skylon space plane - final

8/10/2019 Skylon Space Plane - Final

1/27

SKYLON Space Plane

Date of Submission: April 1, 2014

by

_______________________________________

Shelby [email protected]

_______________________________________

Kevin [email protected]

_______________________________________

Eric [email protected]

Submitted to: Dr. Michael FabianDepartment of Aerospace and Mechanical Engineering

College of Engineering

In Partial FulfillmentOf the Requirements

OfAE 495S

Advanced Space PropulsionSpring 2014

Embry-Riddle Aeronautical UniversityPrescott, Arizona


2/27

AE 495S SKYLON Space Plane 2

1.0 Introduction

The concept of humans traveling beyond the Earths atmosphere was first realized in 1961 with the infamous flightof Yuri Gagarin, otherwise known as the first person in space. Since then, people have had a profound interest inmapping and traveling the stars. Leaving the safe haven that is Earth, however, is no simple feat, requiring countlesshours of research, development, and testing to even get near a launch platform. Costing upwards of $10,000 to

launch just one pound of cargo, however, space travel is an expensive venture. Therefore, the key component in thedevelopment of new spacecraft and space technologies is reducing the weight requirement, thus reducing the cost.

Since the birth of space travel, companies have been striving to develop a cost effective way to launch people andcargo into space. Many known and reliable launch vehicles can be used for only one launch due to the harshenvironment of launch and the design for demise that many of the vehicle components contain. With the exceptionof the Space Shuttle, every launch vehicle in employment uses a single-use, multi-stage rocket system. This meansthat during flight, conventional rockets consume fuel and then discard the empty fuel stages. Because of this, a large

portion of conventional rockets are destroyed after one use, a practice that is not cost effective.

The Space Shuttle, developed by the National Aeronautics and Space Administration (NASA), was first launched in1981. The Space Shuttle was a revolutionary new form of space transportation because of its partial reusability.While the Shuttle itself was reusable, there was a lot of cost and time required to manufacture new fuel, completeinspections, and rebuild certain components for each launch. As a result of such high costs, space became thedomain of government-funded organizations. In order to continue developing space technology and space travel,companies must now find new and innovative ways reduce cost and vehicle turn-around time. Reaction EnginesLimited (REL) believes they have a solution for this new and improved space vehicle.

Reaction Engines Limited is a private company founded in the United Kingdom in 1989 with the sole purpose ofdeveloping air-breathing rocket technology. Founded by Alan Bond, Richard Varvill, and John Scott-Scott, RELcontains a burning belief that if humanity is to conquer space, it has to find a better way of getting there thansimply lighting the fuse on expensive fireworks [ 2].In order to do so, REL researches space propulsion systems inorder to aid in the development of the SKYLON space plane, which is currently in its early phases of development.

The SKYLON space plane is a concept for a single stage to orbit (SSTO) vehicle that is capable of traveling to LowEarth Orbit. Single stage to orbit refers to a reusable launch vehicle that does not jettison hardware throughoutlaunch, meaning the only loss to the system is exerted propellants [3]. Because a large percentage of current launch

vehicles are comprised of propellant, this greatly reduces not only the cost, but the size and weight requirements ofthe vehicle. The goal of SKYLON is to provide a new source of space transportation with an unpiloted and reusablesystem. As Alan Bond says, "Expendable rockets can never deliver a credible transportation system. It is just toolabor intensive to build a vehicle of that complexity and then throw it away after one flight. Therefore, SKYLONwill be cost effective and efficient, reducing both the cost to launch to orbit and the time it takes to prepare thevehicle to be used again. SKYLON will do this by replacing the traditional launch vehicle engines with modified jetengines similar to those on commercial aircraft and developing the system to have horizontal launch and landingcapabilities.

2.0 Background

Single stage to orbit vehicles are not unheard of in the space community. In fact, single stage to orbit vehicles have been in use for quite some time for travel on and around other planets. The infamous Apollo programs lunar

module was a single stage to orbit from the moon, and many of Russias Luna robotic spacecraft use single stage toorbit vehicles as well [3]. It is on Earth that these systems become more complicated. Because of the forces exertedon the vehicles during launch to orbit, developing a vehicle that can survive launch and reentry forces is a dauntingtask. A task, however, that has been tried a few times before. Many companies and governments have attempted tocreate single stage to orbit vehicles. Rockets that launch and land vertically, air-breathing scramjet vehicles, nuclearvehicles, and jet engine vehicles have been under development for the sole purpose of the single stage to orbitvehicle that can reduce cost and increase efficiency.

In the past, many organizations working to develop single stage to orbit vehicles have studied hydrogen and oxygenrocket engines [4].However, this propulsion type has a relatively low specific impulse. When used for air-breathing


3/27


engines, the Hydrogen/Oxygen propulsion system is more feasible, but still contains a low thrust to weight ratio andlimited Mach range [5]. As a result, air-breathing engines alone cannot propel a system into orbit. A combination ofair-breathing engine and rocket engine, however, may be the most realistic way to get to space. The SKYLON space

plane is making strides in the movement towards a working single stage to orbit vehicle utilizing this enginecombination. The minds behind REL and SKYLON, however, have been working for decades to perfect the idea.

Before SKYLON came into development, the founders of REL worked on a similar project called HOTOL.Standing for Horizontal Take Off and Landing, HOTOL was Alan Bonds first air -breathing space plane concept.Development of the concept began in 1982 when the British government gave funding to a joint design teamcomprised of Rolls-Royce and British Aerospace engineers. John Scott-Scott, one of the founders of REL, was oneof the engineers from British Aerospace. HOTOL utilized a Rolls-Royce RB545 engine that used an air, liquidhydrogen, and liquid oxygen propellant system. In flight, this plane would take off and land from a runway similarto a commercial airliner. When the plane reached between Mach 5.0 and Mach 6.0, however, it would transition to

pure rocket propulsion [6]. The HOTOL space plane was to be 63 meters in length, with a seven meter diameterfuselage and a 28.3 meter wingspan. During design, the engines and liquid oxygen tanks for the space plane were

placed at the rear of the fuselage. The cargo bay and Hydrogen tanks were placed in the forward portion of thefuselage. Figure 2.1: HOTOL Design shows the exterior design of the plane, with the air-breathing enginesnoticeable at the rear underbelly of the vehicle.

