Supergun Group 1

Supergun Satellite Launch System

Multidisciplinary Design Project –Group 1 Inception Report

Aidan Cookson, Jack Feakins, Kin Man Benjamin Lee, Sophie McElhill, Alex Oses,

Eleri Williams

27/10/2014

FAO: Prof. Roger Webb, Prof. Neil Downie, Dr. Chris Bridges & Dr. Ignazio Cavarretta

This inception report reviews the current progress towards a non-rocket launch of a useful

payload to orbit and proposes technical options and design considerations for an

economical, environmentally sustainable and reusable launch system. The proposed launch

system is comprised of a hydrogen light gas gun launching a projectile at hypersonic

velocities to reach Low Earth Orbit. This report discusses technical issues and risks such as

launch systems, hypersonic projectile design, high temperature and high g survivability.

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Contents

1. Introduction ........................................................................................................................... 2

1.1. Background .................................................................................................................... 2

1.2. Scope of Project .............................................................................................................. 2

1.3. Design Specification ....................................................................................................... 3

1.4. Project Market ................................................................................................................ 3

1.5. Analysis of Competitors ................................................................................................. 3

2. Key technologies and considerations .................................................................................... 5

2.1. Launch System ............................................................................................................... 5

2.1.1. Ballistics .................................................................................................................. 5

2.1.2. Booster ..................................................................................................................... 9

2.1.3. Structure & Mechanical Design ............................................................................ 10

2.2. Projectile ....................................................................................................................... 11

2.2.1. Aerodynamic Effects ............................................................................................. 11

2.2.2. Thermal Effects ..................................................................................................... 13

2.2.3. Orientation and Stability ........................................................................................ 14

2.2.4. Structural Architecture .......................................................................................... 14

2.2.5. In-Barrel ................................................................................................................ 15

2.3. Projectile On-board Electronics.................................................................................... 15

2.3.1. Projectile Electronic System Overview ................................................................. 15

2.3.2. Guidance, Navigation & Control (GNC) ............................................................... 16

2.3.3. Telecommunications System ................................................................................. 16

2.3.4. Propulsion / Reaction Control System ................................................................... 18

2.3.5. High/Low Temperature Electronics Survivability ................................................. 18

2.3.6. High g Launch Electronics Survivability .............................................................. 18

2.4. Trajectory Simulation ................................................................................................... 20

2.5. Location & Structures ................................................................................................... 21

2.5.1. Considerations ....................................................................................................... 21

2.5.2. Requirements ......................................................................................................... 22

2.5.3. Locations Considered ............................................................................................ 23

2.5.4. Recoil Absorption System ..................................................................................... 23

2.5.5. Launch Site & Ground Work Considerations ........................................................ 24

3. Project Risk Management .................................................................................................... 25

4. Project Organisation & Management .................................................................................. 27

5. Bibliography ........................................................................................................................ 29

Appendices .............................................................................................................................. 31

Appendix A ......................................................................................................................... 31

Appendix B .......................................................................................................................... 32

Appendix C .......................................................................................................................... 32

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1. Introduction

1.1. Background

Ever since the first satellite launch in 1957, rockets have been used to launch satellite payloads

into orbits around the Earth for purposes such as communications, navigation, and earth

observation including weather and research (NASA, 2007). In 1966 the High Altitude Research

Project (HARP), led by ballistics engineer Gerard Bull, used a very large naval gun to achieve

a record non-rocket launch of a 180 kg projectile at 3,600 m/s to space, reaching an altitude of

180 km (Astronautix, n.d.). The HARP project was intended for research into the ballistics of

re-entry vehicles; however it is known that Gerard Bull had envisioned the use of the system to

launch satellites into space at a lower cost than conventional rocket launches (Astronautix, n.d.).

Technology has advanced significantly since 1966, leading to satellites that are smaller and

lighter, and navigation systems for launch projectiles that can be made using solid state

electronics capable of withstanding high acceleration forces. It is thought that the ‘supergun’

non-rocket launch method can be used in combination with an orbital insertion rocket booster

and modern technologies to provide a lower cost and more efficient small satellite launch

method compared to conventional rocket launches.

