Amid spacecraft propulsion methods, solar sailing is often the forgotten alternative amongst the dominant systems such as chemical and electrical propulsion. An elegant form of traversing the cosmos, the sail is propelled by solar radiation rather than using any form of fuel, utilising Newton’s second law; the photons from the sun transfer their momentum to the reflective sail, providing acceleration and pushing the sail forward.
Figure 1: Solar Sail Model (adapted from W.A.Hollerman 2003)
Figure 1 shows a sail on a three dimensional axis at an arbitrary distance r from the sun moving along the +x axis due to the force from solar photons.
The concept of solar sailing has been in circulation since the 1970s (and early ideas since 1700s) and there has been an ever increasing amount of significant research done in this field. There have been a number of missions to test the viability of solar sailing; the first successful demonstration of using solar sail as the main propulsion system was JAXA’s IKAROS spacecraft launched in 2010, NASA’s test of solar sail deployment with the launch of NanoSail-D2 also in 2010, the Lightsail project designed by The Planetary Society to analyse sail performance; Lightsail 1 launched in 2015 and Lightsail 2 to launch in 2017. Solar sail spacecraft encounter the same challenges as current spacecraft while in the space environment such as thermal control, solar radiation damage, and micro-meteorite impact, but there are unique challenges that arise due to the use of a large sail. Sail attitude control, sail deployment and structure, and sail materials are just a few areas where complications could appear. The main reason for interest in solar sail propulsion is the constant acceleration the spacecraft can gain allowing it to potentially achieve extremely high speeds near 0.05c (Matloff & Mallove 1981), this has huge implications for future missions. Optimization for solar sail type missions will possibly be simpler than other types since no propellant has to be carried improving the weight and cost benefit., however this will be expanded upon later in the review. The following pages will review these solar sail specific challenges to systems and optimization and also the advantages that this propulsion system possesses.
2. Overview of the systems of Solar Sails
A common misconception is that the solar wind propels the solar sail, however this is not the case. Rather, the solar sail is propelled by sunlight, that is by photons and not the charged plasma of the solar wind. The Sun outputs constant sunlight and there will always be a pressure force on the sails due to sunlight, according to the inverse square law the intensity is inversely proportional to the square of the distance of the craft. Solar radiation can be used for both orbital manoeuvres and attitude control. The reflective surface must be very large in order to maximize the minute momentum carried by each single photon, Kapton and Mylar are the most common materials used to construct the sails due to their thermal stability, tensile strength and chemical nature (McInnes 2004). As the review will later cover, there are several sail configuration types the main ones being heliogyro and square. A project in the 1970s carried out by NASA JPL (Friedman 1976) assessed the feasibility of a solar sail mission to Halley’s comet, one of the research areas of the project culminated in the design of a heliogyro sail configuration in addition to analysis of a standard square sail. JAXA’s IKAROS mission is currently using a square sail and has been successful in demonstrating solar radiation thrust (Tsuda et al 2011).
Attitude control is another major system that will be addressed, alternatively to using propellant based thrusters or RCS, a solar sail spacecraft will be able to utilise unconventional attitude control methods with its sails and using solar radiation. The heliogyro has a distinctive attitude control system, using cyclic pitching of its six blades “to generate torques about its spin axis” (Wilkie et al 2013, p.2) to create individual direction thrust components. Other types of attitude control include Reflective Control Device, Control vane method and the sliding masses method. The chosen attitude control method will depend on the design of the sail itself, each having their own benefits for their respective configuration.
2.1. Heliogyro design
2.1.1. Sail configuration, structure and deployment
Early NASA research on Heliogyro design was started by MacNeal in 1967 (MacNeal 1967) and further expanded upon during the Halley comet rendezvous mission proposal in 1977. The first report and study was produced by MacNeal for analysis of an alternative design to the square sail. His design consists of a two-blade system with an area of 180,000 ft2, length of 326 ft for a single blade and a chord of 4.84 ft, this design can be altered but this example was a suitable compromise between the performance of the sail and the weight of the spacecraft. A key observation for general solar design was that in order to overcome dynamic pressure due to atmosphere the minimum operating altitude should be 800km (Gordon 1961).