During testing it was determined that this configuration caused severe malfunctions during ascension. During this

portion of flight, the center of pressure shifted ten meters forward due to a variety of factors including a largefuselage cross section and small wing cross section ratio, a large forward fuselage overhang, and a large Machrange. As a result of this pressure shift, the center of gravity shifted to the rear end of the plane. These shifts in

pressure and center of gravity would make the plane unstable during flight if not resolved [3]. As a result, theHOTOL space plane was redesigned several times in order to remedy the center of pressure and center of gravi tyshifts. However, there was little success and the government withdrew its funding in 1988. Unfortunately, the designfor HOTOL was almost complete. However, the aerodynamic problems and speculative nature within the project

proved to be its downfall.

With the loss of government funding, Alan Bond and John Scott-Scott developed REL to continue the space plane project. British Aerospace did offer a HOTOL 2 concept in 1991. The HOTOL 2 redesign would utilize a moreconventional Liquid Hydrogen and Liquid Oxygen engine and be launched from the back of a modified aircraft.This concept, however, was never developed further. This left REL as the only company exploring the continuationof the HOTOL air-breathing space plane concept. After its development in 1989 , RELs goal was to fix the

problems of HOTOL and develop a better mode of space transportation for the world.

3.0 Concept

If completed, the SKYLON space plane could provide humans with a fast, reusable, and reliable means of gettinginto space. At its core, SKYLON is a [sleek redesign of the HOTOL project. Utilizing a lengthened, slenderfuselage, the SKYLON body is 83.3 meters in length and 6.75 meters in diameter. While SKYLON has a similarfuselage diameter to HOTOL, it is considerably longer. Just as with the fuselage diameter, the wingspan for HOTOLand SKYLON are similar in length, with SKYLON containing a 25.4 meter wingspan. The wing section ofSKYLON, however, was moved forward on the plane. To reduce the lift problems of HOTOL, SKYLON s w ingswere moved forward on the fuselage to near the midpoint of the length and in line with the payload bay. The engineswere also moved in order to reduce the massive moments that HOTOL experienced during flight. The overall

orientation of SKYLON can be seen in Figure 3.1: SKYLON Space Plane. These changes were made in order tokeep the center of gravity near the center of the plane.

The orientation of the propellants and propellant tanks also contributed to some of HOTOLs instability. In order toavoid this, the Hydrogen and Oxygen tanks were cut in half and placed on the two sides of the payload bay, as seenin Figure 3.1. As a result, SKYLON is comprised of two liquid hydrogen tanks and two liquid oxygen tanks, ratherthan one of each. During the air- breathing portion of the planes ascent, the Hydrogen is burned from the rear tank.Because the center of pressure will still shift forward slightly during flight, burning from the back Hydrogen tankensures that the center of gravity moves forward as well. This change results in more control of the center of gravity,which means that the stability and usability of the plane increases [1].


4/27


Because SKYLON completes both ascent to orbit and decent back to Earth, it needs to be different than typicalorbital vehicles. In the past, re-entry vehicles have been blunt in shape. This was done in order to maximize the dragin the Earths atmosp here to slow the vehicle. Because a vehicle that produces minimum drag is needed for ascent,REL had to re-evaluate the reasoning behind the blunt re-entry vehicle in order to ensure that SKYLON cancomplete its mission. In doing so, they determined that the blunt vehicle is not the only option for re-entry. Re-entryvehicles are required to experience very high heats during atmospheric re-entry. As a result, re-entry vehicles must

be developed to withstand these heats. Previous re-entry vehicles were designed so that the underside of the vehiclewould experience the most heat and be able to withstand it through the use of extra heat protection. This extra

protection, though, adds to the systems overall mass, which is a characteristic SKYLON seeks to avoid. BecauseSKYLON utilizes a different body shape and does not contain the extra heat protection on the underside of thevehicle, there was concern that the structure would not be able to sufficiently withstand re-entry conditions. Uponfurther analysis, however, it was determined that the re-entry heats were quite manageable. With this considered, theREL engineers still needed to provide the vehicle with adequate and lightweight heat protection.

SKYLON was developed with three methods for managing the heats experienced by the structure. These include anaeroshell, layers of reflecting titanium foil, and water evaporation. SKYLON s structure is made up of a carbon fiberframe reinforced with plastic struts. The frame is then covered with reinforced glass ceramic sheets, which serve asthe aeroshell, or main form of heat protection. The aeroshell is then topped with many layers of a foil heat shield to

provide additional thermal protection. Within the frame, most of the fuselage is comprised of hydrogen tanks,making SKYLON considerably lighter and less dense during atmospheric reentry. This reduces not only how hot thevehicle gets but also many of the complications associated with very high heats. After thermal analysis, it wasdetermined that SKYLON would experience around one megawatt per square meter of heat transfer. However,standard jet engines and rocket engines could experience up to 100 times more heat transfer during flight. Thismakes SKYLON s heat concerns considerably easier to manage. It was determined that three points on the planewould experience concentrations of heat, as demonstrated by the red circles on Figure 3.2: Re-entry HeatConcentrations. In order to ensure that the heat experienced at these areas was not damaging, around 100 kilogramsof water could be drained over the exterior, which would evaporate and provide cooling.

On paper, SKYLON s design is plausible using current and available technology. However, some of thistechnology, including the air-breathing rocket engine, had not yet been perfected. Therefore, REL had to develop anengine system that could complete SKYLON s mission requirements. The result was the Synergetic Air-Breathingand Rocket Engine, or SABRE. Comprised of a combination of a hydrogen rocket and a jet engine, the SABREengine contains the best of a plane and a rocket in one. Over twenty years in the making, the SABRE engine

contains a pre-cooler and turbine that enables it to travel at up to five times the speed of sound. As seen in Figure3.3: SABRE Engine, the SABRE contains an air intake, a pre-cooler, a compressor, a turbine, and rocket nozzles [1].In air-breathing mode, the engine uses oxygen from the environment as oxidizer to propel the system. When turnedto rocket mode, at around a 26 kilometer altitude, the engine utilizes an on-board oxygen supply as oxidizer. Theoverall system uses sub-cooled liquid hydrogen as a fuel.