1.2. Scope of Project

The aim of this project is to take a fresh look at the non-rocket launch method pioneered by

Gerard Bull using a ‘supergun’, investigating key issues such as costs, feasibility and hazard

analysis. The fundamental requirements are outlined below;

To design a launch system for the insertion of useful payloads into Low Earth Orbit,

using a combination of ballistics and rocket propulsion.

To design a more economical and environmentally sustainable launch system than

conventional rocket systems.

The project must consider the construction, regular use, decommission and mission

abort scenario of the launch system and payload to prevent any avoidable damage to

the environment or living creatures as a result of the project.

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1.3. Design Specification

The specification for the project is outlined as follows;

The mass of the useful payload must be greater than 1 kg.

The projectile must be capable of communicating with ground control.

The payload must be capable of achieving stable attitude in orbit.

The projectile should ensure the safe transit of the useful payload to the desired orbit

and have safety features in place for mission abort, should the projectile fail in transit.

The launch system must be capable of launching repeatedly to ensure economic

viability.

The launch system, projectile and all associated systems must comply with all

relevant safety and environmental regulations including: UK Outer Space Act 1986,

Explosives Handling & Storage Regulations, and Launch Regulations.

1.4. Project Market

The current satellite market is estimated at approximately$190 billion USD (Emerick, 2014),

(The Tauri Group, 2013), (Northern Sky Research, 2012). It is estimated that 34-45% of the

market is held by Low-Earth Orbit satellites (The Tauri Group, 2013), (Emerick, 2014). At the

34% estimate this equates to a market share of approximately $65 billion USD. Last year

satellite launches with payloads ranging from 1kg to 50kg, was just below 100 satellites

(SpaceWorks, 2014), this figure is 45% of the total satellites launched in 2013 (SpaceWorks,

2013). The number of 1kg to 50kg satellites launched in the future is predicted to increase

steadily each year, totalling 2000-2750 between 2014-2020 (SpaceWorks, 2014). Evidence

suggests that the nanosatellite (1kg to 10kg range) will be a key shareholder in this predicted

number with the increase of nanosatellite launch attempts between 2012-2013 being 330%,

compared to a slight decrease in the number of microsatellite launch attempts (SpaceWorks,

2014).

1.5. Analysis of Competitors

Currently the most common method for small satellite launch is as a piggyback payload. This

is where a multi-stage rocket is used to launch a main payload, and any extra capacity on-board

the rocket projectile is utilised by small satellites. The launch costs associated with this method

are estimated at $22,000 to $29,000 USD per kg (Ley, 2009).

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Another launch method under development for delivering small satellites to Low-Earth Orbit

or Sun Synchronous Orbit is Super Strypi. Super Strypi is a three stage rocket which utilises

legacy rocket designs and has an expected payload capability of 250kg to Sun Synchronous

Orbit at 400km (Gunter, 2014). The expected (recurring) cost per mission is $5 million USD,

giving an approximate launch cost of $20,000 USD per kg (Schindwolf, 1998).

The largest source of expenditure in multistage rocket launches is in the first stage rocket which

is typically destroyed by aerodynamic and impact forces after use. There are several programs

which look to alleviate these costs by introducing re-usable first stage rockets or combined

ducted rockets with a second stage chemical rocket. The SpaceX program is an example of re-

usable first stage rockets using controlled re-entry (SpaceX, 2014). The Skylon project is

another proposal for a re-usable design, incorporating air-breathing rocket engines into a space

plane, which is intended to reduce costs to approximately $1,100 USD per kg (Reactionengines,

2012).

The development of gun-based launchers is another option for reducing expenditure by

replacing the first stage rocket. Quicklaunch Inc., founded in 2010 by John Hunter, is the most

significant and feasible contribution towards the idea of a gun based launcher of satellites,

leveraging experience from the Super High Altitude Research Program (SHARP). The project

proposes a water based hydrogen gun of 1.1 km length, designed to produce 6 km/s initial

projectile velocity with a second stage rocket booster to reach orbital velocity. The predicted

launch cost for this project is $1,100 USD per kg (Quicklaunch, 2012).

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2. Key technologies and considerations

2.1. Launch System

The launch system must fire the projectile to Low Earth Orbit and potentially Sun Synchronous

Orbit altitudes using methods which are more economically viable than conventional staged

rocket propulsion. The system is to consist of two main launch techniques; a ballistic element

and a single stage rocket element.

