How my Space Journey Began : Designing a Conceptual Spacecraft

How my Space Journey Began: Designing a Conceptual Spacecraft – BY B.KANESH

My first encounter with space technology goes back to my polytechnic years (I was about 19).  Being one of the former space challenge 2010 winners and also the pioneer group of students to participate in the space domain. My team and I had to design a conceptual SSTO (Single stage to orbit) spacecraft highlighting several of the key systems to operate the craft. To date, its still one of the best teams I have worked with. The great enthusiasm the team members had despite the topics being beyond their domain of expertise.  Two models were built. One by hand, and the other by SLS Additive manufacturing. Almost 72 hours of building time was spent to display a prototype model at several exhibitions. The idea was to design a spacecraft that would travel to the low earth orbit and back while landing on the runway like a normal aircraft. Our team won 2nd runners up at SSC2010 and also a 2nd prize at the International Space Exploration and innovation contest in China.

Disclaimer: This article is to share the design that was derived at the point of the competition. The article should not be used as a technical guide. I am writing this article to help students further explore spacecraft design and the depict the thought process the team had in 2010. Having an idea of where to start would allow you to create much greater deliverables that what I have presented in this article. 


Conceptual Design Approach that was used

  • Functional Requirements of Spacecraft
  • Functional Requirements of an Aircraft
  • Design Elements
  • Permutations of Design
  • Selection of Design
  • Modelling
  • Aerodynamics and Testing (Not Included)
  • Modifications to Design
  • Prototype / Simulation

Functional Requirements of Spacecraft

Rendered Graphic of SSTO-SX4 Design
Rendered Graphic of SSTO-SX4 Design on Re-Entry
SX4 Payload Capsule
SX4 Experimentation/Debris Cleaning/Satellite deployment Payload
SX4 Passenger Transport Payload
SX4 – Payload Bay, with Doors Open
“Team Astronuts” with Hand Built Prototype Model
X Planes Simulation Modelling
Exhibition on Engineers Day@SCAPE


The only SSTO spacecraft we were exposed to at that time was the EADS Astrium Suborbital Space Plane. However, we had very little information on the design. The SX-4 Meteor design had 2 Pratt & Whitney CECE( Common Extensible Cryogenic Engines) Rocket Engines that either used a hydrogen-oxygen or methane-oxygen fuel combination. The air-breathing subsonic Turbofan engine option was to have a Honeywell aerospace TFE731-20. The turbofan has a sea-level thrust of 4400lb(2000kg) of thrust and 870lb of thrust at a maximum altitude of FL40(Most common height limit for commercial engines are around Flight level 32000ft-40000ft. The SX-4 will execute a stall maneuver while two (CECE) are fired, a new deep-throttling, 30,000-pound(13600kg) thrust for the 2 engines is used to propel the aircraft into the low earth orbit for a temporary period of time.

I’m still planning on calculating the escape velocity to reach the 100km mark for this configuration since I am much wiser now. Will keep this section updated on that soon.



CECE Engine Specifications

SX-4 Meteor Spacecraft Specifications

Aircraft Length :
Wing Span :
Thrust to Weight:
Wing Area:
Wing Loading:
Aspect Ratio:
Taper Ratio:
MAX  Takeoff  Weight
MAX Passengers
MAX Satellite Payload
200KG w/pax
SPS System Weight

Design Elements

Sears Hack Body At subsonic speeds. the pressure drag arising from the thickness of the body or wings is negligible so long as the shapes are sufficiently well streamlined to avoid flow separation. The sears hack design was chosen as the ideal body geometry for our flight profile which requires the spacecraft to go supersonic during rocket engine flight.

Read more (pdf) Sear-Hack Body

Chines – Commonly seen on the SR71 Blackbird. Sharp edges leading aft on either side of the nose and along the sides of the fuselage. The aerodynamicists discovered that the Chines generated powerful vortices around themselves, generating much additional lift near the front of the aircraft, leading to surprising improvements in aerodynamic performance. The angle of incidence of the delta wings could then be reduced, allowing for greater stability and less high-speed drag, and more weight (fuel) could be carried, allowing for greater range. Landing speeds were also reduced, since the Chines vortices created turbulent flow over the wings at high angles of attack, making it harder for the wings to stall and also creating the ground effect. The Chines act like the leading edge extensions that increase the agility of modern fighters such as the F-5, F-16, F/A-18, MiG-29, and Su-27. The addition of Chines also allowed designers to drop the planned canard for planes. Chines remains an important design feature of many of the newest stealth UAVs, such as the Dark Star, Bird of Prey, X-45 and X-47, since they allow for tail-less stability as well as for stealth. This feature also provides yaw stability to the aircraft even in supersonic speeds.

Jouster/Aerospike– A drag-reducing aerospike is a device used to reduce the forebody pressure drag of blunt bodies at supersonic speeds. The aerospike creates a detached shock ahead of the body. Between the shock and the forebody, a zone of recirculating flow occurs which acts like a more streamlined fore body profile, reducing the drag.

Shock waves develop around aircraft as they near Mach 1. At ground level, these are perceived as a loud double boom or bang. Their intensity varies due to factors such as weather, refraction from different layers of atmospheric density, and size of the aircraft, but in general, from a supersonic aircraft of the size of a civilian airliner, the overpressure created at ground level is enough to at least rattle windows, or in more extreme instances, cause structural damage to buildings.

For example, the sonic boom from the Concorde travelling at a speed of Mach 2 was about 2 pounds per square foot. Because of sonic boom intensity, many countries now prohibit supersonic overflight over land or population centres. The FAA prohibits supersonic flight over land, except in special military flight corridors. The Quiet Spike is a key enabling technology but alone is not enough to reduce sonic booms sufficiently to lift the current prohibition on supersonic overflight. The noise problem with today’s supersonic aircraft is caused by an N-wave pressure change that occurs as shock waves move past the observer. When an aircraft like the Concorde flies at supersonic speeds, conical shocks form at the front and rear of the aircraft and at every discontinuity in the flow along the way (for example, at the engine inlets, wing leading edges, and antennas), because the plane is moving faster than the air can move out of the way. At a certain distance from the aircraft, theIndividual shocks coalesce into two stable conical waves, one at the front of the plane and one at the tail. The effect on the listener is a double boom as each wave passes. According to Coen, shock waves with greater than ambient pressure tend to move faster and collect at the nose, while that with lower than ambient pressure lag to the back. On the ground, the pressure distribution as the waves pass looks like an “N”: As the nose shock passes, the pressure spikes above ambient pressure, ramps down to below ambient, then snaps back to ambient as the tail shock passes. For the Concorde, the initial overpressure (the change from ambient pressure as the front shock passes) is about 2.2 lb/ft2 (or psf) above ambient, while an unmodified F-5E is about 1.3psf. To the ear, a 1-psf pressure change is equivalent to the differential experienced in going down only one flight of stairs—but the effect comes across as a “boom” because it occurs in a fraction of a second. To decrease the effects on the ear, aircraft designers can either reduce the overpressure magnitude or modify the time it takes to reach the peak pressure, in part by preventing the shock waves from coalescing at the front and rear of the aircraft. The overpressure is also linked to aircraft weight, which helps to explain the Concorde’s elevated value compared to that of the F-5E. Psychoacoustics experts attempt to predict how humans will accept or reject noise based not only on sound levels but on how, for example, a boom might rattle the windows. These experts predict that 90% of theU.S. population would not be annoyed if boom levels could be limited to a maximum overpressure of between 0.3 and 0.6 psf, according to Coen, who adds that such levels could be as obscure as distant thunder or a car door slamming. From a regulatory standpoint, the FAA is accelerating its effort to see if existing noiseMetrics will be usable for testing supersonic boom aircraft and to determine how to rank annoyance for low-boom waveforms. The agency is teamed with NASA, Transport Canada, Penn State University, Stanford University, Purdue University, Boeing, Gulfstream, Wyle Labs, Lockheed Martin, and Cessna to study the issue.

