It takes typically ten years to establish acceptance of a material for aerospace applications. Space applications, by comparison, require rather small volumes of such materials and so manufacturers cannot therefore rely on space alone for their business. In addition, emerging materials tend to be much too expensive in the shorter term for a particular application when compared with established materials, the development costs for which have already been wriitten of. The experience gained in using an established materialis another important factor.
Consequently,there is an urgent need to find ways and means of reducing such costs, which involves promoting broader access to the emerging technology.The establishment of a bigger market and a bettersharing of development costs would seem to require improvements in technology transfer. Forums for co-operation and cross-fertilisation must be exploited to identify a wider range of opportunities to extend the use of new materials and their applications technology.
There has already been some success in the transfer of technology of structurally related space application materials to other disciplines. These include the use of high-modulus,carbon reinforced plastic in ground-based applications, surgical uses for shape-memory alloys, and industrial use of rocket-motor static seal technology.
In March this year, ESA co-sponsored a Symposium, with BRITE/EURAM, EUREKA and SAMPE, with a view to stimulating such collaboration. Russian contributions were also actively encouraged. The figures used to illustrate this article are drawn from the Proceedings of that Symposium. ESA continues to support such bodies as AECMA and has its own Advanced Structural Materials Information Exchange Group, which endeavours to coordinate such activities. It has been active in stimulating and reviewing information for ESA's Structural Handbooks.
In the short term, the thrust must be to conserve and consolidate the most promising developments and to foster the wide spread and cost-effective utilisation of this materials technology.
As materials reach an appropriate stage of development, they can be candidates for use in:
'Smart structure' developments are expected to find application in the realisation and maintenance of reflect or profiles, dimensional stability of optical payloads, active noise and vibration control, and in-orbit health monitoring and possibly control of unmanned spacecraft. The ability to monitorthe health of large cryogenic composite fuel tanks is a typical long-term objective.
Long-life orbiting structures will require dimensionally stable materials. To date, such needs have largely been met with carbon-fibre epoxy resin composites (CFRP), but the existing epoxy composites may not sustain such requirements for larger components and extended lifetimes under thermal cycling.
Opportunities exist to replace first-generation CFRP with stronger ultra-high-modulus CFRP materials utilising cyanate ester or toughened epoxy resin systems. The cyanate esters offer reduced moisture absorbency and better resistance to micro-cracking under severe thermal cycling.
A recent ESTEC study showed that low instances of micro-cracking could be achieved in the thermal cyclic range of - 180 degC to + 130 degC.
However, the precise 'prepreg' processing conditions applied by individual contractors have proved very influential in determining the integrity of the thin (300 micron) 0 degree /90 degrees laminates.
Optical structures require a high degree of dimensional stability, so that the propensity of CFRP to outgas moisture and micro-crack may prove unacceptable. Composites with more stable matrices such as metals and ceramics are proposed as solutions here, particularly for Low Earth Orbit (LEO) applications.
Carbon-fibre-reinforced magnesium and aluminium composites have received attention by of fering opportunities for a high degree of Coefficient of Thermal Expansion (CTE) control. Considerable efforts have been expended in developing processing techniques to overcome the difficulties of fibre-matrix reactions and poor transverse strengths. The isssue of galvanic corrosion remains, but may be overcome if the composite is coated and always kept in a controlled low-moisture environment.
Carbon-fibre/ceramic-matrix composites (C- C and C- SiC), which have been developed for high-temperature applications, can be modified for use in optical structures such as mirrors and optical benches.They offer very high stiffness and the stability of materials that have been processed at high temperatures well beyond those likely to be experienced by orbiting structures. However, all such ceramic- and metal-matrix composites are very expensive to produce.
There is concern that polymeric materials in particular may experience accumulated physical damage during prolonged use on low Earth-orbiting structures. Contributing conditions include the hard vacuum , radiation, thermal cycling and micrometeoroid impacts, coupled with atomic-oxygen attack. These provide conditions for progressive surface degradation, which may be significant over a 25-year lifetime compared with the 5 to 7 years of exposure typical of today's orbiting spacecraft.