2.1.1. Ballistics

There are several different types of high power gun technologies that are readily available

including: electric launch, light gas guns, powder guns and RAM accelerators. For this project

powder guns and RAM accelerators have been deemed insufficient.

2.1.1.1. Electromagnetic Launchers

Electromagnetic launchers use electromagnetic induction to accelerate the projectile to the

desired muzzle velocity using the principles of Faraday’s law and Lorentz force. There are three

main configurations for an electromagnetic launcher: linear motors/mass drivers, railguns and

coilguns. Linear motors are limited to low velocities, with the fastest applications reaching just

over 100m/s (McNab, Electromagnetic Augmentation Can Reduce Space Launch Costs, 2013).

Coilguns are also limited to very low delta-v with the highest recorded coilgun launch over two

decades ago reaching a delta-v of approximately 1km/s for a small projectile (Kaye, 1995).

Very little research into coilguns has been conducted in recent times.

The main issue facing railgun launch systems is energy storage. These launch systems require

very high power levels to achieve the required delta-v to reach orbital altitude. Table 1 shows

a study of a number of electromagnetic launchers, scaled by diameter of the bore db. The table

gives values for diameter of the bore db, gun barrel length s, launch mass mt, muzzle kinetic

energy Emuz, current I, the gun rails back EMF Vb, the linear current density in the rails I*, and

the estimated instantaneous power requirement for the gun (McNab, Pulsed power options for

large EM launchers, 2014). The study uses the assumptions that the gun has a round bore and

is fired vertically. For scaling purposes there are a number of set variables which may be found

in the original text along with the mathematical relationships between these parameters. For a

projectile around the chosen useful payload reaching a muzzle velocity of approximately

1900m/s, using the simple relationship [1] a power requirement of 11.7GW is required. Where

v is the muzzle velocity, a is the average acceleration, given as 250km/s2 in the original text,

and s is the gun barrel length:

𝑣2 = 2𝑎𝑠 [1]

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The calculation of 11.7GW ignores the additional dry mass of the projectile and the mass of

the rocket booster required for orbital injection. With inclusion of this additional mass we can

see that the power requirements for an electromagnetic launch system will be in excess of

70GW. The energy system would also need to supply currents of several mega-amperes

which current technology is not capable of producing on the scale of this application (McNab,

Pulsed power options for large EM launchers, 2014).

Table 1: Parameter scaling with gun bore diameter (McNab, Pulsed power options for large EM launchers,

2014).

db

(mm) s (m) mt (kg)

Emuz

(MJ) I (MA) Vb (kV)

I* (kA/m)

IVb

(GW)

10 0.6 0.01 0.001 0.09 0.02 10.8 0.002

20 1.2 0.06 0.02 0.24 0.09 15.2 0.02

40 2.4 0.46 0.27 0.68 0.37 21.5 0.25

60 3.6 1.54 1.39 1.24 0.83 26.4 1

80 4.8 3.66 4.4 1.91 1.48 30.5 2.8

100 6 7.14 10.7 2.67 2.31 34 6.2

120 7.2 12.3 22.2 3.51 3.33 37.3 11.7

140 8.4 19.6 41.2 4.43 4.54 40.3 20.1

160 9.6 29.3 70.2 5.41 5.93 43.1 32.1

200 12 57.2 171 7.56 9.26 48.2 70

2.1.1.2. Light Gas Guns

A light gas gun uses a compressed gas with low molecular weight, such as Hydrogen or Helium,

to propel a projectile. A two-stage light gas uses conventional gun powder to cause rapid

expansion in the combustion chamber which pushes a piston, which then compresses the gas.

When the pressure in the gas chamber reaches a prescribed value, a diaphragm separating the

gas chamber and the gun barrel ruptures. This allows the gas to expand, propelling the projectile

down the barrel. Figure 1 shows a schematic diagram of this process. This type of gun can reach

muzzle velocities in excess of 7.5km/s (NASA, n.d.) and has been proposed as a suitable gun

type for space launches by projects such as the Super High Altitude Rearch Project (SHARP)

and the Quicklaunch project. Figure 1 shows a schematic of a light gas gun at each stage.