On the SX-4, The jouster through computer technology will adjust itself based on the temperature and pressure, factors that determine the speed of sound and the sonic boom cone shape. The computer will deploy the jouster right before entering the sonic boom. Retract when the supersonic flight is complete, to avoid structural weakness in tensile strength. This dynamically changing length further improves the dynamic capabilities of the SX-4.

Meso-Flaps Approaching supersonic speeds, the boundary layer interaction between the plane surface and the airflow begin exerting great forces on the airframe. This results in increased drag and flow separation and inhibits the full aerodynamic efficiency of conventional supersonic aircraft. The system will get rid of the costly, complex and heavy bleed systems of the current supersonic air-crafts. Use of Aero elastics such as (NiTinol) will eventually adjust itself to optimize its effects on reducing boundary layer intensification.

Blended Design Body The Blended Design learned from Boeing X-48 would allow significant payload advantages in strategic airlift/air freight. Moreover, fuel efficiency will increase.  The design factor would also help reduce drag compared to conventional designs. The blended wing design would also increase the overall lift of the aircraft. The modelling of the blending design at that age was a nightmare. The model would have required an immense amount of curve definitions to achieve the blended body. The student edition of the CAD software could only do so much.

Chevrons Chevrons technology can be seen on 787 Aircraft. Chevrons will reduce the noise produced by the engines. This means it will be much quieter as the aircraft takes off the airport in any location.


The air inlets of the turbofan engine will be upgraded to the spike system. The spike allows the engine to experience speeds up to Mach 3.3(proven). While the intake of air is still subsonic at Mach 0.5 At the front of each inlet was a sharp, pointed movable cone called a “spike” that was locked in the full forward position on the ground or when in subsonic flight. During acceleration to high-speed cruise, the spike would unlock at Mach 1.6 and then begin a mechanical (internal jackscrew powered) travel to the rear.

Air Spike Operation – SR71

Wing Design

The Suitable characteristics after extensive research have shown that The low-sweep laminar design uses thin supersonic sections that support a favourable gradient from leading edge to trailing edge, maintain small cross-flow, and are compatible with extensive NLF. Several application studies have suggested that such a concept has significant performance advantages for supersonic flight.

Our Aircraft Design has the canard configuration, while problematic for many subsonic designs, must be viewed differently for supersonic aircraft. A low aspect ratio highly swept canard is much less destabilizing than a higher aspect ratio surface that might be considered on a subsonic counterpart. An ideal longitudinal distribution of lift and area is often more easily accommodated by moving the wing aft, reducing control authority of a Conventional tail. The aft location of the large wing also reduces cabin intrusion. In some cases, when the maximum usable CL is limited by the ground angle or approach Attitude and when the use of high lift devices on the wing is limited (as may be the case for low aspect ratio, highly swept wings), the canard design can provide an increase in Usable take-off or landing CL.

The angle of attack set at α= 8.35

Sweep: 60°

The Area Rule (Aerodynamics) 

At high-subsonic flight speeds, supersonic airflow can develop in areas where the flow accelerates around the aircraft body and wings. The speed at which this occurs varies from aircraft to aircraft and is known as the critical Mach number. The resulting shock waves formed at these points of supersonic flow can bleed away a considerable amount of power, which is experienced by the aircraft as a sudden and very powerful form of drag, called wave drag. To reduce the number and power of these shock waves, an aerodynamic shape should change in cross-sectional area as smoothly as possible. This leads to a “perfect” aerodynamic shape known as the Sears-Haack body, roughly shaped like a cigar but pointed at both ends.

The area rule says that an aeroplane designed with the same cross-sectional area distribution in the longitudinal direction as the Sears-Haack body generates the same wave drag as this body, largely independent of the actual shape. As a result, aircraft have to be carefully arranged so that large volumes like wings are positioned at the widest area of the equivalent Sears-Haack body, and that the cockpit, tailplane, intakes and other “bumps” are spread out along the fuselage and or that the rest of the fuselage along these “bumps” is correspondingly thinned.

The area rule also holds true at speeds higher than the speed of sound, but in this case, the body arrangement is in respect to the Mach line for the design speed. For instance, at Mach 1.3 the angle of the Mach cone formed off the body of the aircraft will be at about μ = arcsin (1/M) = 50.3 deg (μ is the sweep angle of the Mach cone). In this case, the “perfect shape” is biased rearward, which is why aircraft designed for high-speed cruise tend to be arranged with the wings at the rear. A classic example of such a design is Concorde.

Sources: Research Ref. (Design of aircraft by Thomas C Corke) and aircraft performance and design by (John D.Anderson, JR).

Materials Technology 

Following the study of the shuttle orbiter aircraft. The materials technology to our SSTO Aircraft is very similar to its design. Thus we decided to study the materials found on the orbiter spacecraft very closely. Materials used for fabrication must withstand operational temperature requirements, loads, contamination, life expectancy and natural space environments. Properties to be considered include mechanical properties, fracture toughness, flammability, corrosion, stress corrosion, thermal and mechanical fatigue properties. Materials shall be selected to ensure maximum life and minimum maintenance. Materials which are not expected to meet design life requirements but must be used for functional reasons shall be identified as limited-life items requiring maintainability. We are presuming that by the time this aircraft is in production, there should be lighter and stronger options of materials.

Low Earth Orbit

Materials exposed in the Low Earth Orbit (LEO) environments shall be selected to perform in that environment for their intended life cycle exposure. The Critical properties of the material shall survive exposure to LEO environments of atomic oxygen, solar ultraviolet radiation, ionizing radiation, plasma, Vacuum, thermal cycling and contamination. Meteoroids and orbital debris shall also be considered.


