This paper explores different types of rocket propulsion systems and assesses their benefits and drawbacks. The paper also compares different types of rockets, such as Single Stage rockets and Two Stage rockets, before proposing a propulsion system for a future Single Stage to Orbit (SSTO) rocket. This system shall meet the requirements of both reusability, cost effectiveness, high performance and environmental friendliness.
The allure of space travel pervades cultures and nations; the idea of exploring our cosmos is prevalent even today – more than 50 years after Neil Armstrong’s “giant leap for mankind”. The feasibility of future manned spaceflight is dependent on many factors – one of which is the creation of a suitable propulsion system.
A desirable propulsion system enables a rocket to fly as fast as possible, for as long as possible whilst being cheap and environmentally-friendly. Achieving such a set of criteria poses numerous challenges in itself.
The propulsion systems of aircraft and spacecraft are alike in many respects: both must maximise efficiency and performance whilst minimising cost. However, the difference between the two systems lies in the nature of space itself. Aircraft engines work by sucking in air using massive fans located at the front of the engine. After the air is sucked in, it is compressed and sprayed with fuel (usually kerosene). This mixture of air and kerosene is then ignited by an electric spark and the burning gases are pushed out the back of the engine, propelling the aircraft forward (forward thrust). This phenomenon relies on the combustion of atmospheric oxygen and kerosene.
However, as space is a vacuum, atmospheric oxygen isn’t present and thus engine fans are impractical. Hence, rocket engines instead have to carry their own oxidiser  in addition to carrying propellant (fuel). However, carrying one’s own oxidiser increases the weight of the rocket – thus more propellant is required to provide sufficient thrust to overcome the increased weight. Engineers therefore look to maximise the thrust produced by a propulsion system whilst minimising the weight of the overall spacecraft . This idea is fundamental to rocket propulsion.
Spacecraft must escape Earth’s gravitational field to reach space but achieving this requires an escape velocity of 11,000 metres per second. For comparison, the SR-71 Blackbird (the fastest ever aircraft) had a top speed of 980 m/s (11 times lower) . Getting such a high escape velocity requires a lot of thrust and thus a lot of propellant.
This fuels another important consideration in rocket propulsion: the propellant mass fraction (PMF). The majority of a rocket’s mass is propellant: SpaceX’s Falcon 9 rocket has a PMF of 0.96 – just 4% of the rocket’s mass can be devoted to its payload and structure. In comparison, an average car has a PMF of 0.03 and a fighter jet 0.30. Too low a PMF and one doesn’t have enough propellant to reach escape velocity but too high a PMF and the design proves impractical: a PMF of 0.99 leaves only 1% of the rest of the rocket’s mass for its structure and payload (control systems, crew etc.). One cannot design the structure and payload to weigh 1% of the rocket’s total mass without adversely affecting the integrity of the rocket (or greatly increasing its weight) – hence engineers must find a balance between high and low PMF.
Payload mass refers to the mass of the equipment and personnel carried by a spacecraft. The greater the payload mass, the more ‘useful’ a spacecraft is (the more things it can carry). However, maximising payload mass is a big challenge in modern spaceflight as this requires reducing the structural mass in a way that doesn’t compromise the robustness of the spacecraft. By reducing structural mass, one can reallocate the weight savings to bolster payload mass, thereby increasing utility.
Specific impulse (Isp) is a measure of how efficiently a rocket uses propellant (by definition, the total change in momentum per unit weight of propellant consumed). The higher the Isp (measured in seconds), the more efficiently a propulsion system uses fuel hence designers strive to maximise the Isp of rocket fuels to improve rocket performance. As explored in the next section, a higher Isp also means that the rocket flies faster.
Konstantin Tsiolkovsky was a Soviet rocket scientist known as the ‘Russian father of rocketry’ for his numerous contributions to space exploration: one of which was deriving the classical rocket equation. Tsiolkovsky’s Rocket Equation (Eq. 1) is one of the fundamental equations of rocket science as it links the desired change in velocity to the mass of the propellant and the specific impulse of the fuel.