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Figure 1: Schematic of light gas gun (Bernier, 2005)

In any type of gun the velocity of the projectile cannot exceed that of the propellant gases. Light

gas guns are able to reach higher velocities than traditional gas guns as the lower molecular

density gives the gas a higher velocity at a given temperature, enabling the projectile to be

propelled faster.

The limiting factors in the velocity of the projectile are the speed of sound of the gas and the

specific heat ratio of the gas. Assuming a single-stage light gas gun, the maximum velocity of

the projectile is given by [2], where 𝑣𝑙𝑖𝑚 is the limiting muzzle velocity, 𝑎0 is the speed of

sound of the gas and 𝛾 is the specific heat ratio of the gas (Bernier, 2005).

𝑣𝑙𝑖𝑚 =2𝑎0

𝛾−1 [2]

The maximum velocities for hydrogen and helium at a temperature of 300K are given in Table

2. It can be seen that hydrogen can give sufficient muzzle velocities for the required application.

Table 2: Maximum projectile velocity

Gas γ a0 (m/s) vmax (m/s)

Hydrogen 1.41 1260 6200

Helium 1.66 970 2930

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To estimate performance and requirements of the gas gun, [2] has been used to calculate the

required initial pressure of the hydrogen. This equation has been developed for a single stage

light gas gun, therefore the effect of the piston and the gunpowder has not been considered and

the gas is assumed to be a perfect gas with no losses due to friction or heat transfer (Bernier,

2005).

[3]

Where p0 is the initial pressure, a0 is the speed of sound in the gas, v is the muzzle velocity, γ is

the ratio of the specific heats, M is the mass of the projectile, L is the barrel length and S is the

barrel cross sectional area. The mass of the projectile has been assumed to be 150kg and a bore

of the barrel has been assumed to be 0.8m. The required velocity has been assumed to be 6km/s.

The pressure required to reach this velocity has been calculated as a function of barrel length

and gas temperature; the results are plotted in Figure 2. From this data, at a temperature of 900K

and a barrel length of 700m the pressure required is 367MPa. This is an achievable pressure

and is similar to that used by the SHARP gun, which had a hydrogen pressure of 405MPa

(Astronautix, n.d.).

Figure 2: Plot of barrel length against required pressure for a range of temperatures

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2.1.2. Booster

A booster rocket will be incorporated into the projectile launch system to achieve orbital

velocity at the required altitude, if necessary there will be an additional booster stage to assist

reaching the required altitude. There are several classifications of rockets including ducted,

chemical, nuclear rockets and electrical rockets. For the purpose of this project ducted rockets

have been deemed unfit as their effective altitude limit is approximately 14-20km. Alternative

electric propulsion systems characteristically have low thrust and therefore low acceleration,

which makes them unsuitable for this application. Nuclear rockets are also unfeasible due to

the complications associated with them such as power control, radiation and intense heat.

Technology readiness is also a major factor in the decision hence chemical rockets have been

selected for this project.

2.1.2.1. Chemical Rockets

Chemical propulsion rockets create propulsion from hot gas expansion created by the

combustion of chemical products. There are several types of chemical rocket which may be

considered for this application, these are;

Liquid bipropellant – This propulsion system consists of a liquid fuel and a liquid oxidiser

which are mixed in a combustion chamber, ignited and exhausted through a nozzle to create a

hot gas jet thrust. These can be controlled by using a pressurised gas feed or a pump feed system

depending on the fuel burn rate required.

Liquid monopropellant – A liquid monopropellant system uses a liquid fuel which also

contains an oxidising species. The liquid produces a hot gas when properly catalysed which

produces thrust through a nozzle.

Solid propellant – Solid fuel rockets give a continuous thrust once ignited until the fuel is fully

combusted. These rockets may not be adjusted or re-ignited once extinguished. The combustion

takes place in the same chamber as the fuel storage and therefore the system is very simple and

does not require additional pressurised chambers or valves.

Hybrid – hybrid propellant rockets have a combination of both a liquid and solid propellant.

Typically they consist of a solid fuel grain with a liquid injected oxidiser. Similarly to a solid

propellant rocket the solid fuel combusts in the same chamber in which it is stored. This type

of rocket may be controlled to give variable thrust and also exhibit stop and start capability.