The SABRE engine contains many components that set it apart from other attempts to develop a successful air- breathing rocket engine. However, at its base, it is the mastery of the combined engine cycles that set SABRE up forsuccess. The air-breathing and rocket engines are combined by a series of common components. The ultimate goalof this combination is to maximize functionality in order to keep the system weight and installation drag to aminimum [5]. Because the air-breathing mode uses the environment around it for oxidizer rather than carrying it on

board, SABRE enables a total system mass reduction. Despite numerous attempts to develop these combinedengines, SABRE is the first of its kind that utilizes realistic components and current technology. Therefore, SABREis the only credible engine in its field . SABREs overall goal is to provide the SKYLON plane a reusable means ofgetting to and from space. Therefore, the success of the engine, and therefore SKYLON, could greatly reduce thecost of space transportation. However, in order to utilize the same combustion chamber as a rocket engine, the air-

breathing portion of the engine must be able to compress the air to the same pressures as the rocket cycle [5]. Inorder to compress the air to this pressure, it must be gathered and cooled.

During flight, the axisymmetric intake draws in air. Although the air entering the engine is quite fast, the air intakeslows the air, enabling it to move through the engine at relatively low Mach numbers. When the air is slowed,however, kinetic energy is released in the form of heat. Because the air is slowing so much and so quickly, the heat


5/27


in the engine skyrockets, reaching upwards of 1,000 degrees Celsius. As previously mentioned, however, in orderfor the engine to run, the air entering the combustion chamber must be cooled. This is where the pre-cooler steps in.

The minds in REL knew that the only way their engine could work was if they developed something that coulddrastically cool the air in a fraction of a second. If the oxygen were to be cooled to a liquid form, which is themethod of some air-breathing engine designs, then a high fuel flow would be required in order operate the engine. Ahigh fuel flow results in more fuel consumption, which ultimately requires more fuel and added mass. Therefore, the

pre-cooler within the SABRE cools the air to its vapor point, or the point just before liquefaction. This reduces notonly the need for hydrogen fuel flow, but also the additional cooling requirement. Therefore, the pre-cooler forSABRE was developed to reduce the temperature from 1,000 degrees Celsius to minus 150 degrees Celsius in ahundredth of a second [1]. The need for this sort of cooling is a restricting element for most aircraft traveling at veryhigh speeds because of the structural damage such heats can cause. In addition, cooler air is easier to compress to ausable pressure for the rocket combustion chamber. It is how REL made their pre-cooler; however, that sets it apart.

Heat exchangers work by passing fluid or gas through tubes that are surrounded by another liquid or gas, resulting ina transfer of heat. Fins are commonly used in heat exchangers, too, to increase efficiency [1]. For the purposes ofSABRE, however, fins contain too much weight. Therefore, REL made the tubes smaller and developed a way tomake the walls of the tubes thinner. This meant that they could increase the amount of tubes in the system withoutadding too much weight. The benefit of smaller tubes is an increase in surface area, meaning that there is more areafor the heat transfer to take place. After much trial and error to manufacture the tubes required to make SABREwork, REL found out how to make one millimeter diameter tubes with walls thinner than a strand of human hair [1].The tubes were then mounted in a spiral arrangement, as seen in Figure 3.4: SABRE Pre-Cooler. This combinationof smaller tubes and a spiral arrangement allows for an immense amount of cooling in a smaller area.

Overall, the pre-cooler is part of a closed cycle helium loop. The on-board helium is cooled using the cold hydrogen.The helium is then put though the pre-cooler to cool the incoming air. After leaving the pre-cooler, the helium isfurther heated by the pre-burner product [1]. This heating of the helium gives it enough energy to then drive theturbine. Therefore, described as a loop, as the air is cooled, it heats the helium that powers the engine machinery,which uses the cooled air to power the turbine. In addition, during flight the air intake takes in more air than isneeded for the compressor. This excess air is passed around the pre-cooler and fed to the burner duct. This helps torecover drag losses [8].

It is this cooling combined with the high compression ratio of the air that enables the air-breathing component of

SABRE to achieve a higher thrust to weight ratio. In order to accomplish a high specific impulse there must be a lowdemand for fuel flow [5]. This requires an advanced thermodynamic cycle with a range of heat exchangers. Inaddition to the pre-cooler, a heat exchanger with the purpose of cooling hot helium with the cold liquid hydrogenon-board is used. The combination of the pre-cooler and the more conventional heat exchangers keep the SABRE ina working state, allowing the engine to perform its mission.

Together, SABRE and the SKYLON space plane are a realistic solution to the need for a more cost effective way toget to space. Once in production, the SKYLON space plane will revolutionize the space industry. In 2012, RELcompleted a successful scaled test of their revolutionary pre-cooler, putting them on the forefront of the technology.Tested in a Viper jet engine, the heat exchanging system was successfully tested in a working environment, provingthe concept worked. Since then, REL continues to develop and test their technology. With SKYLON becomingmore plausible, more thought has gone into its mission. When in use, SKYLON will be operated as an orbitaltransport. Therefore, its large and adaptable payload bay was made to accomplish many missions. The design for the

payload bay was determined using existing payload requirements. Thirteen meters long and five meters in diameter,the payload bay is a U shaped area with two opening doors. The orientation of the SKYLON payload bay can beseen in Figure 3.5: Payload Bay. Able to accommodate over 33,000 pounds, the bay can carry both people andcargo. As previously stated, the payload bay is placed almost midway in the fuselage and over the wing section. The

bay contains two payload connections points on each end, and, if payload mass constraints are met, both can be usedsimultaneously for one large or two smaller payloads [1].

Because of the shape and size of the payload bay, it can accommodate cargo, multiple satellites, and humantransportation. This makes SKYLON an efficient and useful system. In addition to the two connection pointsdiscussed, the payload bay can also be fitted with mounts for multiple satellites. Although SKYLON is meant for


6/27


Low Earth Orbit transportation, the plane can also be used for satellites and systems requiring geostationary orbits.This can be done by attaching these satellites to a second stage booster, which would be used for orbital transfer. Itis also possible to reuse these upper stages by collecting them after use. In addition to satellite and cargodeliverance, SKYLON could also be used to build infrastructure in space.

4.0 Key Physics and Engineering

The SABRE engine's Combined Cycle is a dual mode engine that, at take-off, is a pre-cooled air-breathing turbo-ramjet and a closed cycle rocket engine after reaching Mach 5. Utilizing the same hardware for both engine cycleskeeps system weight and drag from extra engines or parts down to a minimum. Using the same hardware for bothcycles means that the air must be compressed to similar pressures as in a rocket cycle though. During the first halfof flight, the SABRE engine in air-breathing mode will intake about 1250 tons of air. 260 tons of this air is usableoxygen that the space plane will not have to carry throughout its' ascent [2]. Unlike the LACE, the oxygen in theSABRE's cycle will remain in a gaseous state. Alan Bond and Reaction Engines Ltd. did this to fix the LACE's highfuel flow problems.