Folding is an unwanted consequence in sail deployment as it creates wrinkles, reducing the effective surface area and hence efficiency. The heliogyro can avoid this consequence by unrolling the long rectangular sail material curled on spools in the central body. However, this could have a possible disadvantage because of larger space need to store the spools, any extra dimensions will result in increased structural weight. Centrifugal force is the most suitable method for rigidizing the sail and can be applied for all sail types. Other suggested methods are compression resistant and pneumatically stabilised structural members, also electrostatic or magnetostatic forces to create tension in the structure. The latter two methods require a magnitude of voltage and current too high to be practicable. Three important deformation types that were studied by MacNeal were the “vertical (or flapwise) deflection, in-plane (or chordwise) deflection and twist” (MacNeal 1967 p.10). Coning angle, , expresses vertical deflection and small values are wanted. The coning angles is proportional to the radius of the radius of the blade configuration, if the angle is large then the deviation between the root and the tip is also large and this will lead to chordwise deformation. MacNeal calculated the torque required to pitch the blades, with the value of .002o in-lbs being relatively very small. This was due to two effects found in the equation, the most important being that if the blades were oscillated at one cycle per revolution while in pitch, the inertia moment would be cancelled out by the restoring moment. Bending stiffness was found to be an issue but this depends on the material chosen and is only important when complemented with the analysis of deformation.
The last area of focus is deployment, as discussed earlier the blades are deployed using centrifugal force. A rocket motor is used to provide a small amount of force to rotate the spacecraft starting the process, at certain length after unrolling, the blades are pitched collectively and photon pressure is used to complete the deployment. Formulas can be used to calculate the deployment phase and the time required for the example was 71.2s, this value accounts for zero trailing edge compression of the sail blades. MacNeal concludes by suggesting that long-duration missions are the best scenario for the Heliogyro including mission such as station-keeping interplanetary manoeuvres. The performance criteria will have definitely improved since MacNeal but this data on the Heliogyro is still relevant. Further study by NASA has been carried out and smaller scale missions have been made possible by the advances of lightweight satellites like the CubeSat (Wilkie et al 2013).
2.1.2. Attitude control
Since the Heliogyro has a similar configuration to helicopter blades, MacNeal proposes to use helicopter controls to generate the control responses of the solar sail. The combined pitch of all the blades will produce a rolling moment on the spacecraft, the spin-up time for MacNeal’s example vehicle was 3.34 days, using ideal illumination angle the precessional manoeuvre rate was .0143 per day. This value requires very little correction for a long voyage, large variations in spin rate must be minimized by “periodically reversing the signs of the cyclic and collective pitch angles during the precessional maneuver” (MacNeal 1967 p.21). This is the main attitude control method for Heliogyro configuration and it allows the spacecraft to adjust its direction quickly, which would be efficient for orbit changes (Wilkie et al 2013).
2.2. Square sail design
2.2.1. Sail configuration, structure and deployment
The square design is one of the simplest designs along with the disc sail. The structure is simply a square sheet with the corners attached to a boom ends which converge into the centre mast (Daniels et al 1977). The IKAROS mission has physically shown the capability of this design and has gathered in-orbit data. A mission to Mars was also designed by Eastridge using a square sail design. Eastridge chose the square sail design mainly due to its simplicity, not only in design but also in deployment (Eastridge et al 1989). The square sail also has a high area to mass ratio therefore to the benefit of maximizing velocity. A free sail with no booms has been considered but is not durable enough for long duration missions. Most square sails will use a boom or bridge structure with trapezoidal or triangular sections constructing a square. IKAROS has this exact structure; four trapezoidal sail sections connected by bridges, further strengthened by eight tether lines connecting the sail to the central hub and four flexible harnesses (Tsuda et al 2011). Additional strengthening techniques could include reinforcing the sail by increasing material thickness towards the edges or using stiffened material along the edge tendons to distribute load across the sheet more evenly (Daniels et al 1977). The sail will have to have a cutout in the centre to allow space for the mast and harnesses.