The system requires only one valve to control the injection of liquid oxidiser and two storage

chambers, making it a relatively simple design. Some disadvantages of hybrid rocket propulsion

are variations in specific impulse known as engine chugging (Biblarz, 2010).

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Table 3 shows a comparison of the aforementioned rocket systems. It is clear from this table

that a hybrid propellant rocket will be the most suitable for this mission, and as of such shall be

adopted in continuation of the project.

Table 3: Comparison of chemical rocket propellant arrangements (Boiron & Cantwell, 2013) (Jubb, 2011)

(Biblarz, 2010).

Type of Rocket Specific

Impulse (sec)

Thrust to Weight

Ratio kN/kg

Technology

Readiness

Chemical – Solid or Liquid

Bipropellants 200-468 10-2 - 100 Flight Proven

Liquid – Monopropellant 194-223 10-1-10-2 Flight Proven

Hybrid Propellant 300-400 10-1 Flight Proven

2.1.3. Structure & Mechanical Design

Based on the analysis of alternatives in Section 2.1 the proposed launch system will use light

gas gun technology with hydrogen as the propellant gas. The following parameters will be

considered during optimisation of the gun design;

Initial temperature, pressure and volume of the light gas.

Mass, diameter and distance travelled of the piston.

Type of gun powder used, its mass and the combustion chamber volume.

Pressure at which the diaphragm ruptures and the projectile is realeased.

Barrel vaccuum system.

Design of the projectile.

Barrel thickness.

2.1.3.1. Gun Support Structure

One option for the support of the gun barrel is to install it within the ground at the launch

location. The main advantage of this is it provides uniform support along the length of the gun

barrel in addition to thermal shielding; this limits the thermal gradient of the structure. There

are associated disadvantages to this option such as construction challenges and inability to

adjust the inclination and orientation of the barrel; this may limit launches to a small range of

orbital altitudes.

Another option for barrel support is the use of mechanical structures such as trunnions, truss

supports or hydraulic cylinders. These options facilitate rotation and launch attitude adjustment

however considerations will be necessary for thermal shielding of the barrel. The barrel can be

made from a light material or made thinner incrementally along its length.

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To mitigate strains from launches an effective recoil system, a sufficiently strong structure, and

an incredibly straight projectile acceleration path are required.

2.1.3.2. Barrel Thermal Shielding

To reduce thermal strain on the barrel a cooling system will be implemented using thermal

shielding. This can be achieved by painting the barrel with silicate white paint or covering the

barrel with Aluminised or Silver coated FEP mitigating the power usage associated with a heat

exchange system (Savage, 2011).

2.1.3.3. Attitude Control

Optimal launch attitude has been calculated at 20-30° from the normal to the Earth’s surface

(vertical). For a varying attitude gun the support system will include attitude control, a sketch

of this proposed design may be found in Appendix B.

2.2. Projectile

The objective of the projectile is to safely deliver the useful payload to orbit. The constraints

within which the useful payload will conform are: it must have a mass less than 20kg, and fit

within a 30cm in diameter cylinder of length 30cm. It must also be able to sustain acceleration

forces of 1000g (P.Cox, 2014).

2.2.1. Aerodynamic Effects

Upon leaving the gun barrel, the projectile is expected to be travelling at hypersonic velocities

and therefore key aerodynamic characteristics must be considered. The projectile will create

shock waves on or before the leading edge and thick, viscous, turbulent boundary shock layers.

Furthermore, temperature rises significantly across the shock wave. This makes projectile

design key to maintaining aerodynamic efficiency and ensuring projectile survivability.

For a sharp body the shockwaves formed will be attached oblique shock waves, either

compressing or expanding the flow depending on the direction of the angle the hypersonic flow

is required to turn through. The flow behind an attached oblique shock wave typically remains

supersonic. When the deflection angle is greater than that of the maximum achievable by the

hypersonic flow, a detached shock wave will form in front of the body, which is known as a

bow shock, see Figure 3. Behind a detached oblique shock wave there is a region of subsonic

flow, before a transition returning to supersonic flow.

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Figure 3: Left, oblique shock wave attached to the nose of a sharp body. Right, bow shock wave detached

from the surface of a bluff body (Hartman, 1968).