As a spacecraft re-enters the earth’s atmosphere, it is travelling very much faster than the speed of sound. The aircraft is said to be hypersonic. Typical low earth orbit re-entry speeds are near 17,500 mph and the Mach number M is nearly twenty-five, M < 25. The chief characteristic of re-entry aerodynamics is that the temperature of the flow is so great that the chemical bonds of the diatomic molecules of the air are broken. The molecules break apart producing electrically charged plasma around the aircraft. The air density is very low because re-entry occurs many miles above the earth’s surface. Strong shock waves are generated on the lower surface of the spacecraft.

The four most important parts of the spacecraft will get heated while the re-entry of the spacecraft into the atmosphere. Those parts are the tip of the nose of the spacecraft and the leading edges of the wings. If ordinary ceramic tiles are used for these parts, they are bound to disintegrate as they are exposed to heat of 3000° Fahrenheit. Engineers then developed a non-parse composite RCC. It is distinguished by its colour, grey. They are much larger, stronger impact-resistant and able to tolerate much higher temperatures than ceramic tiles. They possess great mechanical properties such as the strength of carbon-carbon with unidirectional reinforcement fibres is up to 700MPa. Carbon-carbon materials retain their properties above 2000°C which is 3632° Fahrenheit. While RCC has the best heat protection characteristics, it is also much heavier than the silica tiles and FIB blankets, so it is limited to relatively small areas.       Weight: RCC: 1,986 kg/m³ (124 lb/ft³).LI-2200 tiles: 352 kg/m³ (22 lb/ft³) FRCI tiles: 192 kg/m³ (12 lb/ft³)LI-900 (black or white) tiles: 144 kg/m³ (9 lb/ft³)FIB blankets: 144 kg/m³ (9 lb/ft³)

Reinforced Carbon-Carbon

Reinforced Carbon-Carbon (RCC). The light grey material which withstands reentry temperatures up to 1510 °C (2750 °F) protects the wing leading edges and nose cap. Each of the Orbiters’ wings has 22 RCC panels. These panels are about 0.635 cm (0.25 inch) to 1.27 cm (0.5 inch) thick. T-seals between each panel allow thermal expansion and lateral movement between these panels and the Orbiter’s wing.

RCC is a laminated composite material made from graphite rayon cloth and impregnated with a phenolic resin. After curing at high temperature in an autoclave, the laminate is pyrolized to convert the resin to carbon. This is then impregnated with furfural alcohol in a vacuum chamber, then cured and pyrolized again to convert the furfural alcohol to carbon. This process is repeated three times until the desired carbon-carbon properties are achieved.

To provide oxidation resistance for reuse capability, the outer layers of the RCC are converted to silicon carbide. The silicon-carbide coating protects the carbon-carbon from oxidation. The RCC is highly resistant to fatigue loading that is experienced during ascent and entry. It is stronger than the tile and is used around the socket of the forward attach point of the Orbiter to the External Tank to accommodate the shock loads of the explosive bolt detonation. The RCC is the only TPS material that also serves as structural support for part of the orbiter’s aerodynamic shape: the wing leading edges and the nose cap. All other TPS components (tiles and blankets) are mounted onto structural materials that support them, mainly the aluminium frame and skin of the orbiter.

A brief explanation of the procedure: The large round panels of RCC are fit together. These panels are at a high cost exceeding $100,000 per panel. The main reason would be because there is a long heat treatment process allows the carbon to form into larger graphite crystals. TCLs made of RCC materials are used as gap fillers between the panels. Behind them are the insulation materials that are the final barriers, keeping the heat from reaching the Aluminium wing.

Silica tiles

This system is used to radiate or reject the heat away from the vehicle which is able to withstand heat up to 3000°. They are the first reusable material that can be used in space. Ceramic tiles made of silicon fibres derived from sand.

The steps of production are as follows: The silicon fibre is mixed with water and other materials and then shaped into soft blocks. Excess water is squeezed out and the mixture is then bounded in a heating process. The material consists of parse ceramic material comprised of mostly air voids of light-weight and heat resistant silicon-fibres. Individual tiles are then cut from large blocks and machined into specific sizes and shapes depending on which part of the shuttle. It is baked at a temperature of 2300° Fahrenheit. The block dissipates heat very efficiently such that it can be held seconds after being removed from the surface. They are then coated with black glass-based liquid. Black radiates heat and the coating keeps the plasma particles from sticking onto the tiles surface top and sides. The tiles can damage easily and can be cracked by applying pressure by finger. They are fragile. And estimated a shuttle comprises up to 20,000 tiles for one shuttle. Each tile is hand bonded by a rubber based adhesive.

Nomex Felt Coating

Nomex Felt Reusable Surface Insulation (FRSI). The white, flexible fabric offers protection at up to 371 °C (700 °F). FRSI covers the Orbiter’s wing upper surface, the upper payload bay doors, a portion of the OMS/RCS pods, and aft fuselage.

The preferred core material is aluminium, while aramid fibre paper (Nomex), which is quite common in aircraft construction, is not often used in space applications because of the environmental conditions.


Titanium can be alloyed with iron, aluminium, vanadium, molybdenum, among other elements, to produce strong lightweight alloys mainly for the aerospace (jet engines, missiles, spacecraft).

A metallic element, titanium is recognized for its high strength-to-weight ratio. It is a strong metal with low density that is quite ductile (especially in an oxygen-free environment), lustrous, and metallic-white in colour.  The relatively high melting point (over 1,649 °C or 3,000 °F) makes it useful as a refractory metal.

Properties: Commercial (99.2% pure) grades of titanium have an ultimate tensile strength of about 63,000 psi(434 MPa), equal to that of common, low-grade steel alloys, but are 45% lighter. Titanium is 60% denser than aluminium but more than twice as strong as the most commonly used 6061-T6 aluminium alloy. Certain titanium alloys (e.g., Beta C) achieve tensile strengths of over 200,000 psi (1,400 MPa). However, titanium loses strength when heated above 430 °C (806 °F).

It is fairly hard although not as hard as some grades of heat-treated steel, non-magnetic and a poor conductor of heat and electricity. Machining requires precautions, as the material will soften and gall if sharp tools and proper cooling methods are not used. Like those made from steel, titanium structures have fatigues limit which guarantees longevity in some applications.


Aluminium is the most abundant material in the Earth’s crust,  and the third most abundant element therein, after oxygen and silicon. It makes up about 8% by weight of the Earth’s solid surface. Aluminium is too reactive chemically to occur in nature as a free metal. Instead, it is found combined in over 270 different minerals. The chief source of aluminium is bauxite ore.

Properties: Aluminium is remarkable for its ability to resist corrosion due to the phenomenon of passivation and the metal’s low density. Structural components made from aluminium and its alloys are vital to the aerospace industry and very important in other areas of transportation and building. Its reactive nature makes it useful as a catalyst or additive in chemical mixtures, including ammonium nitrate explosives, to enhance blast power. Another important property of aluminium alloys is their sensitivity to heat. The low melting point of aluminium alloys has not precluded their use in rocketry; even for use in constructing combustion chambers where gases can reach 3500 K. The Agena upper stage engine used a regeneratively cooled aluminium design for some parts of the nozzle, including the thermally critical throat region.