Similar to the heliogyro, the most effective method for deployment is using centrifugal force. Other methods such as extendable booms are less practical as they require added mechanisms to operate which increases weight. The deployment of IKAROS was carried out in two stages; the first stage, also known as the quasi-static phase, is release the tip masses to extend the bridges into a cross shape, the second stage, dynamic phase, releases the guide rollers and the sail unfolds by centrifugal force. The in-orbit deployment was carried out over a period of 29 days. On May 21 2010, it launched with an initial spin rate of 5 RPM which was reduced to 2 RPM over 4 days. The tip masses were released at this speed after which the spin rate was increased again to 25 RPM. First stage deployment took 7 days to complete until June 9 when the second stage deployment was initiated. The second stage deployment is very quick, only taking a few seconds. The spin rate is finally decreased to 1 RPM (Tsuda et al 2011). It is important to dampen the spin rate after deployment in order to reduce stress and strain on the structure, the attitude control system can reduce the spin rate of the sail. Square sails are relatively simple in design and can be constructed in a number of ways to suit a mission, this is a major advantage of a square configuration.
2.2.2. Attitude control for general square sail design
RCS is the standard system on attitude control on most spacecraft, it can be used normally for a solar sail spacecraft but there are fuel-less attitude control methods which can only be used for low-thrust systems such as a solar sail. RCD, Reflectance Control Device, is the system used by IKAROS. These are reflective liquid crystal sheets embedded within the sail that can change its optical reflectance when a voltage is applied, synchronising the operation of these sheets allows the attitude to be controlled. IKAROS aims to flyby Venus and so was orientated towards the sun, the change rate data recorded in-orbit was precisely the same as the design objectives, thus RCD is a viable system for attitude control (Tsuda et al 2011).
IKAROS in-orbit attitude control data (Tsuda et al 2011)
Fu has reviewed several methods for rigid sail attitude control (Fu et al 2016). One method is control vanes; this method is a very similar concept as the RCD system used by IKAROS. Control vanes are sheets of material at the corners of the sail which can be rotated on two axes, they are made out of the same material as the main sail. Four control vanes equals eight degrees of freedom, more can be added for greater control. The main disadvantage is that if some vanes are shadowed, it will affect the generation of torque of the shadowed vanes (Kun 2015). The shifted wings method involves shifting different sections of the sail to create a difference between the centre of mass and the centre of pressure, the shifts occur in the same plane so there would be no effect on the spin torque. The tilted wings method is similar but instead of shifting the sail sections it rotates them by rotating the tips attached to the corners of the sail, it can be used in conjunction with the shifted wings method (Wie 2004). Using independent masses can offset the centre of mass and change the direction of thrust components. The gimballed mass method has a mass offset from the sail attached by a rod, it creates a body torque due to the offset of the centre of mass. Adjusting the centre of mass to the centre of pressure will create body torque which will control the attitude. The final method is the Sliding masses method; masses have one degree of freedom along the booms and slide to change the position of the centre of mass. Fu believes that the sliding masses method has “great potential for use in future solar-sail missions.” (Fu et al 2016). The methods explained by Fu et al have not been tested like the RCD and so would need to be tested in an in-orbit environment before being considered as a viable attitude control method.
2.3. Sail material
2.3.1. Solar system plasma environment effect on solar sails
The plasma environment is the predominant threat to the performance of the solar sail. Douglas analysed the effects of the plasma environment on a solar sail system. A major source of damaging plasma arises from the solar wind, it can cause electrostatic charging of the sail which can lead to unstable oscillations and interference of electrical systems (Douglas et al 1977). To understand the solar wind data must be collected on this plasma, theoretical models exist at 1 A.U, this data can be extrapolated to 3 A.U. Most of the calculations done by Douglas et al will use the data within this range.