4.1 Combined Cycle Rocket Engine

In a normal turbo-ramjet, the air passes through an inlet and is compressed. The compressor is driven by a gas

generator fed by on-board fuel and oxidizer. In the SABRE engine though, pre-cools the air with a light-weight pre-cooler made of consisting of many thousands of small bore thin wall tubes. The pre-cooler utilizes a closed Braytoncycle power loop of helium as seen in Figure 4.1: SABRE cycle [9].

The helium is cooled at HX4 by on-board liquid hydrogen which acts as a heat sink. The cooled helium is thencompressed, then extracts the heat from the incoming air via the pre-cooler. After its temperature is stabilized inHX3, the helium turns the turbine that drives the turbo-compressor as well as turning the turbine that circulates thehelium. The turbo-compressor compresses the air to the appropriate pressure for use in the engine's combustionchamber. Therefore in this cycle, the major heat exchanger roles are: "extracting heat from incoming air in the pre-cooler, topping up cycle flow temperatures to maintain constant turbine operating conditions and extracting rejectedheat from the power cycle via regenerator loops for thermal capacity matching" [5].

The intermediate fluid helium was interposed between the incoming air flow and on-board hydrogen heat sink to

prevent hydrogen embrittlement of materials, as well as acting as a safety barrier between fuel and oxidizer due toits inertness and specific thermal capacity. Since the specific thermal capacity of helium is in between air andhydrogen's, capacity matched heat exchangers were easier to design. In addition to these reasons, helium's highspecific heat ratio "reduces the cycle pressure ratios by a factor of two keeping the size of ducting small" [5].

Figure 4.1: SABRE cycle conditions were found using a Mach 5 design case and an isentropic analysis coupled withrealistic performance for each part. Isentropic conditions yields the maximum air compression ratio possible, i.e.,the overall entropy rise of the isolated system is zero. The realistic performance calculations coupled into theanalysis included entropy generation due to a finite driving temperature difference, or T, irreversibilities due to

pressure loss and friction, and associated polytropic efficiencies for the compressors and turbines. A polytropicefficiency is the ratio between the polytropic work to the actual work done by a component. These performancecalculations contribute to entropy rises and thus "can be considered as lost capacity to do work" [5].

For a compressor, entropy rise per unit mass flow is found using the following equation:

2c

1

ln (1 ) pS T

cm T

[5]

For a turbine, entropy rise per unit mass flow is found using the following equation:


7/27


2

1

1ln T p

T

S T c

m T

[5]

Using the 2 previous equations, Table 4.1: Engine Component Attributes was constructed, and lists the realistic performance attributes for the engine s components. The entropy of each unit was evaluated using isentropic

conditions and the efficiencies from Table 4.1 applied. Combining these, the total entropy rise of the combined airand hydrogen streams are found.

The compression ratio was then found using the following equation:

212

exp exp1 1

air

in

air out air real ideal

air H in air air p pair air

S T S PR PR

T T mc mc

[5]

The analysis done by Reaction Engines Ltd. yielded a pressure ratio of 3330 under isentropic conditions and 202 forthe SABRE engine under realistic conditions. This is a marked improvement over other jet engines like those used

by Boeing or Airbus which have pressure ratios around 30-40.

At Mach 5.14 and 28.5 km altitude, the intake closes and the SABRE engine uses on-board liquid Oxygen as itsoxidizer for the liquid bi-propellant system. Since the air-breathing mode takes the SKYLON vehicle to this highspeed, the second half of the mission requires less mission velocity. The majority of the V is achieved in this halfthough, as the SKYLON vehicle reaches orbital velocity in this half. Though this is true, the rocket cycle must still

be as efficient as possible to make it to orbital speeds.

4.2 Pre-Cooler

The key enabling technology of the SABRE is the pre-cooler. The rate of heat transfer, , per unit heat transfer area,A, has the following equation:

QU T

A

where T is the mean temperature difference between the two fluids exchanging heat, and U is the overall heattransfer coefficient which incorporates air flow thermal resistance, helium flow thermal resistance, and theresistance of the separating solid surface [10].

One way to maximize the rate of heat transfer per unit heat transfer area is to increase the fluid flow velocity. Thiscould result in an overall negative effect on performance through the increase in frictional power loss though.Instead a very large heat transfer surface area is necessary. Since tube diameter is inversely proportional to the fluidheat transfer coefficient, reducing the diameter of the flow passages allows for a compact high surface area densityand improved heat transfer per unit area.

In the SABRE pre-cooler, many thousand small bore tubes are brazed together to form a matrix. The matrices are

modularized, wrapped into a cylindrical drum, and arranged into a counterflow design with the hot air in an externalcrossflow.

The matrix of tubes seen in Figure 4.2: Pre-cooler was chosen to give the lightest matrix due to the high pressuredifference between the two fluid streams. Fins and other extended heat transfer surfaces were not used due to theirhigh weight to surface area ratio. The counterflow of helium minimizes the finite driving temperature differencewithin the pre-cooler. The modular design wrapped into a cylindrical drum with the hot air in an external crossflowyields a large frontal area that prevents excessive engine nacelle wave drag by reducing flow constriction into theturbo-compressor, as well as automatically forcing new boundary layers to grow on each tube [10].


8/27


Development of the pre-cooler started with a small experimental pre-cooler in the 1990's called the JMHX [11].From there, a pre-cooler was made in 2001 to test frost control in a wind tunnel. After this came a single Viper jetengine module in 2004, 2 SCIMITAR modules during LAPCAT in 2008, a prototype production line of multipleSABRE pre-cooler modules, and a full-scale Viper jet engine pre-cooler in 2012.

The JMHX was a small high pressure staggered tube matrix heat exchanger researched and manufactured at BristolUniversity in the 1990's. Development tested brazing techniques and materials to bind the small tubes together, aswell as the elements used for the intermediate interposed stream. The JMHX had 0.38 mm diameter stainless steeltubes with 50 m walls. Its layout of 10 layers alternating between 41 and 42 tubes resulted in 415 tubes being

brazed together to form a 40mm by 40mm wafer that was 4mm thick. This thin wafer can be seen in Figure 4.3: TheJMHX [11].