Due to their high tensile strength to density ratio, high corrosion resistance, fatigue resistance, high crack resistance, and ability to withstand moderately high temperatures without creeping, titanium alloys are used in aircraft, armour plating, naval ships, spacecraft, and missiles. For these applications titanium alloyed with aluminium, vanadium, and other elements is used for a variety of components including critical structural parts, fire walls, landing gear, exhaust ducts (helicopters), and hydraulic systems. In fact, about two-thirds of all titanium metal produced is used in aircraft engines and frames. The SR-71 “Blackbird” was one of the first aircraft to make extensive use of titanium within its structure, paving the way for its use in modern military and commercial aircraft. An estimated 59 metric tons (130,000 pounds) are used in the Boeing 777, 45 in the Boeing 747, 18 in the Boeing 737, 32 in the Airbus A340, 18 in the Airbus A330, and 12 in the Airbus A320. The Airbus A380 may use 146 metric tons, including about 26 tons in the engines. In engine applications, titanium is used for rotors, compressor blades, hydraulic system components, and nacelles. The titanium 6AL-4V alloy accounts for almost 50% of all alloys used in aircraft applications.

Drawbacks: The spacecraft’s aluminium structure cannot withstand temperatures over 175 °C (350 °F) without structural failure. Aerodynamic heating during reentry would push the temperature well above this level in areas, so an effective insulator is needed. This is when the thermal protection system comes into use. One important structural limitation of aluminium alloys is their fatigue strength. Unlike steels, aluminium alloys have no well-defined fatigue limit, meaning that fatigue failure will eventually occur under even very small cyclic loadings. This implies that engineers must assess these loads and design for a fixed life rather than an infinite life.

Aluminium honeycomb skin

Spacecraft structures can be composed of sandwich plates in different basic geometrical forms. Its mechanical properties, in particular, its strength and stiffness under tension, compression, shear and bending loads, depend decisively on the sheet thickness, the cell size and the thickness of the honeycomb cell walls-apart from the pure material properties as such. Due to the manufacturing of the honeycomb core from individual thin ribbons which are folded and glued together, the mechanical properties in the length directions differ from those in the width directions. Since adhesion between the honeycomb core and the sandwich face sheets is achieved in means of an adhesive film which creeps into the fillets between the inner side of the face sheet and the cell walls perpendicular to it and hardens there, the cells would be hermetically closed if venting were neglected. By perforating the ribbon material in the course of honeycomb core manufacturing it is assured that after launching of the spacecraft the air enclosed in the cells can quickly evacuate through small holes of slits. A large number of sandwich parameters makes it possible to be tailored to a variety of sandwich properties but makes analytic calculations complex.


A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys typically have a matrix with an austenitic face-centred cubic crystal structure. A super alloy’s base alloying element is usually nickel, cobalt, or nickel-iron. Superalloy development has relied heavily on both chemical and process innovations and has been driven primarily by the aerospace and power industries. Typical applications are in the aerospace, industrial gas turbine and marine turbine industry, e.g. for turbine blades for hot sections of jet engines.

Examples of superalloys are Hastelloy, Inconel, Waspaloy, Rene alloys (e.g. Rene 41, Rene 80, Rene 95, Rene 104), Haynes alloys, Incoloy, MP98T, TMS alloys, and CMSX single crystal alloys.

Superalloys are commonly used in gas turbine engines in regions that are subject to high temperatures which require high strength, excellent creep resistance, as well as corrosion and oxidation resistance. In most turbine engines this is in the high-pressure turbine, blades here can face temperatures approaching if not beyond their melting temperature. Thermal barrier coatings (TBCs) play an important role in blades allowing them to operate under such conditions, protecting the base material from the thermal effects as well as corrosion and oxidation. Additional applications of superalloys include the following: gas turbines (commercial and military aircraft, power generation, and marine propulsion); space vehicles; submarines; nuclear reactors; military electric motors; chemical processing vessels, and heat exchanger tubing.

Non-metallic Materials

Elastomeric Materials

Flexible Insulation Blankets (FIB). Developed after the initial delivery of Columbia. The white low-density fibrous silica batting material has a quilt-like appearance. The vast majority of the LRSI tiles have been replaced by FIB blankets. They require much less maintenance than LRSI tiles yet have about the same thermal properties.


Polyvinyl chloride, (IUPAC Poly(chloroethanediyl)) commonly abbreviated PVC, is a thermoplastic polymer. It is a vinyl polymer constructed of repeating vinyl groups (ethenyls) having one of their hydrogens replaced with a chloride group. Polyvinyl chloride is the third most widely produced plastic, after polyethylene and polypropylene. PVC is widely used in construction because it is cheap, durable, and easy to assemble. PVC production is expected to exceed 40 million tons by 2016. It can be made softer and more flexible by the addition of plasticizers, the most widely used being phthalates. In this form, it is used in clothing and upholstery, and to make flexible hoses and tubing, flooring, to roofing membranes, and electrical cable insulation. It is also commonly used in figurines and in inflatable products such as waterbeds, pool toys, and inflatable structures.

Fiber Reinforced Plastics

Two other types of tiles, known as FRCI and TUCI (Fibrous Refractory Composite Insulation and Toughened Unipiece Fibrous Insulation), which protect against temperatures between 1,200 and 2,300 degrees Fahrenheit (650 and 1,260 degrees Celsius), are also used in small numbers. FRCI is used in a few areas and TUCI is used on the extreme back of the Orbiter, near the engines. The forward nose cap is made of a material called Reinforced Carbon-Carbon, or RCC. RCC covers the highest temperature areas of the Shuttle and is also used on the leading edges of the wings.

Windows in Space

The SX-4 windows afford an excellent view, and they block out harmful Ultraviolet (UV) rays from the Sun, survive the heat from reentry, and withstand the pressure and temperature differentials throughout the flight. The XP windows and doors are a “plug” design with layers of high-temperature glass and Lexan. Special coatings on the windows keep out harmful UV rays. They provide redundant protection from both the heat of reentry and double-walled containment for the internal pressure.

Windows Skirt Al 2219-T851
Thermal control Goldised Kapton Multi-Layer Insulation blanket
Windows Fused Silica and borosilicate glass
MDPS shutters Al-6061-T6, AL 7075-T7352 and Kevlar/Nextel sheets

Fused Silica and borosilicate glass

Silica (SiO2) is one of the chief constituents of the earth’s crust. It is present in various forms, the most being quartz which is crystalline in character. Typical examples are siliceous sands and rock crystal. There are also various other crystalline forms such as tridymite and cristoblite. All types when fused at 2000°C give a vitreous material. Fused Silica Glass is a unique material with an unrivalled combination of purity, high-temperature resistance, thermal shock resistance, good electrical insulation, optical transparency and chemical inertness. This material is widely used in the production processes of the semiconductor industries. The outstanding characteristic of silica glass is its very high degree of purity (99.99% SiO2).