The contribution to drag due to a shock wave is called wave drag, this can be calculated from

the deflection angle of the flow and the pressure drop across the flow (Grabow, 1965).

In addition to this there will be another component of drag called the parasitic drag. This is

comprised of skin friction and pressure drag. The skin friction is related to Reynolds number in

turbulent flows by the 1/7th power law. This component of drag will therefore scale with the

characteristic length of the projectile. Pressure drag is related by the pressure coefficient of the

projectile, which is determined by boundary layer reattachment due to projectile shape.

At sufficiently high Mach number the lift, pressure and drag coefficients of an object behind a

shock wave become Mach number independent (Anderson J. D., 2001). An estimated drag

coefficient for the projectile has been chosen as 0.6 based on a cone with wedge angle of 60°

(Anderson J. D., 2001).

The design of the projectile will be a trade-off between the optimal aerodynamic profile and a

geometry which limits the thermal energy absorbed by the projectile, to ensure payload

survivability.

Figure 4: Conceptual CAD design of the supergun projectile

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Figure 4 is an initial design for the projectile. It depicts a pointed nose with a slender body and

a rear ‘boat tail’ design, with fins for stability. Shown in white is a component for thermal

shielding, in green are the electronics, control systems and avionics, in red is the useful payload

bay, in bronze is the liquid oxidiser pressure vessel of the hybrid rocket, in blue is the solid fuel

and combustion chamber vessel, and in black is the rocket nozzle.

Once an initial design has been created it is possible to use computational fluid dynamics to

compare analytical solutions for the coefficient of drag. Computational results may also be used

to predict the temperature on the surface of the projectile.

2.2.2. Thermal Effects

Across an oblique shockwave, there is an almost instantaneous rise in temperature, resulting in

the conduction of extreme heat to the skin of the projectile. From isentropic flow relationships,

it can be shown that the temperature downstream of a shock wave at Mach 10 is more than 10

times greater than the upstream temperature. Another aspect to consider is the temperature

caused by the stagnation point on the nose of the projectile. The process of losing the air flow’s

kinetic energy results in calculable localised high temperatures. Furthermore, thermal

conduction on the skin of the projectile can be significant (Allen, 1958).

When in orbit, one side of the projectile will be subject to solar radiation this will create large

temperature gradients which must also be taken into account.

A thermal shield on the nose of the projectile will be required to reflect and dissipate heat away

from the internals of the projectile. There are a range of methods for thermal shielding, three of

these methods are;

Thermal barrier coatings such as yttria-stabilized zirconia.

Thermal radiation reflection using elements such as gold and silver.

Ablative thermal shields such as AVCOAT.

Insulation will also be applied to any exposed structural architecture to ensure no significant

thermal gradients will affect the integrity of the core structure. The selection and fixing of such

insulation will be driven by the operational performance of each option and the availability and

cost.

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2.2.3. Orientation and Stability

The orientation and spin of the projectile must be stable after launch to enable a predictable

altitude and trajectory. Once the projectile is launched, stability and orientation will be obtained

passively through the aerodynamic design and stability mechanisms. When in orbit, active

orientation control systems will be required to make translational as well as rotational

adjustments.

The two main concepts for ensuring this while maintaining low drag are;

Gyroscopic stability obtained through rotating the projectile about its primary axis.

Centre of pressure position.

Gyroscopic stability is common place in the design of artillery shells. Most cannons will have

a rifled barrel to impart a spin on the projectile. Once spinning the angular momentum of the

projectile means that any perturbations will create gyroscopic procession. This phenomenon is

typically not an issue for artillery shells due to their limited time in the air before hitting a target.

This act of procession will increase the drag acting on the projectile.

Moving the centre of aerodynamic pressure behind the centre of mass will increase

aerodynamic stability. To move the centre of aerodynamic pressure, drag must be increased

towards the rear of the projectile. This can be achieved through the application of fins to the

rear; the overall coefficient of drag will also increase. The in-flight integrity of the fins will be

critical to the survivability of the entire projectile, thus must be designed to withstand the

aerodynamic pressures associated with hypersonic flight.