Tsiolkovsky’s Classical Rocket Equation: this shows that the change in velocity (Delta V) is equal to the exhaust velocity (Vex) multiplied by the natural logarithm of the ratio of a rocket’s initial mass to its final mass
The faster the rocket, the better, thus to achieve a high delta V (to make the rocket fly faster), a high exhaust velocity and/or a high initial mass to final mass ratio is necessary. By decreasing the structure and payload mass of the spacecraft, the PMF increases without changing the mass of propellant, thereby increasing mass lost and thus increasing initial mass to final mass ratio: increasing delta V. Exhaust velocity can be related to specific impulse through Eq. 2 below.
Eq. 2 shows that exhaust velocity is equal to Earth’s gravitational acceleration (g0) multiplied by the specific impulse of a fuel. Thus, the higher a fuel’s Isp, the higher the exhaust velocity of the working fluid and (from the Rocket Equation) the higher the consequent change in velocity.
Thus, minimising the structural mass and maximising the Isp increases delta V.
Eq. 3 shows that the rate of change of mass of the propellant (dm/dt – the derivative of mass) – or how quickly the propellant is being used up – is calculated by dividing the thrust produced (F) by the specific impulse of the fuel (Isp). Thrust is the driving force of the rocket. A low propellant consumption rate is beneficial as this means that the same mass of fuel powers the rocket for longer. This can be achieved by, again, increasing the specific impulse of the fuel (propellant) one is using. Therefore, using a high Isp fuel both increases engine performance and decreases propellant consumption rate, thereby achieving the criteria of efficiency and performance.
Figure 1: Graph of maximum velocity against initial mass to final mass ratio with varying specific impulses.
Figure 1 aptly shows the extent to which specific impulse can impact delta V: by doubling the specific impulse from 350s to 700s, delta V subsequently doubles. Currently, chemical rockets (which are the traditional rockets currently in use) have a maximum specific impulse of 450s, but this graph is testament to the impact that small increases in Isp can have on delta V and rocket performance.
Along with a high Isp, propellants (fuel) should be dense – maximising the mass of propellant in a given volume also maximises energy release per unit volume – thus increasing the efficiency of the rocket. This means that a denser fuel needs a lower volume of propellant to yield the same increase in velocity. The propellant mass remains the same here but by reducing the volume of propellant, one can use a more compact structure – which reduces the structural mass – providing scope for the payload mass to increase.
Another aspect of rocket propulsion systems is regression rate/burn rate. This is a measure of how quickly an exposed propellant surface burns (rather than how quickly it is consumed). A high regression rate propellant burns vigorously – increasing the energy produced by burning the propellant (in a given time) – hence a high propellant regression rate is desirable to maximise thrust .
Figure 2: Mind Map diagram of the various types of rocket propulsion systems
|Specific Impulse||High (around 450s)||Low (around 270s)||Medium (around 350s)|
|Ubiquity||Very common||Somewhat common||Relatively uncommon|
Chemical propulsion produces thrust via a chemical reaction (unlike non-chemical systems). Air-breathing engines are similar to jet engines in that they use atmospheric oxygen as oxidiser. This article focuses mainly on chemical propulsion systems as they are the most common.