It also has excellent thermal properties with an extremely low coefficient of expansion of 0.55 x 106cm/cm°C (0-300°C). This makes the material particularly useful for optical flats 7 furnace windows, where its minimal sensitivity to thermal changes is of benefit. Another related property is its high resistance to thermal shock. Thin sections can be heated and cooled rapidly without cracking. Some technical references report, heating the material to 1100°C, then plunging into cold water with no adverse effects.

Borosilicate glass is widely used for laboratory glassware, either mass produced or as custom made. Borosilicate glass has excellent thermal properties with its low coefficient of expansion and high softening point; it also offers a high level of resistance to attack from water, acids, salt solutions, organic solvents and halogens. Resistance to alkaline solutions is moderate and strong alkaline solutions cause rapid corrosion of the glass, as does Hydrofluoric acid and hot concentrated Phosphoric acid.

Systems in Spacecraft

Course Control 

Course control is carried out using sun sensors, simple devices consist of solar cells and baffles. A crude estimate of sun incidence angle is derived from the differential voltages generated by opposing cells (Figure 3). They are often used as part of the hardwired emergency sun re-acquisition system, dedicated to re-aligning the spacecraft with the sun when the main control mode fails.

Spacecraft also utilise other bright objects such, as the earth or bright stars, for crude attitude calculation. Earth sensors scan the sky detecting the planet edges by thermal emission. Orientation information is generated to approximately 0.05, but larger errors can be caused by anomalous atmospheric conditions. Crude star sensors record the position of a particular pre-determined bright celestial body, using its location to align spacecraft axes. These instruments are used for course attitude stabilisation, after the acquisition. The Global Position System (GPS) is also finding increasing application in spacecraft. With two antennae, satellites in earth orbit may not only locate the position, but also calculate orientation. The system is limited to systems requiring only a crude attitude lock (Ý0.5o), but is inexpensive and uses well-proven technology.

Cold Gas Propulsion System

We will be using the 3 Moog 58-126 cold gas thrusters using helium/nitrogen as a propellant for the SX-4. Basically, it works by propelling accelerated gas molecules through the nozzle. The energy need for propulsion is contained in the tank in the form of air pressure. The higher the pressure the more thrust produced.

Data Acquisition and Vehicle Health management system

SX-4  features a Data Acquisition System (DAS) being built by ARINC Company, that collects data from hundreds of sensors placed throughout the vehicle and stores it onboard the vehicle as well as transmits it to the ground. This data is used for an Integrated Vehicle Health Management System (IVHMS) in which the computers onboard SX-4 look for anomalous behaviour from the vehicle’s systems and structure and attempt to diagnose what maybe causing the undesirable readings. If there is any concern, the IVHMS can warn the pilot immediately. The IVHMS will also allow ground staff to determine if a system appears to need routine maintenance or repair. This system makes SX-4 safer to fly and saves valuable time and parts cost, by servicing parts before they fail, and by knowing which areas require service, rather than undertaking exhaustive gremlin chasing between flights. These hardware and software technologies will be embedded in the SX-4’s subsystems, maintenance operations, and launch and mission operations elements, and will provide both real-time and life-cycle vehicle health information which will enable informed decision making and logistics management. Most major subsystems, processing elements, and operations elements are likely to benefit from IVHM technologies, including: propulsion, power, structures, thermal protection systems, avionics, Orbital Maneuvering Systems (OMS), Reaction Control Systems (RCS), crew systems, vehicle turnaround/logistics, launch operations, and mission operations. Another IVHM benefit will be reduced fatigue accumulation rates of systems and structures due to less cycling. Cost benefits include: significantly reduced processing and operations manpower, predictive maintenance for main engines and other high maintenance systems, lower DDT&E costs due to factory enabled Built-In Test (BIT). The improvements in vehicle turnaround and cost will be due to prognostic and diagnostic capability resulting in less actual maintenance being performed due to the detailed vehicle health knowledge from IVHM. A collateral benefit is that there will be fewer maintenance discrepancies and more of the vehicle/component lives used during actual operation.

Electrical Power System Cooling – Using Ammonia

The Electric Power System (EPS) components onboard the SX-4  must be cooled to sustain the space research experiments and prevent system failures due to overheating throughout the spacecraft. The Photovoltaic Thermal Control System’s (PVTCS) radiator rejects heat into space to keep the power system cool.

As a mechanically pumped, single-phase system, the PVTCS is part of the Thermal Control System (TCS). It can be controlled manually by the astronauts or remotely from the ground via the Photovoltaic Control Unit (PVCU). Using ammonia coolant, the PVTCS keeps the primary EPS components within their proper temperature range by transporting excess heat from the electrical equipment assemblies, batteries and radiators into space.

The PVTCS consists of three main parts: the Integrated Equipment Assembly (IEA) structural framework, the Pump Flow Control Subassembly (PFCS), and the Photovoltaic Radiator (PVR). The cooling system plugs into the IEA framework. The PFCS controls the flow of ammonia coolant to the TCS while the PVR rejects the heat from the photovoltaic electronics into deep space. The PVTCS components work together to help maintain the functionality of the EPS and its related systems while ensuring the safety of the astronauts.

Electrical Power System

Fuel Cell Name: Proton exchange membrane fuel cell

Electrolyte Polymer membrane (ionomer) (e.g., Nafion or Polybenzimidazole fiber)

Qualified Power (W) :100 W to 500 kW

Working Temperature (°C) :(Nafion)50–120 (PBI)125–220

Electrical efficiency  Cell: 50–70% System: 30–50%

An Electrical Power system will be vital for the aircraft like any other. A smaller version for the aircraft should be made. A cooling system has also been chosen for this. The SX-4 spends much of its time flying either exoatmospherically or in glide as it returns to base. During these portions of the flight, it will have no running jet engines to run generators or other power producing devices and therefore uses stored energy systems. For this purpose, the SX-4 is equipped with a series of large Lithium-Ion batteries. These batteries provide power to SX-4’s systems at 28volts and 270volts Direct Current (DC). The electric flight control actuators use the 270VDC while the computers and other systems use the 28VDC. The Lithium Ion power system is redundant and fault-tolerant and is designed to be recharged onboard the plane.

Emergency Systems

The overall safety of the spacecraft is protected by an emergency mode, which automatically activates if other control modes fail. Often implemented as an entirely hardwired system, or at least securely partitioned from the main attitude control task in the attitude processor, the emergency mode prevents untimely loss of the spacecraft in the event of failure or glitch (single event up-set). It employs separate sensors ensuring that in the event of a single sensor failure, the spacecraft may be recovered. Attitude determination and control is usually minimal, the system simply locking solar arrays onto the sun, minimising angular rate to prevent damage to deployed booms and sensors, notifying ground control of emergency mode activation and awaiting further instructions.