Brazing methods investigated included brazing temperature, brazing material, and application to industrial furnaces.The tubes were brazed at 910C instead of the normal 1,000C brazing temperature due to erosion of the tubesoccurring at high temperatures. The material used was a nickel phosphorus eutectic alloy manufactured byColomony as Nicrobraz 10. The method used to braze the tubes together was an electro-less nickel platingtechnique that catalytically deposited a nickel phosphorus braze alloy onto the surface to be brazed. This woulddeposit constant thickness of material with even distribution due to invariance in electric field strength over complexsurfaces like anodic electroplating would. From the research, it was found that normal vacuum furnace must beoutfitted with a molecular oxygen getter to reduce partial pressure of oxygen by 100 times below standard vacuumatmosphere to reduce chromium oxide from forming during production process.

The JMHX was built for temperatures of 100 K N 2(g) or He(g) to cool 1,000 K N 2(g). The nitrogen had a maximumchange in temperature of over 200 K while the helium had a maximum change in temperature of over 500 K. Theuse of helium resulted in about 1.8 GW/m^3 in power transfer rates, while nitrogen only resulted in about 0.88GW/m^3 [11].

The Viper engine demonstration module had 0.98 mm diameter tubes with 40 m walls. Its tube layout of 460 tubesthat were 2.2 m long resulting in about 1100 m of tubing being drawn. This module investigated assembly andmanufacturing issues. The 2 full scale SCIMITAR engine modules made during the LAPCAT - Long-termAdvanced Propulsion Concepts and Technologies. These modu les used 0.88mm diameter tubes with 40m wallthickness, the same size bores that the SABRE engine would use. These modules were used to investigate brazingunder industrial process conditions.

The 9% SABRE scale Viper pre-cooler had to scale tubes, but 9% as many as the SABRE would have. It operatedat same the temperatures, pressures, mass fluxes, and Reynolds numbers as the full scale SABRE would. The testmodule was about 50 kg but contained about 50 km of heat exchanging tubing. The tests that occurred in 2012investigated aerodynamic stability and uniformity, structural integrity, freedom of vibration across a wide range ofthe flight envelope, and cryogenic cooling [12].

5.0 Potential Uses

5.1 LAPCAT A2

The LAPCAT A2, seen in Figure 5.1: LAPCAT A2, or Long-term Advanced Propulsion Concepts and Technology,is the entirely air-breathing application of the SABRE engine and lightweight heat exchanger technology, optimizedfor atmospheric flight only. The LAPCAT A2 would utilize the Scimitar engine, seen in Figure 5.2: ScimitarEngine. The four Scimitar engines would propel the LAPCAT to cruising speeds of up to Mach 5 using liquidhydrogen instead of normal hydrocarbon fuels. With a cruising speed of Mach 5, travel time would be reduced in"long-distance flights, e.g. From Brussels to Sydney, to less than 2 to 4 hours" [2]. The layout of the engine would

be similar to the SABRE in that it would have a pre-cooler, but would also contain "a high bypass airflow permittingefficient subsonic flight and moderate take-off noise" [2]. The Scimitar engine would be designed for a much longerlife than the SABRE though, since the LAPCAT would need to sustain Mach 5 cruise.


9/27


The A2 vehicle configuration proposed by REL allows necessary subsonic and supersonic lift/drag ratio for efficientcommercial application. It is "designed to have adequate control authority about all axes to handle engine-out and toachieve pitch trim over the full Mach range" [2]. The LAPCAT would be sized to carry 300 passengersapproximately 20,000 km in both subsonic and supersonic flight. This would allow the vehicle to service a largenumber of routes and avoid supersonic flight over populated areas.

5.2 Orbital Base Station An orbital base station (OBS) has been conceptualized since the 1950's and would enable "assembly andmaintenance of a Cis-Lunar transportation infrastructure and integration of vehicles for other high energy spacemissions" [13]. The OBS would be modular in nature and be assembled using multiple launches of the SKYLONspace plane. The facilities of the OBS put forth by REL, seen in Figure 5.3: Orbital Base Station, would incorporatean assembly dock, longitudinal rails to provide internal tether attachments, manipulators to handle and assemblevehicle structures, and habitation modules.

Since the OBS would be dependent on the SKYLON launch vehicle, the payload capabilities of the vehicle are alimiting factor. This would mean the station would need to be modularized for accommodation in the payload bay,taken into orbit, and assembled piece by piece. Advantages of doing this include "standardized units that arecompletely reusable and replaceable with the aim of lowering the cost of space operations by means of minimalhardware development programs" [13].

Assembly of a space transport system inside the proposed dock, which would be "a large cylindrical space-framestructure with two large doors on either end incorporating a skin of aluminized Mylar" [13], would utilize standardtankage modules. These modules could be stacked in parallel to form a transfer stage. The amount of modules usedwould be dependent on the given mission but could look like Figure 5.4: Modular Construction of a Vehicle Stage.A SKYLON vehicle would dock using an exterior manipulator arm to engage the docking mechanism located in theundercarriage wells instead of performing a hard docking maneuver. The payload would then be transferred insideusing the arm through one of the transfer doors into the interior.

5.3 Project Troy

Project Troy is a Mars mission concept created by REL to determine the role that the SKYLON space plane take inthe exploration of Mars. The size and mass of components of a credible Mars mission were therefore estimated andcompared to performance capabilities of the SKYLON. The mission architecture selected included maximumcoverage of Mars, safety of the crew, and minimization of cost. This yielded a mission requiring 2,300 tons ofhardware placed into orbit for each mission phase. The mission would span 14 months on the Martian surface with18 explorers covering 90% of the surface. The mission "would have surface and orbital resources enabling extendedstop-over to wait relief should major equipment failures occur" [14].

The two phase mission includes an automated phase and a manned phase. The first phase would deliver equipmentlike surface habitats and power supplies to the surface of Mars, two years prior to the manned phase. This wouldallow "a working surface base and orbital facilities to be established and checked out in advance so that an abortedexploration has the maximum chance of survival and a range of predetermined back-up options" [14].