Figure 3: Space Shuttle taking flight – note the white solid boosters
Solid propulsion systems (like those used in the white boosters of the Space Shuttle in Figure 3) have both propellant and oxidiser in the solid state. Solid rocket engines work by using an igniter to set fire to a solid propellant grain, which contains both fuel and oxidiser moulded into a cylindrical shape. As the grain itself contains the necessary ingredients for combustion (oxidiser and propellant), once it is lighted, it shall continue to burn. This means that the propellant burn rate cannot be altered once ignition is initiated (the rocket cannot be throttled up or down) nor can the combustion process be stopped and started again (not restartable). The combustion of the propellant grain releases toxic nitrogen oxides and hydrogen chloride gas into the atmosphere – both of which combine with water vapour in the air to form acid rain. Thus, in a time where environmental sustainability is pivotal, solid rockets do not meet that criteria. The main problem with solid rocket motors, however, is their poor performance: solid propellants have substantially lower Isps than their hybrid and liquid counterparts making them highly inefficient. To rectify this, engineers can ‘dope’ the propellant grain with metal additives such as aluminium powder, which can increase both specific impulse and burn rate. Thus, solid rockets are cheap and uncomplicated but suffer from subpar performance and toxicity to the environment .
Liquid rockets, like those attached to the three glowing nozzles on the Space Shuttle itself (in Figure 3), have been the primary propulsion system in spaceflight for many years now and for good reason . Liquid rockets store propellant and oxidiser in separate compartments (1 and 2 respectively in the figure below) and siphon these into a combustion chamber, where the mixture is ignited. Some liquid rockets are pressure-fed, meaning that the compartments are highly-pressurised to force the oxidiser and propellant into the combustion chamber, but most liquid rockets use turbo-pumps to do this instead – which are lighter. Liquid rockets are ubiquitous as they deliver the highest specific impulse of any chemical rocket (up to 450s) and are throttleable, despite their increased complexity and cost .
Figure 4: Schematic of hybrid rocket
Hybrid rocket engines usually have a cylindrical solid propellant (fuel) grain which lines the combustion chamber and a liquid oxidiser. Combustion works by the oxidiser being injected into the cylindrical grain before the mixture is ignited (see Figure 4). Hybrids are a cost-effective intermediary propulsion mode reaping both the benefits of solid motors (simplicity) and those of liquid engines (restartability/throttling). In addition to these, hybrid motors have their own benefits:
- Inherent safety – Oxidiser and propellant cannot intimately mix – thus preventing an overmixing of the two which could lead to an explosion caused by the spike in energy and pressure release
- Performance enhancement – It is easy to alter the performance of a hybrid engine by ‘doping’ – such as the metal additives we discussed earlier. Moreover, by using a gelled propellant grain (gelation), one can increase both the density of the fuel and the Isp.
- Light – By not requiring the same extent of piping that a liquid motor design does, hybrid rocket engines are usually lighter than their liquid counterparts.
Hybrid engines have two main drawbacks:
- Hybrid fuels tend to have lower specific impulses than comparable liquid fuels
- The propellant grain has a lower regression/burn rate than that of solids (a typical hybrid propellant grain has a regression rate 1/10th that of a typical solid grain – 0.1cm/s for the hybrid and 1 cm/s for the solid). To compensate for this, engineers are forced to increase the surface area of the propellant grain (by making it star-shaped), but this only adds weight. Efforts to increase regression rate by 50-100% have so far not been possible without compromising safety– hence future hybrid powered spacecraft must overcome this challenge.
Despite these setbacks, the benefits of hybrid motors make them an attractive option for the future; the idea of cheap, reusable, environmentally friendly rockets can be feasible with an increase in regression rate and specific impulse – hence this article shall focus on hybrid rockets given the potential rewards of this concept.
Non-chemical propulsion systems, although very much in the development phase, are an exciting prospect to come given the massive advances they bring in specific impulse. This section shall look at two major non-chemical propulsion types: nuclear and electrical.