While attitude determination subsystems are designed with a wide variety of requirements, missions and spacecraft, the sensors suites they utilise are often very similar. The following sections describe three categories of sensor devices: rate sensors, coarse pointing sensors and fine pointing sensors.

Caution and Warning System

The primary caution and warning system is designed to warn the crew of conditions that may adversely affect SX-4’s operations. The system consists of hardware and electronics that provide the crew with both visual and aural cues when a system exceeds predefined operating limits. The primary system’s visual cues consist of four master alarm lights, a 40-light array on panel F7 and a 120-light array on panel R13. The aural cue is sent to the communications system for distribution to flight crew headsets or speaker boxes.

The C/W system interfaces with the auxiliary power units, data processing system, environmental control and life support system, electrical power system, flight control system, guidance and navigation, hydraulics, main propulsion system, reaction control system, orbital manoeuvring system and payloads. The audio alarms are classified as an emergency (class 1), C/W (class 2) and alert (class 3).

The emergency alarms consist of a siren (activated by the smoke detection system) and a klaxon (activated by the delta pressure/delta time sensor that recognizes a rapid loss of cabin pressure), and they are annunciated by hardware. The siren’s frequency varies from 666 to 1,470 hertz and returns at a five-second-per-cycle rate. The klaxon is a 2,500-hertz signal with an on/off cycle of 2.1 milliseconds on and 1.6 milliseconds off, mixed with a 270-hertz signal with a cycle of 215 milliseconds on and 70 milliseconds off.

Area’s of concern on emergency Systems

Life support systems safety

  • Atmospheric conditioning and control
    • Atmospheric conditioning
    • Carbon dioxide removal
  • Trace contaminant control
    • Trace contaminant control methodology
    • Trace contaminant control design considerations
  • Assessment of water quality in the spacecraft environment: mitigating health and safety concerns
    • Scope of water resources relevant to spaceflight
    • Spacecraft water quality and the risk assessment paradigm
    • Water quality monitoring
  • Waste management

Emergency systems

  • Space rescue
  • Personal protective equipment

Collision Avoidance systems

  • Docking systems and operations
    • Orbital mating
    • Design approaches ensuring docking safety and reliability
  • Descent and lading systems
    • Parachute systems
    • Known parachute anomalies and lessons learnt

Robotics systems safety

  • Generic robotic systems
    • Controller and operator interface
    • Arms and joints
    • Drive system
    • Sensors
    • End effector
  • Identification of hazards and their causes
    • Electrical and electromechanical malfunctions
    • Mechanical and structural failures
    • Failure in the control path
    • Operator error
  • Hazard mitigation in design
    • Electrical and mechanical design and redundancy
    • Operator error
    • System health checks
    • Emergency motion arrest
    • Proximity operations
    • Built-in test
    • Safety algorithms

Meteoroid and debris protection

  • Risk control measures
    • Maneuvering
    • Shielding
  • Emergency repair considerations for spacecraft pressure wall damage
    • Balanced mitigation of program risks
    • Leak location system and operational design considerations
    • Ability to access the damaged area

Noise control design

  • Noise control plan
    • Acoustic analysis
    • Testing and verification
  • Noise control design applications
    • Noise control at the source
    • Noise control at the receiving space
    • Post-design noise mitigation

Materials safety

  • Toxic off-gassing
    • Materials offgasing controls
    • Materials testing
    • Spacecraft module testing
  • Stress-corrosion cracking
    • Prevention of stress-corrosion cracking
    • Testing materials for stress-corrosion cracking
    • Design for stress-corrosion cracking
    • Requirements for spacecraft hardware
    • Stress-corrosion cracking in propulsion systems

Oxygen systems safety

  • Oxygen pressure system design
    • Design approach
    • Oxygen compatibility assessment process
  • Oxygen generators
    • Electromechanical systems for oxygen production
    • Solid fuel oxygen generators (oxygen candles)

Avionics safety

  • Electrical grounding and electrical bonding
    • Control of electric current
    • Electrical grounds can be signal return paths
    • Where and how electrical grounds should be connected
    • Defining characteristics of an electrical bond
    • Types of electrical bonds
    • Electrical bond considerations for dissimilar metals
    • Electrical ground and bond connections for shields
  • Safety-critical computer control
    • Partial computer control
    • Total computer control: Failsafe
  • Circuit protection: Fusing
    • Circuit protection methods
    • Circuit protectors
  • Electrostatic discharge control
    • Various levels of electrostatic discharge concern
  • Arc tracking
    • Characteristics or arc tracking
    • Likelihood of an arc tracking event
    • Prevention or arc tracking
    • Verification of protection and management of hazards
  • Corona control in high voltage systems
    • Design criteria
    • Verification and testing
  • Extravehicular activity considerations
    • Displays and indicators used in space
    • Mating and demating of powered connectors
    • Single strand melting points
    • Battery removal and installation
    • Computer or operation control of inhibits
  • Spacecraft electromagnetic interference and electromagnetic compatibility control
    • Electromagnetic compatibility needs for space applications
    • Basic electromagnetic compatibility interactions and a safety margin
    • Mission-driven electromagnetic interference design: The case for grounding
    • Electromagnetic compatibility program for spacecraft
  • Design and testing of safety-critical circuits
    • Safety-critical circuits: conducted mode
    • Safety-critical circuits: Radiated mode
  • Electrical hazards
    • Electrical shock
    • Physiological considerations
    • Electrical hazard classification
    • Leakage current
    • Bioinstrumentation
    • Electrical hazard controls
    • Verification of electrical hazard controls
    • Electrical safety design considerations
  • Avionics lessons learned
    • Electronic design
    • Physical design
    • Materials and sources
    • Damage avoidance
    • System aspects

Software systems safety

  • The software safety problem
    • System accidents
    • The power and limitations of abstraction from physical design
    • Reliability versus safety for software
    • Inadequate system engineering
    • Characteristics of an embedded software
  • Best practice
    • Management of software-intensive, safety-critical projects
    • Basic system safety engineering practices and their implications for software-intensive systems
    • Specifications
    • Requirements analysis
    • Model-based software engineering and software reuse
    • Software architecture
    • Software design
    • Design of human-computer interaction
    • Software reviews
    • Verification and assurance
    • Operations

Battery safety

  • General design and safety guidelines
  • Battery models
  • Battery types
  • Hazard and toxicity categorization
  • Battery chemistry
    • Alkaline batteries
    • Lithium batteries
    • Silver zinc batteries
    • Lead acid batteries
    • Nickel cadmium batteries
    • Nickel metal hydride batteries
    • Nickel hydrogen batteries
    • Lithium-ion batteries
  • Storage, transportation, and handling

