies when illuminated by beams of EM radiation, typically emitted by a communication antenna or a nearby electronic device. Figure18.8show a typical the result for an EMC radiated test on a payload to be operated on the ISS.
18.4.5 Radiation Hardness Tests
As discussed in the previous paragraphs, the space environment is particularly harsh for operating microelectronics devices, due to the presence of single, heavily
90 80 70 60 50 40 dBµV/m 30 20 10 0 –10 –20
14 kHz 100 kHz 1 MHz 10 MHz 100 MHz 1 GHz 10 GHz
Fig. 18.8 Typical radiated EMC tests result for a payload to be operated on a LEO orbit on the ISS [19]. The thin line presents the e.m. field limits which should not be exceeded as described in Table18.3
Table 18.3 Radiated EMC limits for the tests described in Fig.18.8
Frequency range Emission [dBμV/m] Antenna
14 kHz–10 MHz 56 rod—vertical
10 MHz–259 MHz 56–86 (16 dB/decade) biconical—horiz/vert
259 MHz–10 GHz 46–72 (16 dB/decade) double ridge—horiz/vert
10 GHz–20 GHz 76 horn—horiz/vert
ionizing particles which can deposit large amount of charge in the bulk, inducing short circuits or spurious currents in the solid state circuits. The total dose collected during a space flight is relatively small, mostly in thekrad range, so the radiation damages are mostly due to Single Event Effects (SEE).
Depending on the type of circuits and on their construction technology, the sensitivity to ionizing radiation can be very different. In order to select families of commercial circuits which are more insensitive than others it is necessary to run testing campaigns, comparing the behavior of several different chips when exposed to low energy ion beams. The type circuits which showslatch up sensitivity or abnormal behavior only at high Linear Energy Transfer (LET) are the one which are radiation hard and can be used in space. In order to ensure statistical significance of the radiation hardness measurement, several chips of each type should be tested (typically>5).
Often, it is possible to protect the circuit by limiting the current which can flow through the power lines, by the use of an active switch which temporarily cut off the voltage to stop thelatch upeffect. In order to develop and implement a protection scheme it is then important to understand which area of the chip is sensitive. Nowadays it is also possible to perform in laboratory part of these test
by mean of IR laser beams which are absorbed by the silicon and can deposit a controlled amount of energy into the bulk simulating the charge released by a low energy ion [21].
All microelectronics components used in a space experiment must be radiation hard. In addition, the design of the on board electronics should include multiple redundancy since the radiation damage due toSEEis a stochastic process.
Space qualification of radiation resistant devices for a mission not only requires the understanding of damage mechanisms [22], but also the knowledge of local particles (and species) intensities [23, 24] and, in addition, of dose amounts, deposited via ionization and non-ionization energy loss (NIEL) processes. The latter mechanism is that one responsible for displacement damages particularly relevant for semiconductor devices. Only recently, the SR (screened relativistic) NIEL treatment framework has allowed a comprehensive calculation of NIEL doses imparted by electrons, protons, ions and neutrons in any material and compound [25]. SR-NIEL treatment is currently embedded in ESA transport codes, like GRAS [27] and MULASSIS [28] as well as in GEANT4 and it is available at the SR-NIEL and SPENVIS websites.