Nuclear Thermal Rockets (NTRs) are a proposed form of future propulsion utilising nuclear fission to propel the rocket. In an NTR, a propellant such as liquid hydrogen is passed through a nuclear reactor (where nuclear fission of Uranium-235 occurs: each U-235 atom splits into two smaller atoms, releasing heat energy in the process). In the reactor core, the hydrogen is heated to a high temperature by the heat from the fission reaction, before the hot gases are expelled to produce thrust. Hence, no chemical reaction takes place in an NTR. NTR rockets generate an Isp of around 825s, nearly double that of a liquid rocket but in doing so, NTRs operate at very high temperatures (as Isp is proportional to the square root of the working fluid temperature), hence the engine must be constructed with material that can both withstand these temperatures and retaining lightness and strength. However, the main issue with NTRs is potential radiation fallout – necessitating in the instalment of heavy neutron shielding to protect the crew from nuclear meltdown. Moreover, the 1972 United Nations Liability Convention dictates that if, for example, a US spacecraft crashes into London, it is the US government who must pay the bill. However, the subsequent radiation fallout from a crashed NTR rocket would be both catastrophic and very expensive – thus although NTR motors are a promising concept, the risks involved with using nuclear reactors outweigh the gains in specific impulse.
As the name suggests, electric propulsion (EP) systems use electrical power to accelerate a propellant. EP systems greatly increase the speed at which the propellant is thrust out the back of the spacecraft – up to 20 times faster than a conventional chemical rocket. Consequently, EP systems have much higher Isps than conventional chemical systems: the Nexis thruster (the result of a collaborative effort between the California Institute of Technology’s Jet Propulsion Laboratory and Boeing) can sustain specific impulses of over 10,000 seconds (nearly 20 times a conventional Isp). EP systems work by ionising a heavy gas such as Xenon before accelerating the ions in an electric field (with high propulsive efficiency) and using a magnetic field to focus the ion beam (rather than a conventional nozzle) upon leaving the spacecraft. However, EP systems are subject to major drawbacks:
- EP systems don’t work unless in a vacuum – Ion engines don’t function in the presence of external ions outside the engine, thereby rendering the system useless for taking off from Earth
- EP systems yield very low thrust – we can deduce this through the propellant consumption rate formula: EP systems have a high propulsive efficiency and thus a low consumption rate (dm/dt is low) and have extremely high Isps, thus thrust generation is minimal. This essentially prohibits EP spacecraft from taking off anywhere with a gravitational field (as the thrust cannot overcome the weight of the spacecraft) but doesn’t prohibit the use of such spacecraft for interplanetary travel
Thus, neither EP nor NTR systems are feasible for manned spacecraft launches from Earth as their drawbacks essentially proscribe their usage for this purpose.
The next section explores another fundamental aspect of spacecraft design: rocket staging.
|Characteristic||Single Stage to Orbit (SSTO)||Two/Three Stages to Orbit (TSTO)|
|Gross Lift off Weight (GLOW)||Higher||Lower (less propellant needed)|
|Complexity||Less complex||More comples|
|Cost||Lower (from reduced complexity)||Higher|
|Risk||Lower risk||Higher risk (from complexity)|
|Propellant Mass Fraction (PMF)||Higher (due to high GLOW)||Lower|
|Reusability||Higher reusability||Lower reusability|
Figure 5: Diagram of a three stage to orbit (TSTO) rocket
Rocket staging separates the rocket into smaller individual stages, each carrying its own propellant and structure. Each stage is discarded once all its propellant is exhausted. The two main types of rocket staging are:
- Single Stage to Orbit (SSTO): no staging occurs on SSTO rockets – all the required fuel is carried onboard and there is no discarding of the structure once finished. Despite the concept not being new, a viable SSTO concept hasn’t been found as of yet: the aim being for a reusable SSTO where relaunching is done by simply refuelling the rocket.
- Two/Three Stage to Orbit (TSTO): the rocket is divided into three stages with each stage’s design being customised for its purpose (Figure 5). For example, the first stage of a TSTO rocket is specifically designed to get the rocket to escape velocity as quickly as possible whilst upper stages are designed to operate in the vacuum of space. Anything more than three stages is simply impractical hence we shall only compare SSTOs to TSTOs. Most modern rockets utilise a multi-stage approach – the reasoning behind this is explained in the following section.