Mechanical systems safety

  • Safety factors
    • Types of safety factors
    • Safety factors typical of human rated space programs
    • Things the influence the choice of safety factors
  • Spacecraft structures
    • Mechanical requirements
    • Space mission environment and mechanical loads
    • Project overview: successive designs and iterative verification of structural requirements
    • Analytical evaluations
    • Structural test verification
    • Spacecraft structural model philosophy
    • Materials and processes
    • Manufacturing of spacecraft structures
  • Fracture control
    • Basic requirements
    • Implementation
  • Pressure vessels, lines and fittings
    • Pressure vessels
    • Lines and fittings
    • Space pressure systems standards
  • Composite overwrapped pressure vessels
    • The composite overwrapped pressure vessel system
    • Monolithic metallic pressure vessel failure
    • Composite overwrapped pressure vessel failure modes
    • Composite overwrapped pressure vessel impact sensitivity
  • Structural design of glass and ceramic components for space system safety
    • Strength characteristics of glass and ceramics
    • Defining loads and environments
    • Design factors
    • Meeting life requirements with glass and ceramics
  • Safety critical mechanisms
    • Designing for failure tolerance
    • Design and verification of safety critical mechanisms
    • Reduced failure tolerance
    • Review of safety critical mechanisms

Containment of hazardous materials

  • Toxic materials
    • Fundamentals of toxicology
    • Toxicological risks to air quality in spacecraft
    • Rick management strategies
  • Biohazardous materials
    • Microbiological risks associated with spaceflight
    • Risk mitigation approaches
    • Major spaceflight specific microbiological risks
  • Shatterable materials
    • Shatterable materials in a habitable compartment
    • Program implementation
    • Containment concepts for internal equipment
    • Containment concepts for exterior equipment
    • General comments about working with Shatterable materials
  • Containment design approach
    • Fault tolerance
    • Design for minimum risk
  • Containment design approach
    • Containment environments
    • Design of containment systems
  • Safety controls
    • Proper design
    • Materials selection materials compatibility
    • Proper workmanship
    • Proper loading or filling
    • Fracture control
  • Safety verifications
    • Strength analysis
    • Qualification tests
    • Acceptance tests
    • Proof tests
    • Qualification of procedures

Propellant systems safety

  • Solid propellant propulsion systems safety
    • Solid propellants
    • Solid propellant systems for space application
    • Safety hazards
    • Handling, transport and storage
    • Inadvertent ignition
    • Safe ignition systems design
  • Liquid propellant propulsion systems safety
    • Containment integrity
    • Thermal control
    • Materials compatibility
    • Containment control
    • Environmental considerations
    • Engine and thruster firing inhibits
    • Heightened risk
    • Instrumentation and telemetry data
    • End to end integrated instrumentation, controls, and redundancy verification
    • Qualification
    • Total quality management
    • Preservicing integrity verification
    • Propellants servicing
  • Hypergolic propellants
    • Materials compatibility
    • Material degradation
    • Hypergolic propellant degradation
  • Propellant fire
    • Hydrazine and monomethylhydrazine vapor
    • Liquid hydrazine and monomethylhydrazine
    • Hydrazine and monomethylhydrazine mists, droplets and sprays

Pyrotechnic safety

  • Pyrotechnic devices
    • Explosives
    • Initiators
  • Electroexplosive devices
    • Safe handling of electroexplosive devices
    • Designing for safe Electroexplosive device operation
    • Pyrotechnic safety of mechanically initiated explosive devices

Extravehicular activity safety

  • Extravehicular activity environment
    • Extravehicular activity space suit
    • Sensory degradation
    • Maneuvering and weightlessness
    • Glove restrictions
    • Crew fatigue
    • Thermal environment
    • Extravehicular activity tools
  • Suit hazards
    • Inadvertent contact hazards
    • Area of effect hazards
  • Crew hazards
    • Contamination of the habitable environment
    • Thermal extremes
    • Lasers
    • Electrical shock and molten metal
    • Entrapment
    • Emergency ingress
    • Collision
    • Inadvertent loss of crew

Emergency, caution and warning system

  • System overview
  • Historic NASA emergency, caution and warning systems
  • Emergency, caution and warning system measures
    • Event classification measures
    • Sensor measures
    • Data system measures
    • Annunciation measures
  • Failure isolation and recovery

Laser safety

  • Optical spectrum
  • Biological effects
  • Laser characteristics
    • Laser principles
    • Laser types
  • Laser standards
    • NASA Johnson space center requirements
    • ANSI standard Z136-1
    • Russian standard
  • Lasers used in space
    • Radars
    • Illumination
    • Sensors
  • Design considerations for laser safety
    • Ground testing
    • Unique space environment

Crew training safety: An integrated process

  • Training the crew for safety
    • Typical training flow
    • Principles of safety training for the different training phases
    • Specific safety training for different equipment categories
    • Safety training for different operations categories
  • Safety during training
    • Training, test or baseline data collection model versus flight model: type, fidelity, source, origin and category
    • Training environments and facilities
    • Training models, test models and safety requirements
    • Training model, test model and baseline data collection equipment utilization requirements
    • Qualification and certification of training personnel
    • Training and test model documentation
  • Training development and validation process
    • The training development process
    • The training review process
    • The role of safety in the training development and validation process
    • Feedback to the safety community from the training development and validation processes

Safety considerations for the ground environment

  • Roles and responsibilities
  • Contingency planning
  • Failure tolerance
  • Training
  • Hazardous operations
  • Tools
  • Human factors
  • Biological systems and materials
  • Electrical
  • Radiation
  • Pressure systems
  • Ordinance
  • Mechanical and electromechanical devices
  • Propellants
  • Cryogenics
  • Oxygen
  • Ground handling
  • Software safety

Fire safety

  • Characteristics of fire in space
    • Overview of low gravity fire
    • Fuel and oxidizer supply and flame behavior
    • Fire appearance and signatures
    • Flame ignition and spreads
  • Design for fire prevention
    • Materials flammability
    • Ignition sources
  • Spacecraft fire detection
    • Prior spacecraft systems
    • Review of low gravity smoke
    • Spacecraft atmospheric dust
    • Sensors for fire detection
  • Spacecraft fire suppression
    • Spacecraft fire suppression methods considerations for spacecraft fire suppression

Environmental Control & Life Support System

The SX-4, Environmental Control and Life Support System (ECLSS) provide or controls atmospheric pressure, fire detection and suppression, oxygen levels, waste management and water supply. The highest priority for the ECLSS is the SX-4’s atmosphere, but the system also collects, processes, and stores waste and water produced and used by the crew—a process that recycles fluid from the sink, shower, toilet, and condensation from the air. The Elektron system aboard sx-4 will generate oxygen aboard the station. The crew has a backup option in the form of bottled oxygen and Solid Fuel Oxygen Generation (SFOG) canisters. Carbon dioxide is removed from the air by the Vozdukh system. Other by-products of human metabolisms, such as methane from the intestines and ammonia from sweat, are removed by activated charcoal filters. The atmosphere on board the SX-4 is similar to the Earth’s. Normal air pressure on the SX-4 is 101.3 kPa the same as at sea level on Earth. An Earth-like atmosphere offers benefits for crew comfort and is much safer than the alternative, a pure oxygen atmosphere.