After calculating parameters of the properties of the solar wind, the effects of sail charging can be analysed. The effects of the solar wind are more significant for the square sail than the heliogyro, the blades of the heliogyro are close to parallel to the solar wind. Charging of the solar sail can be reduced by electrically shorting the front and rear side of the sail. A plasma wake can form behind the spacecraft, calculating the behaviour of the disturbed plasma behind the sail is complicated consisting of relating the Poisson and Vlasov equations, hence the wake will be analysed physically. Electron flux in the plasma was not found to have significant effects due to the low electron density, using an expression for the “Maxwellian components for electron distribution at infinity” (Douglas et al 1977 p.20). Ion flux is another component of the solar wind and was estimated to be zero at the rear of the sail because the solar wind will impact the sail at the front therefore few protons will hit the rear of the sail. For a shorted sail, most plasma current effects such as plasma proton flux are not influential with the exception of secondary emission. The threshold for secondary emission processes occurs around 40-50eV incident electron energy, thus in low solar winds (Te~20eV) secondary emission is negligible, as the solar wind intensifies up to Te~100eV the effects become more prominent. This can lead to the degradation of the oxide layer of a metal oxide, it is better to use smaller pure metal yields to minimize this effect. Another area of research was carried out into discovering if unstable modes of oscillation is caused by solar wind plasma. Although further research is needed in order to find the conclusion, Douglas has reviewed evidence and assumes that there is no substantial influence of the solar wind plasma on sail oscillation (Douglas et al 1977). In conclusion, the plasma environment was found to have no considerable damaging effect on the operation of a solar sail mission. Induction and oscillations caused by the solar wind were both found to be much smaller than original analysis had first determined.
2.3.2. Sail material in design
Solar sail material is the paramount design parameter when designing a solar sail. It will determine the potential thrust of the system and the capability of the spacecraft’s movement. The material will have to be lightweight and durable enough for the solar environment. Rowe and JPL at NASA extensively researched the materials suitable for solar sail, in particular they were evaluating the suitability of materials for the proposed Halley comet rendezvous (Rowe 1978). The original target set for the material was “an ultra light, highly reflecting material system capable of operating at high solar intensity (high temperature, high radiation dose) for long periods of time” (Rowe 1978 p.1). At the time of this research, a target area of 4.5 gm/m2 for the system was set and was required to operate for more than one year at an altitude of 0.3 AU with an equilibrium temperature of 370c. Successful testing was achieved by July 1977, feasibility was proven for an even lower density, 3.3 gm/m2, than the original target operating at 0.25 AU.
Any material that can be produced as a film could be used as sail material but after brief consideration, polymers emerged as the only viable type as they could satisfy the thermal and radiation requirements. Aluminium was considered but to reach the requirements it had to be 0.04mm thickness to meet the target area density, at this thickness its properties degenerate.
Screening of polymer candidates narrowed the choice to two materials, Kapton and B100. These two choices were far ahead in terms of performance than several polymers such as Thermo-plastic polymide P100. Comparing data between Kapton and B100, Kapton has many more superior properties, a few examples are; Kapton has a service temperature range of -270 to 400c compared to -270 to 300c for B100, Kapton also has a much higher tensile strength of 17500 psi at 25c as to 7770 psi at 25c for B100.