This section provides a real-life example of the rocket equation to explain why TSTO rockets are currently favourable.
Consider a rocket carrying 100,000 kg of propellant, 10,000 kg of structure, and 5000 kg of payload. Recall that the relationship between specific impulse and exhaust velocity is:
Assuming that Isp is 450s and g0 is 9.81 ms-2, calculate ΔV given that all the propellant is consumed in one stage?
Here, the initial mass of the rocket is the sum of the propellant, structure and payload – 115,000kg. The final mass of the rocket is just the structure and payload (as the propellant has been used up but the structure hasn’t been discarded) – or 15,000kg. Therefore, the ratio of initial mass to final mass is 115,000/15,000 (or 23/3). Resubstituting these values back into the Rocket Equation gives us…
Figure 6: Compares the schematics of a TSTO before and after the completion of the first stage.
Now break the rocket into two equal stages, with each stage containing half the propellant and half the structure of Part 1 (i.e., 50,000 kg of propellant and 5000 kg of structure in each stage). The 5000 kg of payload goes on the front of the second stage.
Calculate the ΔV of each individual stage and sum these together to find the overall ΔV, comparing this to the ΔV value you found for the SSTO example.
In Figure 6, M01 shows the launch schematic of the TSTO whilst M02 shows the schematic of the second stage onwards.
For the first stage, our initial mass is the same 115,000kg we found in the previous section, but our final mass loses 50,000kg of propellant. We haven’t dropped the structure (MS1) at this stage as the calculations for the first stage assume we’re in beteen the two stages, but before the first stage has actually detached. Substituting these values back into the Rocket Equation, we find…
The calculations for the second stage assume that detachment has now happened, so the initial mass at the second stage is just the payload, the second stage structure and second stage propellant (as the diagram shows).
If we sum the two velocity changes together, we find total ΔV to be 10429 m/s. This is 1,437 m/s higher than the SSTO calculation. Thus, this worked example proves that a TSTO concept creates a higher delta V than an SSTO concept – or that TSTO concepts have greater performance – hence explaining their ubiquity.
Questions adapted from 
The SSTO concept is inherently built on the concepts of cost-effectiveness, reliability and reusability (although reusable TSTOs like SpaceX’s Falcon 9 are emerging) – which are all positive characteristics for a concept to have . Moreover, recent technological advantages have reduced the GLOW (Gross Lift-Off Weight) disparity between SSTOs and TSTOs, with composite material advances leading to significant SSTO GLOW losses in the last 60 years (comparable SSTOs would’ve weighed 2.72 million kgs in 1960 but 0.91 million kgs in the mid-1990s). However, given that the fundamental TSTO design has stayed the same, TSTOs haven’t seen the same level of GLOW reduction. SSTO’s only require the construction of one airframe, thus can be up to 30% cheaper than comparable TSTOs but also more reliable. Stage separation is the 2nd most common failure point for terrestrial launch vehicles thus by removing this factor, overall risk is greatly reduced . Thus, the SSTO concept has promise, but only if the drawbacks of the concept can be addressed:
- Weight (GLOW) – The average SSTO spacecraft has a 75% higher GLOW than a comparable TSTO (as it needs to carry more propellant due to the lack of stage separation), hence continued advances in composite materials are needed to further reduce the GLOW. Moreover, this also necessitates fuels needing to produce a high thrust – to overcome this high weight – so that sufficient resultant force is produced (yielding sufficient acceleration through Newton’s Second Law).
- Payload mass – Due to their weight, SSTO systems require more thrust than TSTOs, thus must carry more propellant thus SSTOs have a high PMF. Typical SSTO PMFs range around the 90% mark – which significantly hinders the payload mass, to the point where SSTO concepts become extremely marginal on payload – dependent on shaving every last kilogram. By weighing more than TSTOs, SSTOs are subject to greater structural loads upon ascent and thus require a more robust structure. Achieving sufficient structural rigidity comes at the expense of reducing PMF and thus delta V. Increasing fuel Isp can compensate for a reduction in initial mass to final mass ratio, thus allowing PMF to decrease, thus buying more payload mass without compromising delta V. Finding a sufficiently high enough Isp fuel to do so is arguably the biggest obstacle to the feasibility of future reusable SSTOs.