The mission would require six Martian transport vehicles, three for each phase. Each Troy vehicle has three stages,an Earth Departure stage, Mars Transfer stage, and Earth Return stage which are shown in Figure 5.5: Troy SpaceVehicle. The vehicles would be assembled in an OBS and brought to orbit using the SKYLON space plane. ForPhase I, approximately 2346 tons would be required to lift to orbit, while Phase II would require approximately2234 tons. This would take about 522 total flights of the SKYLON space plane to bring all the equipment and fuelto orbit. Clearly, only the SKYLON would be able to achieve this with relatively low costs.


10/27


6.0 Limitations

6.1 SKYLON Limitations

Reaching the orbit of the Earth has always been a laborious task with many limiting factors. Weight of cargo andweight of the required propellant are the main limiting factor. A typical rocket requires around 90 percent of themass to be propellant. This immense fuel requirement consumes non-reusable fuel tanks which are ejected in stagesfor weight reduction and a more efficient burn. The jettison of each stage is a risky action and can cause damage tothe rocket if done incorrectly. Ever since the beginning of the space race engineers have searched for ways toimplement refillable rocket fuel tanks and avoid stage jettisons. The SSTO space plane has been a dream of manyengineers and is faced with not just weight limitations but also aerodynamic limitations, structural limitations,engine limitations, and limitations of a reusable fuel along with the storage system of this fuel. As it is with everynew rocket project there is a large research and development cost as well.

The fuel, engine, and development cost limitations for the SKYLON space plane are the main limiting factors. Theoverall manufacturing cost for each SKYLON is estimated at 12 to 14 billion USD [2]. This price is still just anestimation and could decrease as manufacturing techniques improve. Even though the manufacturing cost seemsunreasonable, the SKYLON would be able to decrease the conventional cost to reach orbit by 10 fold or greater. Theestimated cost for flight and payload delivery into orbit is about 35 million USD but could see costs as low as 10 or

even 2 million USD [2]. This may sound like a hefty budget but in reality it is small when compared to NASAlaunches which range in the 100s of millions of USD. Even new private rocket companies have payload deliverycost estimations of 50 to 90 million USD. The SKYLON s refillable fuel systems and SSTO capabilities allow forthe decreased costs but also present structural constraints to the SKYLON as well.

The SKYLON will take off from sea-level and fly to orbit. A large amount of stress will be developed on theSKYLON internal structure. This structure must be lightweight yet incredibly strong. The purposed material for thistask is unidirectional carbon-fiber-reinforced polymer (CFRP). The predicted ultimate stress of this CFRP layup is1,500 MPa, for compression and tension [15]. The SKYLON will use a circular cross-sectional body for simplermanufacturing techniques and for near ideal pressure vessel effects. The CFRP internal structure will consist of 4square-tube longerons at the 45 degree points of the circular cross-section. The longerons will provide bendingresistance for take-off and reentry. Ring frames provide a circumferential strength when subject to the changing

pressures. The ring frames will be spaced 300 mm. apart to provide maximum global rigidity as well as torsional

resistance when coupled with the 4 longerons. The aerodynamic control surfaces at the nose and tail sections requireincreased support structure do to the large Mach and reentry forces. An increase in the number of ring framescombined with monocoque nose and tail cones has been the purposed approach for handling the increased stresses.

The circular design will help idealize the pressurized storage of the liquid oxygen (LOX) and liquid hydrogen(LHX) fuels. The fuels required for the SKYLON s mission profile will consume majority of internal space of theSKYLON. The LHX fuel accounts for most of the internal space since it is burned for both air-breathing and rocketmodes of the SABRE engine. As the SKYLON escapes and reenters of the Earth s atmosphere fluctuating pressureswill stress both the support structure and fuel tanks of the SKYLON. If the internal fuel tanks collapse, do tonegative pressures, the whole body of the SKYLON will collapse. To avoid collapsing the LOX and LHX fuel tanksmust remain pressurized at all times.

The constant pressurization of the fuel tanks is also a constraint to avoid critical damage to the liquid fuel pumps; aswell as, to avoid the dew point temperature of nitrogen gas (77.35 K). If nitrogen condenses as the SKYLONreaches orbit, then upon reentry the nitrogen will rapidly expand and cause depressurization of the SKYLON space

plane. A 10mm. PVC foam insulation covering around the fuel tanks will help to combat this effect. The constrainton the LOX and LHX pumps come from the net-positive-suction-head limits (NPSH) [15]. NPSH limits govern thatthe pressure of the liquidous gas cannot rise above a pressure which will cause cavitation in the pumps. Cavitationwithin a pump can lead to intense bubble damage on the pump blades and pump failure. The NPSH limits allow forthe temperature of the LHX tanks to rise from 16 Kelvin to 18 Kelvin (K). This 2 K temperature rise is equal to anincrease of 3.6 degrees Fahrenheit. The subtle increase in temperature effectively raises the LHX temperature abovethe dew point of Nitrogen to avoid depressurization while also staying within NPSH limits. This temperature rise


11/27


also satisfies an FAA regulation which states that the SKYLON must be able to be delayed on tar mack for 2 hours.The 2 K temperature increase for the volume of LHX fuel would take longer than 2 hours.

The circular cross-section of the SKYLON will also help atmospheric drag. The elongated body and circular cross-section create a Sears-Haak body. The Sears-Haak body is created by reducing the diameter of the fuselage over thewing to make a classic coke-a-cola bottle like shape. The incorporation of the Sears-Haak body helps to reduceMach drag but a corrugated silicon-carbide (SiC) fiber reinforced glass ceramic Aeroshell will increase the skinfriction drag; however, the skin friction will only contribute a small amount of drag and not enough to be considereda major drawback.

The SiC corrugated shell allows for thermal expansion when the SKYLON reenters the Earths atmosphere. TheSiC shell is 25 percent more dense than C/SiC material and has a lower maximum temperature, but SiC is 100 timesless expensive to manufacture. The SiC material was chosen for its cost effectiveness. Do to the SKYLON s fuel

pressurization constraint, a dynamically controlled re-entry keeps aerodynamic heating temperatures below themaximum temperature of the SiC shell (1470 K). Temperature sensors on all the sharp points of the SKYLON willrelay re-entry control information. This dynamic control will be paired with convective gap to help maintain aconstant re-entry temperature of no more than 1100 K. The convective gap will be powered by the low temperatureof the LHX and LOX fuel tanks along with 10 titanium foils [15 ]. The titanium foils have a thickness of 10m andwill be spaced 3mm apart. The foils help create a more consistent temperature gradient between the inside of theSKYLON space plane and the exterior shell. The convective gap, through passive radiation cooling, will also keepthe skin at a predicted 855K during aerodynamic heating at Mach 5 conditions [15].