Kapton itself is not specular enough to reach the requisite optical parameters, the illuminated side needs to be highly reflective to maximize thrust performance, thus the material needs to be coated. The coat not only improves reflectance but also provides protection from UV radiation and charged plasma from the solar wind. Aluminium coating has properties which are desired and also economically viable, silver was also highlighted as a potential coating as it had higher overall reflectance but has a high cost and a high density three times that of aluminium. During the period of Rowe’s research there was no substantial data on the impact of radiation on the sail film however it was estimated that an aluminium coating thickness of 1100-1200Å was sufficient to protect the Kapton film from UV radiation. A protective overcoat of the sail is the final component of the sail design; the overcoat must not hinder the performance of the sail but also aim to increase emittance of the sail. Several methods evaluated to increase emittance included introducing an absorbing layer or carbon coating the film. The most viable method was either to add a layer of high resistivity metal or to use an oxide overcoat (Rowe 1978). In conclusion, Kapton was found to be the best base material, an aluminium coating proved to be an effective coating material and increased emittance. Chromium at thickness of 125Å was the preferred choice of overcoat for the backside of the sail due to the increase of emittance it provided.
3. Overview of trajectories and missions
3.1. Review of past and planned missions
There have been several solar sail missions over the past twenty years and many more planned for the future. There will only be a review on the higher profile missions including IKAROS, NanoSail-D2 and Sunjammer. Their mission structures will be briefly summarised and the data gathered in-orbit will be evaluated.
3.1.1. NanoSail-D2
NanoSail-D was a solar sail project designed to deploy from a CubeSat in 2008 to demonstrate and observe sail deployment and the capability for de-orbit (Johnson et al 2010). The original mission was to launch in 2008 but the launch vehicle failed while ascending. NanoSail-D2 was the second chance to launch in 2011, successfully deploying a CP1, 10m2 sail in approximately 5 seconds. It is the first solar sail to orbit earth and transmitted data over a period of 3 days. Measurements on the rotation rate of the spacecraft revealed that gravity had a larger effect on the torque generated than solar radiation pressure, “Gravity caused unstable rotation in the roll axis for 0.1 days, then regular oscillations resumed and the overall rate of rotation became negative.” (Katan 2012 p.6). For a simulation with only solar torque applied, the solar torque actually rapidly stabilized the sail and generally does not cause instability during full sun exposure. A study on the effect of the position of the centre of pressure was conducted to establish a centre of pressure offset range of the sail plane. It was found to have a very large effect on the period of stabilization, at sample offset of 0.001m it took 0.36 days to stabilize, at an offset of 0.005m the sail stabilized in 42.16 days.
Katan concluded the review of the NanoSail-D2 by stating that the ideal orbit for a solar sail in a Low Earth Orbit could be a polar sun-synchronous orbit because of the stability of the sail in full sun exposure. Centre of pressure offset should not be too large as to prohibit sail stabilization, the maximum for NanoSail-D2 was 0.005m, sails in LEO need rigid structures to minimize flexibility in the centre of pressure. NanoSail proved the viability of using sails as a de-orbit device for space debris, the sail had a high drag are which lead the satellite to de-orbit within 9 months (Katan 2012).
3.1.2. IKAROS
IKAROS was launched with JAXA’s AKATSUKI as a piggy-back payload, it is the first functional interplanetary solar sail spacecraft. This mission was able to demonstrate many effects and performance results, it also validated new technologies such as RCD attitude control and solar power generation by solar cells on the sail. IKAROS used a sheet of Polyimide for its sail. As seen in section 2.2.1 and 2.2.2, RCD was proven to be an effective and viable method for attitude control. After deployment the sail reduces its angular velocity and this process causes oscillations in the system, the structural damping showed its effectiveness at providing stability for the normal spin axis (MacDonald 2014).
Communication was lost with the mission in December 2011 due to insufficient solar power generation, JAXA needed to correctly predict the attitude and orbit in order to re-establish communications the spacecraft (Mimasu et al 2014). They were able to model attitude changes using a logarithmic relationship and the attitude drift. Using table, the attitude drift motion can be interpreted after finding the torque parameters, A, B and C and equations that Mimasu has derived with data recorded during the transit to Venus. Fuel for the RCS was fully depleted and attitude control was purely RCD.