Thus, the feasibility of a future SSTO concept relies upon minimising structural mass and maximising the specific impulse of the fuel used. These allow payload mass to be increased without decreasing the initial mass to final mass ratio or (thus) delta V.
The next two sections look at the ideas of hypergolicity and air-breathing engines: two possible ways in which the drawbacks of the SSTO concept can be rectified.
Hypergolic ignition is where two chemicals react spontaneously with each other without any external stimuli – like the reaction between kerosene and concentrated hydrogen peroxide. A hypergolic system doesn’t require ignition due to the spontaneity of the chemical reaction, thereby saving weight by removing the need for an ignition system. In typical ignition systems, a spark is needed to begin a reaction whilst all a hypergolic system needs is a valve that lets the two substances mix to initiate combustion. This reduces GLOW and risk of failure, as well as allowing for in-flight restartability (the valve can be shut to prevent the two chemicals from reacting, ceasing the reaction process). These attributes make hypergolic propulsion systems attractive for a future SSTO. Hypergolic systems require a number of different characteristics to ensure their feasibility in a future SSTO:
- Ignition Delay – The ignition delay (the time between the reactants being in contact with each other and the combustion reaction starting) should be short (10 milliseconds). If the ignition delay is too long, hard starts can occur – where the two substances overmix, leading to a sudden spike in energy (and pressure) release, potentially causing a rapid explosion of the combustion chamber (referred to as “Spontaneous Self-Disassembly”).
- Storability – Many hypergolic chemicals tend to be stored for a long time and undergo a range of temperature changes hence chemical stability and storability is required in a hypergolic substance
- Toxicity and Reactivity – To minimise the fabrication and handling cost of the hypergolic substances, they should likely be as unreactive and harmless as possible – to reduce the impact of the substance’s use on humans. Consequently, Fluorine isn’t used despite its high performance as it is too toxic and reactive
- High performance – Every propulsion system needs to be high performance, no matter the application, thus maximising Isp remains desirable.
Hypergolic hybrid systems have high regression rates – a huge incentive for using hypergolics in a hybrid system as they effectively fix one of the (aforementioned) drawbacks of hybrid engines. The regression rate can be increased further by gelation of the propellant grain and the addition of metal hydrides – which increases regression rate by 60-400% relative to traditional hybrid rockets as well as increasing specific impulse. Increasing the regression rate removes the need to increase the surface area of the propellant grain (decreasing weight) and reduces the mass of propellant required. Some hypergolic hybrid combinations have an even greater regression rate than paraffin – which is known for its high regression rates. Thus, hypergolic hybrids can mitigate the low regression rates experienced in conventional hybrid systems.
Green hypergolic systems are those which have reduced impact on people and the environment without compromising on performance. Green fuels reduce the hazard risks associated in fabrication, handling, transport, storage and eventual disposal of the fuel – thereby reducing the associated costs involved in each of these steps.
FLP-106 is a green fuel developed by the Swedish Defence Agency (FOI) with a specific impulse of 381s, more than most hybrids . FLP-106 is neither toxic nor carcinogenic nor volatile yet has greater performance than Hydrazine (the most common hypergolic fuel). FLP-106 serves as testament to the idea that the use of green hypergolics need not come at the expense of performance and may potentially reduce launch programme costs.