The high speeds required to reach escape velocity and the dual functionality of the SABRE engine are what drive thelarge amount of fuel needed for a mission. The gross takeoff weight (GTOW) of the SKYLON is 275,000 kg withthe payload being limited to 12,000 kg. The remaining 263,000 kg is divided into 220,000 kg of fuel, and 43,000 kgof structures and flight equipment [9][15]. These immense weights will limit the SKYLON s structure to a 0.48 gcornering maneuver constraint. The heavy weight of the SKYLON will also limit the internal structure to a 1.9 g

pull-up + 0.1 g gust maneuver constraint at Mach 5 flight conditions. This intense speed will cause large stresseswhen maneuvering with the large weights, but the large size and Mach speeds drive the gust conditions to lowervalues. Mach 5 is extremely fast when compared to a wind gust therefore the SKYLON feels minimal gust effects.The stress on the LHX tanks when fully pressurized is limited to 2 gs, which is the gust limit added to the pull -upmaneuver limit for the structural frame. If the tanks are at there depressurized pressure limits than the load limit forthe LHX tanks falls to 1.25 g.

The issues of center of gravity shifts for the fuel and payload have also been accounted for. The constant pressurization of the fuel tanks will keep the center of gravity of the fuel from shifting. The center of gravity shiftcaused by removal of the payload will be minimized by placing the payload center of gravity and total SKYLONcenter of gravity locations as close to the aerodynamic center of the wing as possible. This aerodynamic center,center of gravity placement will depress the movement of the overall center of gravity location when the payload isunloaded. The delta shape of the wing will provide favorable Mach aerodynamic center shifts. Also aiding theSKYLON with stability for both loaded and unload configurations is the curved shape of the SABRE engines. Thecurvature provides a stabilizing pitching moment about the aerodynamic center for all loaded, unloaded, subs-sonic,and super-sonic flight conditions. The staggering mission profile of the SKYLON space plane will limit it to amaximum of 200 flights before retirement. The intensity of these flights will also require that the SKYLON beserviced for 2 days after each flight.

6.2 SABRE Engine

The SABRE engine is a feat of engineering which will be one of the greatest aerospace developments in the 21 st century. The engine is capable of both an air-breathing mode and a rocket mode. This dual functionality requiresthat the SABRE engine achieves heat transfer rates which are 30 times great than previously thought possible formodern engines. The intake air will not only be entering the engine at a drastic speed but also a drastic temperature.


12/27


At Mach flight condition intake air must be slowed down to below Mach speeds. If it is not the engine will be blownout and all thrust will be lost. The SABRE engine will use a 2 shock-axisymmetric closable intake cone. The 2shock cone focuses 2 Mach waves into one another to create reflected shock waves which reflect to form an obliqueMach wave, thus slowing the super-sonic flow to sub-sonic flow conditions. The slowing of the flow essentialdiffuses the pressure and causes a large increase in temperature.

The SABRE engine intake cone is of variable geometry in the sense that it will be closable to form a smooth surfacewith no intake, when the SABRE engine transitions into rocket mode. This lowers the Mach wave drag penalty ofthe external engines for the remaining super-sonic rocket mode. However, the movable cone increases weight andmanufacturing cost. Motors must be added to control the opening and closing of the cone and will therefore addweight to the overall design. The CFRPs special shape for each concentric piece could very well cause an increasein manufacturing costs. Part reliability for the motors and moving cone parts could also develop issues for futuremaintenance costs.

The intake temperature after diffusing is predicted to be upwards of 1350 K at Mach 5 speeds. The 1350 K is the predicted maximum temperature for the intake. The SABRE engines pre-cooler must lower the 1350 K intake air to124 K before the air enters the turbo-compressor. This is a heat transfer equal to 400 mega-Watts of power and willcause the largest source of irreversibility throughout the engine cycle. Liquid helium is purposed to be usedthroughout the pre-cooler tubes to drop the in-take air temperature. Helium was seen to have a higher heat transferthan liquid nitrogen when tested by Reaction Engines LTD [10][11]. The turbo-compressor will use this dense, coldair to re-compress the helium coolant used in the pre-cooler and to also drive the liquid helium pump for the pre-cooler before entering the engines preburner. The driving of the turbo -compressor by the intake oxygen will raisethe oxygen temperature to 693 K, where it will then be combustible with the LHX fuel in the engines preburner. Inair-breathing mode the compressed air is limited to 700 K and the LHX to 876 K [10][11]. These temperature limitsare set so that the nozzle jacket of the combustion chamber can be implemented in air-breathing mode.

The intense temperature gradient across the pre-cooler must not cause a large drag penalty. Reaction Engines LTDdocumented that the purposed staggered layout of the coolant tubes caused minimum drag by creating a counterflow effect. Minimum drag is also important in combating the effects of pre-cooler frost over. The 124 K lowertemperature limit of the pre-cooler is far below the freezing point of many molecules in the atmosphere. If thesegaseous molecules freeze over with in the pre-cooler than large drag penalties will be seen from the engine. Freezeover will also cause a loss of efficiency in the engines performance, if not stalling the engine all together. Thefreezing over of the pre-cooler could even cause internal damage of the engine components and result in the

destruction of the engine.

1350 K bypass flow will be mixed with preburner exhaust gases for an assisting ramjet burner sequence. This willhelp bust the thrust power of the SABRE engine. The burning of exhaust gases in the ramjet increases the enginesefficiency, making the SABRE a more effective and environmentally conscious engine. The byproduct of the LHXand LOX burns is pure water, even in air-breathing mode. This production of pure water is a very beneficial factorand helps further the favorability of this engine system.

6.3 Brazing Techniques

Brazing is a process like welding, in which 2 metal bodies are joined by a high heat. Before brazing the pre-coolercomponents must be electroless nickel plated for improved heat resistivity. The electroless plating process was seento be more beneficial then anodic electroplating. Anodic electroplating is the plating process used by large car

manufactures. The anodic process uses a magnetic field to attract nickel, but small ridges and holes in the metalcause varying magnetic fields. The varying magnetic fields lead to a poor nickel deposit over the plating surface.