Orbit predictions were based the attitude profile found previously using table. To improve accuracy of the predictions, the rear side sail dynamics were calculated using residual data from Doppler radar and optical parameters were estimated. This orbit prediction can be used for a sail with RCD as its main attitude control system. IKAROS is currently still in flight and JAXA continues to monitor its progress and alter its course. JAXA’s future solar power missions to Jupiter and the Trojan asteroids will use and build upon the research of IKAROS.
3.1.3. Sunjammer
L’Garde Inc designed a solar sail mission called Sunjammer to travel to and maintain a sub-L1 point, to investigate station-keeping potential of a solar sail (Barnes et al 2014). It was planned to launch in 2015 but was cancelled several months before, however the study behind the mission is still relevant in planning interplanetary solar sail trajectories. It was to launch onboard a Falcon 9 to Geostationary Transfer Orbit, detach from its carrier and travel to its sub-L1 within 30 days. Sunjammer deploys a 1200m2 Kapton sail using inflatable booms and controls its attitude using the control vane method.
The proposed trajectory would aim for a sub-L1 Artificial Equilibrium Point, starting from a a ballistic orbit to a solar sail phase and then finally reaching the sub-L1 point or a Halo orbit in the region. Heiligers analysed the mission using the restricted three body problem and assumed that the launch starts from a midnight geostationary transfer orbit. The transfer modle used by Heiligers is composed of two main phase, a two-body ballistic phase and an interplanetary solar sail phase. Optimization of the orbits were solved with PSOPT, a pseudospectral method in C++.
Orbit trajectories between the targets of reaching the sub-L1 point and the region Halo orbit were compared. It required less V for a real sail model to reach Halo orbit than to target the sub-L1 point and was considerably quicker, however for an ideal sail model the V for sub-L1 point was lower but still much slower than a trajectory to the Halo orbit (Heiligers et al 2014).
3.2. Optimal orbits
3.2.1. Inner solar system
Sackett analysed the optimal planetocentric trajectories in 1977 just after the proposed Halley Comet Rendezvous mission. A computer program called SUNSPOT was used to optimise trajectories with limited performance and specifically missions to the inner solar system. Due to the low levels of thrust acceleration, a solar sail orbit would require a spiral phase to escape planetary sphere of influence, the report by Sackett focuses on this spiral phase of the escape trajectory (Sackett 1977). SUNSPOT was built upon an earlier program for trajectories for electric propulsion, it uses the method of averaging and equinoctial orbital elements. A shooting method is used to optimize overall trajectory by integrating state and costate from initial time to the final time.
Three methods are used within the computer program the first of which is Averaging, which uses an averaged Hamiltonian to shorten computing time. The second is the solution of the two-point boundary value problem, once solved it will generate a locally optimal extremal trajectory. The third is to use numerical methods, i.e. a Newton-Raphson iterator, to calculate the sensitivity matrix by running similar trajectories over time each with different initial values.
The model used in the code was a square sail, the model used was not idealized and its forces were modelled before optimizing its orbit. A limitation to the code is that the escape trajectory cannot be fully defined therefore trajectories to sub-escape conditions are considered instead. Sackett has two assumptions for an orbit-to-orbit transfer; the initial orbit is given and the final orbit is specified or its orbital elements are specified. Averaging the trajectories yields a two-point boundary value problem which is solved by an iterative method. For a more accurate model, complexity can be increased by including the effects of shadowing, oblateness and a penality function. A pericentre penalty function is added because a planetocentric solar sail trajectories naturally develops large eccentricity quickly, optimization may lead to trajectories that actually intersect the planet’s surface. For the case of oblateness, analysis was focused on earth-centred trajectories. The sail has no thrust when it is shadowed, therefore when entering a planets shadow the entry and exit angles into the shadow need to be calculated in order to perform averaging.