Thus, green hypergolic systems have the potential to rectify the drawbacks of both hybrid and hypergolic concepts, without compromising on performance – making them a viable future propulsion option.
|Advantages of air-breathing engines||Disadvantages of air-breathing engines|
|Greatly improved specific impulse (from 450s to nearly 3500s)||Engine Performance is independent of speed and altitude (altitude ceiling due to the vacuum of space)|
|Weight savings (no oxidiser is required) which can then be reallocated to reinforce the structure and increase payload capacity||Thrust generation decreases with altitude|
|Reduced Gross Lift off Weight (GLOW) – ESA hypothesise that a SABRE-like engine could halve the GLOW of modern spacecraft through weight savings||Propulsive efficiency cannot reach 100% whilst rocket propulsion efficiency can|
|Reduced mission cost and increased usability – a consequence of the weight savings and reduced GLOW mentioned above. This could lead to future aircraft-like deployment from runways.||Air-breathing engines cannot be used on their own for space-craft and must be combined with another propulsion system (this need not be chemical, however)|
An air-breathing engine is very similar to a jet engine in that it uses inlets or ducts to suck in the air before using the oxygen from the air as oxidiser to burn the fuel – removing the need to carry oxidiser thereby saving weight. The table above outlines the main advantages and disadvantages of air-breathing systems.
Evidently, the main advantages with air-breathing engines are the massive increase in specific impulse and the potential to save weight – rectifying the two main drawbacks of the SSTO concept, as discussed earlier. However, spacecraft can’t rely on a solely air-breathing propulsion system as this would necessitate maximising time spent in the Earth’s atmosphere – the opposite of what a spacecraft wants to do.
Therefore, some concepts combine air-breathing systems with a conventional rocket engine – one such concept being the Synergetic Air Breathing Rocket Engine (SABRE) by Reaction Engines (British aerospace manufacturer).
The SABRE engine has two propulsion modes (see Figures 7a and 7b): it can switch to and from air-breathing mode and rocket mode. SABRE uses the air-breathing mode for the initial ascent, taking advantage of this mode’s specific impulse benefits to propel the rocket up to a height of 28.5km. At this point, the system switches to its liquid rocket engine for the rest of the ascent. The SABRE concept combines the high Isp and low weight of air-breathing engines with the high-altitude performance of rocket engines, thus has the potential to ameliorate the inherent compromises of current SSTO concepts . SABRE also serves as testament to the potential of bi-mode propulsion systems where both air-breathing and rocket engines work in tandem. Thus, although SABRE is still in the development phase (due to insufficient funding), it exemplifies the type of propulsion system which could finally make the SSTO concept viable. It remains to be seen whether SABRE can evolve to become mainstream rather than just a prototype as it is now.
Figure 7a and 7b: Compares the two different propulsion systems used in the SABRE engine. Figure 7a (top) shows the air-breathing propulsion schematic whilst 7b (bottom) shows the conventional rocket engine schematic.
The following section puts into practice the ideas discussed in the previous 5 sections to theorise a propulsion concept for a future SSTO.
Like Reaction Engines’ SABRE, this concept utilises air-breathing propulsion for the initial ascent up to a threshold altitude. However, at this threshold altitude, rather than switching to a liquid rocket engine, the system switches to a green hypergolic hybrid rocket engine.
The chosen propellant for such a system would ideally react non-hypergolically with atmospheric oxygen but hypergolically with a chosen oxidiser. This means that rather than having two separate engines (one air-breathing and one rocket), one can have a single propellant which can react with two separate oxidisers. An example of a propellant fulfilling this is kerosene: kerosene can react with atmospheric oxygen (non-hypergolically) but also with concentrated hydrogen peroxide (hypergolically) and thus a separate propellant isn’t necessary. Thus, once the spacecraft reaches a threshold altitude, the air-breathing system ceases and the hypergolic oxidiser starts mixing with the chosen propellant. By blending the two systems together, this concept saves weight by reducing the number of required – critical for any spacecraft concept but particularly for SSTOs.