Experiments aboard the US Long-Duration Exposure Facility (LDEF) and Europe's own Eureca retrievable platform have quantified the levels of erosion that can be experienced by uncoated CFRP. In the worst case, some 100 microns of composite, equating to a single ply, can be removed in just 5 years of exposure to atomic oxygen. For typically thin constructions, the integrity of the structure would be lost before the end of the spacecraft's intended service lifetime. Such erosion can be avoided by adopting one or more of the following options: protective metallic coatings, shielding or reduced angle of incidence to atomic-oxygen flux.
A more ambitious option is to use Anoxic Siloxane 'composites' (described in detail by S. Palsule in ESA Journal No. 2, 1993). These materials have silicon instead of carbon in their molecular structure and it is converted to a non-volatile silicon oxide by the atomic-oxygen flux. A type of 'passivating layer' can be formed which is stable in vacuum and inhibits further penetrative oxidation. Siloxanes can therefore be exploited in the first instance as coatings. They are currently being evaluated as matrices for structural composites with carbon-fibre reinforcement.
Figure 1. Extensive use of diffusion bonding and superplastic forming is made in the manufacture of the Rafale fighter's highly loaded titanium canard. This incorporates a diffusion-bonded spigot and a hexagonal corrugated core structure. A substantial mass saving is realised compared with other manufacturing methods (Photo courtesy of Dassault Aviation)
Within Europe, high-temperature structural materials development has received an impetus in recent years from the Hermes, Sänger and Hotol spaceplane programmes. The potential materials for such applications now include new titanium alloys, Super-Plastically Formed/Diffusion Bonded (SPF/DB) titanium, titanium-matrix composites (TMC), nickel alloys, beryllium and ODS aluminium alloys, as well as the more well publicised C- Cand C-SiC ceramic-matrix composites. The metallic materials offer capabilities in the 300 to 1000 degC range, whilst Ceramic Matrix Composite (CMC) materials in various selected forms can sustain temperatures of 1500 to 2000 degC. The issue of oxidation protection for prolonged use is not to be underestimated. Elaborate protective coatings are expensive to apply and increase the complexities of joining, inspection and verification.
With C-SiC, more recent developments have concentrated on reducing the very high processing costs associated with chemical vapour deposited or infiltrated (CVD/CVI) C- SiC by switching to quicker processing routes such as siliconising, polymer pyrolysis and melt infiltration. All of these materials can provide technical solutions in some form for applications experiencing operating temperatures beyond the capabilities of polymer composites and aluminium alloys. Each material has its own particular attributes and drawbacks, not least in who retains the capabilities to design and manufacture with a specific material.
For the high-temperature composites with continuous fibres, the material microstructures are very complex, as are the progressive modes of failure through multiple crack formation. The integrity of such materials under harsh oxidising conditions remains to be fully resolved, since they are of ten operating close to their physical limits and experiencing micro-structural changes during use.
The short- to medium-term future of many high-temperature materials is very uncertain due to the very limited commercial applications available. In space programmes, the shorter-term opportunities have diminished to primarily those associated with re-entry vehicles, launcher propulsion systems, and planetary probes. Some C-SiC may be used in optical structures. The volume of material used in these outlets will, however, be insufficient to guarantee the commercial viability of all such materials.
The other industrial opportunities for the new materials lie with gas-turbine engine development, high-energy braking systems, and electrical power generation. These applications have different service requirements, however, making the likelihood of common material usage somewhat tenuous.
An obvious conclusion from all of this is that future space programmes using modest quantities of materials cannot readily expect commercial sourcing to a unique specification. The result could be considerable expense in fabricating a limited number of items especially for space application.
With the application of new materials comes the requirement to fabricate structures from them with efficient joining techniques. The assembly costs for such large structures have also to be weighed carefully, particularly for a launcher like Ariane-5, with its many cylinders, bulkheads, frames, adapters and panels. Even with established materials such as aluminium alloys (2219,7475) and steels (D6AC), useful advances can be made by applying modified techniques. For the aluminium alloys, this includes spin-forming/Tungsten Inert Gas (TIG) welding (Ariane-5 tanks), cold forging and machining (Ariane-5 booster separation motors), shot peening/rolled/forged and TIG welding (Ariane-5 tank bulkheads) and electron-beam welding. This, in turn, could lead to the next generation of tank constructions based on cold-formable and SPF/DB titanium alloys, weldable lithium containing aluminium alloys (2195, Weldalite), and filament-wound CFRP constructions. Such technologies could save up to 2 tons on an Ariane-5 vehicle, but the extra development costs would have to be properly identified.