SUNSPOT produced trajectories for earth-centric missions and demonstrated its performance by providing sensible flight times for escape and orbit transfer. The program is capable of optimizing inner plant trajectories but due to declining interest at the time extensive study was not carried out. This program provides a framework if further study was undertaken to explore interplanetary missions, many performance criteria has vastly improved and programs can calculate trajectories rapidly.
3.2.2. Outer solar system
Solar sail missions towards the outer planets is a vastly different scenario than inner solar system missions because of the square decrease is solar radiation pressure as an object moves away from the sun. “For such missions, the solar sail may gain a large amount of energy by first making one or more close approaches to the sun, thereby performing a so-called single or multiple ’solar photonic assist’ (SPA) maneuver, turning the trajectory into a hyperbolic one.” (Dachwald 2004 p.1). The minimum solar distance and the sail lightness dictate the flight time of the mission, both of which depend on the maximum sail temperature. Hence the main constraint for optimization is the temperature of the sail. The method used by Dachwald in his report is Evolutionary Neurocontrol, it is utilised in a program called InTrance.
InTrance was developed by Dachwald to improve upon the local trajectory optimisation methods. It combines evolutionary algorithms and artificial neural networks into ‘Evolutionary Neurocontrollers’ which InTrance uses with its artificial intelligence to optimise trajectories. The input neurons receive current and target parameters which produce direction unit vectors. Further detail is not necessary as the paper focuses more upon the algorithm and the variables that the code uses (Dachwald 2004).
Dachwald used InTrance to model a Neptune flyby to assess ideal vs non-ideal force models, compare optimization constraints and evaluate general features of SPA trajectories. Dachwald made a number of simplifications for feasibility such as assuming invariable optical properties of the sail film over the course of the mission. After optimising trajectories with different characteristic accelerations, the comparison showed that a higher number of SPAs are required as the characteristic acceleration decreases and also for heavier sails. Evaluating ideal reflectivity vs non-ideal, the travel times for non-ideal were 5% longer. Calculated minimum flight times to the other outer planets were very reasonable, near-term solar sails, characteristic acceleration 0.4 mm/s2, are able to reach Neptune in under 10 years and medium-term sails can reach the Edgeworth-Kuiper belt within 10 years. One important factor to consider is that these scenarios are strictly flyby missions because of the sizable hyperbolic excess velocities reached, orbit-to-orbit transfer will require a braking system or a slow down manoeuvre. Dachwald was able to show that even with mid-performance sails, quick flight times were achievable but it depends on the sail material and its ability to withstand high temperatures as an SPA is carried out.
4. Conclusions
As this review has covered, a solar sail spacecraft can be tailored in a variety of ways for specific mission. Sail configuration defines the look and attitude performance of the spacecraft, considerable research has concluded that the two main designs are the heliogyro and the square sail. The heliogyro has shown important potential but the square sail is the design which most missions have used and as such has been tested in space whereas the heliogyro has not.
Sail material characterises the thrust and lifespan of the spacecraft. It must be chosen to withstand extreme radiation levels and be durable for a long period of time. Polyimides are the prime candidate for sail material and the most economically viable. Kapton is a popular choice as it has been proven to operate effectively however IKAROS and NanoSail-D2 have used variants of polyimides i.e. ISAS-TPH and CP1 respectively. Regardless of the type of polyimide used, they are all coated with aluminium to increase reflectivity of the sail. For a future mission, it would be logical to use a square sail configuration made from Kapton or a similar polyimide.
Solar sail missions have been reasonably successful, unsuccessful missions have been mainly due to external reasons such as failure of launch. There have been a wide range of objectives for these missions from station keeping to interplanetary travel and have proven the capability of solar sails. Optimising a mission will require a computer program, there are existing codes which can be used such as Sunspot and InTrance, or a new code could be created based on these and averaging methods. An interplanetary mission is definitely viable using current technology and methods but long term effects on the sail have yet to be measured, creating opportunities for further investigation.