Onboard Oxygen Generating System 

Similar to the elektron, the Russian oxygen and NASA’S Oxygen Generating System. The generator used on board the International Space Station (ISS).The methods uses electrolysis to produce oxygen. The process splits water molecules reclaimed from other uses on board the station into oxygen and hydrogen via electrolysis. The Oxygen is vented into the cabin while the hydrogen is vented into space. If Singapore could develop a smaller version of this we can use it for a longer stay in space. Without the usage of such devices, we have to carry onboard oxygen supply. Look at SX-4 Diagram for location. We can use bottled oxygen or Solid Fuel Oxygen Generation canisters.

Rate Sensor System

For initial acquisition modes, where the spin rate of the spacecraft must be controlled to attain a first inertial lock, sets of rate sensors are employed for each spacecraft axis. Gyroscope devices are widely used. High precision mechanical gyros are common but suffer from high drift rates and are not always reliable. Solid state devices such as ring laser gyros (Figure 2) are increasingly considered for new missions, as are quartz rate sensors.

These devices all measure rotational motion around a principal axis. They are usually mounted in the same configuration as reaction wheels in packs of four: in a tetrahedral arrangement or with three mounted orthogonally and a fourth at a skew angle to allow some redundancy to failure. Since their signal must be integrated to give angular position, signal error makes them unsuitable for measuring absolute angles as the measurement will drift with time. Rate measurement is limited to a predetermined range configured to meet the control system specifications.

Reaction Control System

RCS can be used to control altitude, rotational maneuvers such as yaw, pitch and roll and small velocity changes. The movement made by the RCS can be used for out space rendezvous.

Space Shuttle :

The forward RCS has 14 primary and two vernier engines. The aft RCS has 12 primary and two vernier engines in each pod. The primary RCS engines provide 870 pounds of vacuum thrust each, and the vernier RCS engines provide 24 pounds of vacuum thrust each. The oxidizer-to-fuel ratio for each engine is 1.6-to-1. The nominal chamber pressure of the primary engines is 152 psia. For each vernier engine, it is 110 psia.

The primary engines are reusable for a minimum of 100 missions and are capable of sustaining 20,000 starts and 12,800 seconds of cumulative firing. The primary engines are operable in a maximum steady-state thrusting mode of one to 150 seconds, with a maximum single-mission contingency of 800 seconds for the aft RCS plus X engines and 300 seconds maximum for the forward RCS minus X engines as well as in a pulse mode with a minimum impulse thrusting time of 0.08 second above 125,000 feet. The expansion ratio (exit area to throat area) of the primary engines ranges from 22-to-1 to 30-to-1. The multiple primary thrusters provide redundancy.

The vernier engines’ reusability depends on chamber life. They are capable of sustaining 330,000 starts and 125,000 seconds of cumulative firings. The vernier engines are operable in a steady-state thrusting mode of one to 125 seconds maximum as well as in a pulse mode with a minimum impulse time of 0.08 second. The vernier engines are used for finite manoeuvres and stationkeeping (long-time attitude hold) and have an expansion ratio that ranges from 20-to-1 to 50-to-1. The vernier thrusters are not redundant.

Each RCS consists of high-pressure gaseous helium storage tanks, pressure regulation and relief systems, a fuel and oxidizer tank, a system that distributes propellant to its engines, and thermal control systems


The pressurization system

Each RCS has two helium storage tanks, four helium isolation valves, four pressure regulators, two relief valves, two check valves, two manually operated valves and servicing connections for draining and filling.

The helium storage tanks are composite spheres and consist of a titanium liner with a Kevlar structural overwrap that increases safety and decreases the tank weight over conventional titanium tanks. Each helium tank is 18.71 inches in diameter with a volume of 3,043 cubic inches and a dry weight of 24 pounds. Each helium tank is serviced to 3,600 psi.

Propellant system

The system that distributes the propellants to the RCS thrusters consists of fuel and oxidizer tanks, tank isolation valves, manifold isolation valves, crossfeed valves, distribution lines and filling and draining service connections.Each RCS contains two spherical propellant tanks, one for fuel and one for oxidizer, constructed of titanium and 39 inches in diameter.The nominal full load of the forward and aft RCS tanks in each pod is 1,464 pounds(664kg) in the oxidizer tanks and 923 pounds(418kg) in the fuel tanks. The dry weight of the forward tanks is 70.4 pounds(32kg). The dry weight of the aft tanks is 77 pounds(35kg).

RCS engine

Each RCS engine contains a fuel and oxidizer valve, injector head assembly, combustion chamber, nozzle and electrical junction box.Each primary RCS engine has one fuel and one oxidizer solenoid-operated pilot poppet valve that is energized open by an electrical thrust-on command, permitting the propellant hydraulic pressure to open the main valve poppet and allow the respective propellant to flow through the injector into the combustion chamber. When the thrust-on command is terminated, the valves are de-energized and closed by spring and pressure loads.

Thermal Protection System

When the SX-4 reenters Earth’s atmosphere the leading edges of the wings, tails, flight control surfaces, and nosecone will experience temperatures in excess of 600-700 degrees Fahrenheit. These surfaces are made of titanium which is not only very temperature resistant but light as well. The rest of the vehicle is covered in a special ceramic paint that emits over 93% of the energy it receives. In select areas such as near the rocket engine, ceramic blankets will be used to protect the fuselage from not only the heat of reentry, but also the heat produced by the rocket engine.

This article is written in memory, and in credit to the following team members

Simon Soh


Noorul Ameen

Yogendra Raj


Article on Aerospace Magazine 2011,Singapore
Flying into the future Competition – Chinese Society of Astronautics

Combined Sources

-Metallurgy for engineers ( E.C ROLLASON) FOURTH EDITION.
Pages 327,351
-Aircraft Structures for engineering Students (fourth edition)
Pages 353,356
-(Design of aircraft by Thomas C corke) Whole book
-Aircraft performance and design by (John D.Anderson ,JR) Whole book
-Aircraft Conceptual Design Synthesis by Denis Howe
-Aircraft Design : A conceptual Approach Daniel P.Raymer
-Safefty Design and Space systems Edited by Gary Eugene Musgrave,M,larsen,Tommaso Sgobba

Author: adlabs2

Arrowdynamic Laboratories Pte. Ltd. (ADLABS) was founded in 2008 anticipating imminent technological pervasion. ADLABS operates with two wings, Education and Engineering. Education involves the development of skills training programs and STEM-based enrichment programs for everyone. The engineering wing conducts research and development activities on new prototypes with potential to form into a new branch startup. Our engineering team also offers specialised services for startups, companies and institutions.

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