Where high-temperature materials are to be applied, the main issues relate to connecting hot outer surfaces to cooler underlying airframe structures. Joints may be static or may consist of movable parts and seals. ESTEC-funded programmes have evaluated fasteners made made of C-C, C-SiC, coated nobium alloys and nickel alloys. All can sustain the temperatures involved, but the issue of oxidation and protecting threaded fasteners requires further investigation, as do the appropriate inspection techniques.
In preliminary studies, the use of castable inorganic cements has been demonstrated for a configuration representative of a CMC thermal-protection shingle mounting. By careful design, the cementis injected into a cavity where it solidifies and provides a mechanical taper to lock the shingle in place. The cement sustains any applied loads by compression, for which it is well-suited, as opposed to the shear that bonded joints experience. As ceramic composites are low strain to failure materials, it is preferable to configure joints to operate in compression and not in tension or shear.
Figure 2. Spin-rolling facilitates the production of very large, seam ess, stiffened cylindrical parts, as illustrated by this 6.2m long Ariane component (Photo courtesy of MAN, Germany)
There are now a large number of established Non-Destructive Testing (NDT) techniques for inspecting assembled composite and light-alloy structures. To these have been added modified systems for ceramic and metallic composites, where the defects and features of interest may be similar, but on a smaller scale and in mediums of different signal attenuation and energy absorption. Greater emphasis is now being placed on non-contact and non-coupled (no water) techniques as means of examining assembled configurations. Ultrasonic inspection has always been popular, but is usually a contacting/coupled method. It is now giving way to techniques such as: laser ultrasound, eddy currents, shearography, electronic speckle thermography, transient thermography, photo thermal thermography and stimulated infrared thermography in addition to holographic interferometry, neutron magnetic resonance, and neutron absorption, X-ray techniques based on radioscopy, beta-particle backscattering, refractography, micro-focussing and and tomography have contributions to make for high-temperature composites such as coated C-C and C-SiC. For high-resolution inspection of ceramic composites, computerised tomography is gaining acceptance.
Whilst all of these techniques have something to offer in providing various degrees of resolution, the types of defect and damage that need to be detected remain unchanged. For polymer composites, the critical defects are de-laminations and impact damage, with supporting evidence required to confirm low porosity, correct fibre alignment and no disbonds. With ceramic composites, acceptable porosity levels can be higher, but any variations in bulk density and thickness need to be quantifiable. Severe delaminations and excessive micro cracking are the main features to be looked for before confirming material integrity.
This is a rather special and emerging structural materials development area and often involves the embedding of sensors, detectors and optical fibres in a composite structure. In conjunction with the use of these devices, passive or active control over the structure's behaviour can be achieved using a controlling network. The realisation and maintainance of large deployable reflector profiles could be one such application, while the possibility of monitoring the health of such items as large cryogenic fuel tanks would be a typical longer-term objective. Some more specific examples are given below.
Figure 3. Optical fibres embedded in a woven CFRP skin (Photo courtesy of Westland,UK )
Figure 4. The effects of ageing on bismaleimide CFRP after 100 h at 250 degC (mag x 20). (Photo courtesy of Short Bros.,UK )
These are likely to be critical constituents of a smart structure. In any space applicatlon, they will have to require a very minimum of the spacecraft mass, volume and electrical power budgets. They will have to work for a minimum of five years, and probably much longer in some project applications. They will have to survive the launch environment and the longer-term effects of the in-orbit environment. Some typical sample problems that may have to be addressed are:
Clearly, the importance of some of these questions will be application-dependent, but some general feeling of confidence in the use of 'off-the-shelf' items has to be established before the introduction of optical fibres into composite space structures can be seriously entertained. Work of a similar nature is probably needed for actuators. Westland Aerospace (UK) are conducting practical investigations into the problems of and solutions for realising embedded optical fibres. Some additional work has been conducted for ESA by Dornier(D).