Air-breathing systems are a requisite for a future SSTO concept as only they can produce the necessary specific impulse to permit an increase payload mass and overcome the high GLOW. Moreover, the high Isp reduces propellant consumption rate – meaning less propellant is required, saving weight.
As air-breathing systems cannot produce sufficient thrust above a certain height, they must be ‘paired’ with a rocket engine. Hybrid engines were chosen on the grounds of low cost and low complexity whilst maintaining the ability to be restarted and throttled (that solids don’t have). The reduced complexity again saves weight – even at the cost of reduced specific impulse compared to liquids. However, the Isp issue can be ameliorated by using a doped propellant gel – increasing Isp and fuel density.
Hypergolic hybrid systems increase regression rate (a factor plaguing hybrid concepts) whilst removing the need for an ignition system (reducing weight). The use of a green hypergolic system reduces cost and toxicity whilst preserving propellant performance.
This concept has two main aims: reduce weight and increase performance – hence each part of the concept has been thought out with these aims in mind. The weight saving measures this concept brings about enables the available payload capacity to increase whilst the increase in Isp and regression rate maximise the efficiency of the propulsion system to save costs – facilitating the idea of reusability.
Continued increases in technology (like the invention of composites which made the SSTO dream conceivable) will further improve rocket performance – which only increases the feasibility of such a propulsion concept’s use in the future.
This paper conceptualises a new propulsion system by analysing the requirements and drawbacks of different propulsion and staging types. This exploration has subsequently been applied to formulate a propulsion concept that has the potential to deliver performance-wise, whilst saving weight and being environmentally friendly.
This article explores relatively unconventional approaches to rocket design: the idea of integrated air-breathing hypergolic hybrid SSTOs isn’t as commonplace as the ‘standard’ multi-stage liquid rocket. This article also advocates the use of green fuels, which will only become more pertinent as environmental awareness heightens; the need to minimise environmental impact of spacecraft (from their creation to their launches to their disposal) and the need to create reusable rockets will be a central aspect of future spaceflight.
Only time will tell whether SSTOs ever become a viable means of space transportation but what is known is that if we can develop a propulsion system that can overturn the drawbacks of specific impulse and weight, the SSTO’s inherent benefits of reduced cost and complexity could increase the accessibility of space travel to the general population – thereby realising the dreams of space pioneers like Elon Musk or Konstantin Tsiolkovsky.
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Figure 1: Brian Cantwell, Graph of maximum velocity against initial mass to final mass ratio, Aircraft and Rocket Propulsion, 2010 https://web.stanford.edu/~cantwell/AA283_Course_Material/AA283_Lectures/AA283_Chapter_08_Multistage_Rockets_Brian_Cantwell.pdf
Figure 3: NASA, Space Shuttle Atlantis taking off, 1988, https://commons.wikimedia.org/wiki/File:Atlantis_taking_off_on_STS-27.jpg
Figure 4: University of Waikato – Science Learning Hub, Diagram of Hybrid propulsion system, 2011, https://www.sciencelearn.org.nz/images/412-hybrid-rocket-engine
Figure 5: European Space Agency, Soyuz FG Infographic, 2018, https://www.esa.int/ESA_Multimedia/Images/2018/06/Soyuz_FG_rocket_infographic_and_liftoff_sequence
Figure 6: Brian Cantwell, Launch schematics, Aircraft and Rocket Propulsion, 2010 https://web.stanford.edu/~cantwell/AA283_Course_Material/AA283_Lectures/AA283_Chapter_08_Multistage_Rockets_Brian_Cantwell.pdf
Figure 7a and 7b: Graham Parrish, Rocket engine will need funds to lift off, 2016, https://www.ft.com/content/9b1244f6-78f7-11e6-97ae-647294649b28
About the Author
Ayushman is a student in Year 13 studying Maths, Further Maths, Physics and Chemistry. He’s very interested in engineering but more specifically the spacecraft engineering aspect of the field and wishes to pursue this in the future