The technique for brazing the small pre-cooler tubes to the larger pre-cooler tubes required refinement by ReactionEngines LTD. The brazing process was refined on tubes with a diameter of 0.38 mm. The 0.38 mm tubes are smallerthan the purposed full size SABRE pre-cooler 0.84 mm inner diameter tubes. If the refinement process works on thesmaller 0.38 mm tubes than it will surely work on the full-sized pre-cooler tubes.

The normal brazing techniques for nickel to phosphorous call for brazing temperature of 980 degrees Celsius to1,000 degrees Celsius. This temperature allows for proper eutectic bonding of the metal tube bodies. During brazing


13/27


testing of the small 0.38 mm tubes, the 980-1,000 degree range was seen to cause excess erosion, which clogged the pre-cooler tubes. Wall Colmony perfected the brazing process and found that a proper eutectic formation could beachieved at 910 degrees Celsius with an acceptable 5 m of erosion. The formation does require longer brazing time

but resulted in such favorable effects that the time increase is seen as tr ivial. Colmonoys eutectic formation is nowknown as Nicrobraz 10 and was his doctoral dissertation [11].

7.0 Implementation Year and NASA TRLThe latest tests done by REL occurred in 2012. The test was of a 9% scale SABRE and placed the pre-cooler on aViper engine and investigated aerodynamic stability and uniformity, structural integrity, freedom of vibration acrossa wide range of flight envelopes, and preliminary cryogenic cooling [12]. The tests secured REL 60M from theBritish government to further invest in the project, improve the heat exchanger, wind tunnel test SABRE enginecomponents, and demonstration of the engine. The current timeline for a prototype SABRE is 2017, with test flightsexpected in 2020.

Given the current state of the SKYLON project, a NASA technology readiness level of 5 was approximated by thegroup. There is no actual prototype as of yet and only the pre-cooler has demonstrated to be working outside of alaboratory environment.


14/27


Appendix A:Chapter 2 Figures


15/27


Figure 2.1: HOTOL Design [6]


16/27


Appendix B:Chapter 3 Figures


17/27


Figure 3.1: SKYLON Space Plane [1]

Figure 3.2:Re-entry Heat Concentrations[7]


18/27


Figure 3.3: SABRE Engine [2]

Figure 3.4: SABRE Pre-Cooler [1]


19/27


Figure 3.5: Payload Bay [1]


20/27


Appendix C:Chapter 4 Figures


21/27


Table 4.1: Engine Component AttributesEngine Component Attribute Entropy per Unit Mass Flow (J/kgK)

Precooler T=30 257HX4 T=7 173C1 c=0.88 208C2 c=0.88 97

T1 t=0.9 63T2 t=0.9 8

Figure 4.1: SABRE cycle [5]

Figure 4.2: Pre-cooler


22/27


Figure 4.3: The JMHX [11]


23/27


Appendix CD:Chapter 5 Figures


24/27


Figure 5.1: LAPCAT A2 [2]

Figure 5.2: Scimitar Engine [2]


25/27


Figure 5.3: Orbital Base Station [13]


26/27


Figure 5.4: Modular Construction of a Vehicle Stage [13]

Figure 5.5: Troy Space Vehicle [14]


27/27


References

[1] Shubber, K. (2013, August 13). SKYLON: Alan Bond's mission to replace space rockets with spaceplanes.Retrieved from http://www.wired.co.uk/news/archive/2013-08/12/skylon-alan-bond

[2]Reaction Engines Limited. (2013). Retrieved from http://www.reactionengines.co.uk/

[3] Robertson, D. F. (1994). Commercial Space, Single Stage to Orbit SSTO. Retrieved fromhttp://www.donaldfrobertson.com/ssto.html

[4] Bond, A., &Varvill, R. (2003). A Comparison of Propulsion Concepts for SSTO Reusable Launchers Journal ofthe British Interplanetary Society (56 vols.,Pg 108-117). Abington, Oxfordshire: Culham Science Centre.

[5] Webber, H., Feast, S., & Bond, A. (2008). Heat exchanger design in combined cycle engines. International Astronautical Federation, 8 (4.5.1).http://www.reactionengines.co.uk/tech_docs/Heat%20exchanger%20design%20in%20combined%20cycle%20engines%20IAC-08-C4.5.1.pdf

[6] HOTOL. (n.d.). Retrieved from http://www.astronautix.com/lvs/hotol.htm

[7] Martin, J. (n.d.). SKYLON re-entry - SKYLON space plane gets a thumbs-up (images). Retrieved fromhttp://news.cnet.com/2300-11386_3-10007900-8.html

[8] Bond, A. (2009). SKYLON Users Manual. 0001 , 1.1st ser.

[9] Varvill, R. & Bond, A. (2004). The SKYLON Spaceplane Journal of the British Interplanetary Society (57 vols.,Pg 22-32). Abington, Oxfordshire: Culham Science Centre.

[10] Webber, H., Bond, A., & Hempsell, M. (2007). The Sensitivity of Pre-cooled Air-breathing EnginePerformance to Heat Exchanger Design Parameters Journal of the British Interplanetary Society (60 vols.,Pg 188-196). Abington, Oxfordshire: Culham Science Centre.

[11] Murray, J., Hempsell, C., & Bond, A., (2001). An Experimental Precooler for Airbreathing Rocket Engines Journal of the British Interplanetary Society (54 vols.,Pg 199-209). Abington, Oxfordshire: CulhamScience Centre.

[12] Reaction Engines Ltd. (10 July 2012). Major Advance Towards the Next Jet Engine[Press Release]http://www.reactionengines.co.uk/press_release/MAJOR%20ADVANCE%20TOWARDS%20THE%20NEXT%20JET%20ENGINE%20-%20PRESS%20RELEASE%20-%2010%20JULY%202012.pdf

[13] Feast, S., & Bond, A. (2008). A Design for an Orbital Assembly Facility for Complex Missions. International Astronautical Federation, 8 (3.3.1). http://www.reactionengines.co.uk/tech_docs/Design%20of%20an%20orbital%20base%20facility%20for%20complex%20missions%20IAC%2008%20D3.3.1.pdf

[14] Martin, T., Varvill, R., & Bond, A. (2007) Project Troy, A Strategy for a Mission to Mars.http://www.reactionengines.co.uk/tech_docs/mars_troy.pdf

[15] Bond, A., Varvill, R. (2004) Application of Carbon Fibre Truss Technology to the Fuselage Structure of theSKYLON Spaceplane . Abington, Oxfordshire: Culham Science Centre.

skylon space plane - final

Documents