Current antenna reflectors in Europe rely almost entirely on the use of stiff polymer composite face sheet/honeycomb-sandwich core constructions, where the required surface profile has been realised during manufacture. For most present applications, this will suffice up to a diameter of about 3 m , but where larger diameters are required, the reflector must be stowed compactly during launch and subsequently unfurled once in orbit. This necessitates alternative forms of construction, of which mesh antennas are a typical example. American and Russian versions have been flown and similar technology has also been developed in Europe.
One of the main problems that has to be resolved for unfurlable antennas is the assurance that the necessary surface profile can be realised after deploying the reflector in the zero-gravity conditions of space. Clearly, if the shape can be 'tuned' by the use of sensors and actuators in orbit, this would reduce the complexity of ground testing and perhaps even simplify the method of construction. The challenge is to demonstrate a reliable, cost-effective system whilst not placing additional pressures on the resource budget identified for the 'passive' configuration. The shape could, of course, be realised using various thin-shell and membrane concepts, of which the mesh configuration is just one example.
A study being undertaken by ERA (UK) and Dornier (D) includes some initial assessments of 'smart' technology for antennas and will examine the potential of the competing transducer systems for shape control. A breadboard configuration using co-located piezo-electric transducers for strain control is being developed. In practice, it is difficult with most currently marketed transducers to realise efficient thermal strain control because of the high coefficient of thermal expansion of the transducer material.
The presence of micro-vibrations emanating from on-board sources such as reaction wheels, stepper motors, mechanical relays, etc., can disturb experiments requiring zero-gravity conditions, or can cause jitter that disturbs the imagery of optical payloads.
Considerable information has been gleaned from a micro-accelerometer package flown on the Olympus spacecraft for in-orbit health monitoring/vibration control. Work is continuing to develop smart systems that will employ automatic signal-recognition techniques and exploit both passive and active methods of vibration reduction. For example, many optical payloads employ isostatic mounts to avoid distortions due to quasi-static loads. These are of ten in the form of flexures. They also provide the sole path for in-orbit vibration transmission. Investigations are planned into the use of passive or active damping systems at these locations to eliminate the unwanted vibrations.
Achieving sufficient low-frequency noise reduction through a launcher payload fairing is a difficult task and liable to involve a substantial mass penalty. The use of active control systems employing embedded piezo-ceramic transducers is the subject of current investigations, which it is believed can lead to a more mass-efficient system .
In future air-breathing launchers, similar techniques could be used to reduce the risk of acoustic fatigue in the air intake.
The space industry remains conservative in its selection of materials. Preference has traditionally been given to proven materials with multiple sourcing. These criteria can be fulfilled by steels and aluminium alloys as the established materials of construction. Fibre-reinforced epoxies (e.g. CFRP) in laminate or sandwich-panel form are now in common use, after twenty years of developmentand progressive application.
Recent discontinuation of some industry-standard ultra-high-modulus CFRP prepregs has served to demonstrate the vulnerability of the space industry because it purchases relatively small quantities. With more ambitious space programmes, particularly those requiring high-temperature materials, there is a likelihood that only single sourcing will be available for a specific material or processing technology. It is already the case that the immediate commercial availability of some advanced materials is in considerable doubt; this includes thermoplastic/bismaleimide/polyimide composites, aluminium-lithium alloys, particulate and fibre-reinforced aluminium composites and titanium-matrix composites. There are not enough immediate applications to sustain present commitments in ceramic-matrix composites such as C-SiC, where initial funding froms paceplane programmes has now been scaled down.
There will be an increasing commercial tendency to consolidate on those materials and process technologies for which there are the largest markets. In this respect, future space programmes may be confined in their ambitions to using materials which are not necessarily ideal for the application, but are at least available and not overly expensive.
The potential for the application of smart structures and their related technologies is evident, but faces a range of interdisciplinary problems. It seems likely that 'partially smart' techniques may well be the first applications. As with more conventional structures, their evolution and introduction will depend on theircost-effectiveness and reliability. This applies to the constituent embedded sensors, actuators and the like, and their use of spacecraft resources. These will have to be traded against the capabilities of traditional methods.
Figure 5. Machined all-CFRP core for a highly dimensionally stable reflector. (Photo courtesy of Dornier, Germany)
Figure 6. C/SIC air-intake ramp which operates at 1200 degC (Photo courtesy of Dornier)
ESA Bulletin Nr. 79.