2.2.4. Structural Architecture

Similar to conventional launch vehicles, the projectile will be required to withstand large

vertical loads due to the extreme acceleration; however these will more likely be in order of

1000g compared to a conventional rocket launch peak acceleration of 24g (P.Cox, 2014). Along

with the high acceleration, the resulting extreme aerodynamic pressure acting on the projectile

will require the structure to withstand high hoop stresses. Therefore the projectile’s core

structure is required to be strong as well as light, to facilitate a survivable and efficient launch.

Furthermore, at launch, the rear face of the projectile will be required to absorb the shockwave

and pressure created by the expanding gas behind it, allowing the projectile to reach hypersonic

velocities. External factors such as untracked micrometeorites will require the structure to resist

penetration from high velocity, low mass projectiles.

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Along with securely containing all components of the projectile, the structure will also have to

break apart and separate from the useful payload to release the useful payload into orbit. After

the release of the useful payload, the launch projectile must complete a controlled de-orbit

process through the atmosphere to ensure that no unnecessary debris is left in orbit after the

mission.

Typical rocket launch vehicles use an aluminium lithium alloy such as AL-Li 2195 which is a

high strength light alloy. Structural strength can be achieved using a truss architecture,

minimising the material and therefore mass required. A range of materials will be suitable for

designing the structure of the projectile including Al 6062, Ti 6-4 or composite materials such

as carbon or Kevlar fibre reinforced metals or graphene reinforced carbon.

2.2.5. In-Barrel

To ensure the maximum achievable efficiency at launch using a first stage supergun, the

projectile fired must be machined to a close tolerance of the bore of the barrel. To enable the

projectile to also be of a geometry optimised for aerodynamic effects, a sabot is used to ensure

close tolerances between the barrel and the projectile. Additionally the sabot acts as a barrier

between the gun barrel and the surface of the projectile to ensure the integrity of the projectile

upon leaving the barrel. Through the use of a weak, sacrificial surface it can also limit the wear

experienced by the bore of the barrel. Design of the sabot will be closely linked to the geometry

of the projectile to ensure an even load distribution across the projectile. The sabot will also be

designed to be aerodynamically unstable to ensure that it will quickly tumble away from the

projectile after launch.

2.3. Projectile On-board Electronics

2.3.1. Projectile Electronic System Overview

Figure 5 shows a sub-system overview of the projectile on-board electronics. Key aspects of

the electronic system are detailed in the following sections.

Figure 5: Electronic system overview for projectile

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2.3.2. Guidance, Navigation & Control (GNC)

The information from sensors in the GNC, supply the control systems with feedback

measurements. Additionally, the data can support the ground control station operation to

monitor the progress of the projectile following launch. This section describes the common

sensors used on spacecraft.

2.3.2.1. Attitude sensors

Attitude sensors are used to generate reliable data to describe the physical orientation and

position of the projectile. It is very common for the projectile to have a desired path to follow

designated prior to launch. The sensor data is used in an internal feedback control system to

maintain this path as well as being transmitted to the earth station.

There are two types of attitude sensors typically used for launch projectiles: relative attitude

sensors and absolute attitude sensors. These sensors vary in size, weight and sensing

mechanism.

A gyroscope is one of the most common relative sensors, which senses rotation in three

dimensions (roll, pitch and yaw). Traditional gyroscopes rely on a spinning moment of inertia,

which does not change orientation with the surrounded force acting on it. This is then referenced

to detect the orientation of the projectile.

2.3.2.2. Pressure sensors

Pressure sensors are devices that detect the pressure of gases or liquids. They can be used to

measure the fuel tank pressure and also the internal pressure of the projectile. Pressure sensors

can also support the control system to optimise the combustion of hybrid fuel.

2.3.2.3. Servo type accelerometers

Servo type accelerometers detect the acceleration of an object in three dimensions of

translational movement. Traditional accelerometers used piezoelectric or capacitance

technology to measure acceleration. These traditional mechanisms lead to a small mass

displacement that reduces the accuracy of the accelerometer. In comparison, the servo type

accelerometer is a closed loop device, which keeps internal deflection of the proof mass to a

minimum, hence providing a better accuracy for measurement of acceleration (Wilson, 2004).

2.3.3. Telecommunications System

A communication system is needed for the projectile in order to communicate with ground

control. The system is expected to receive instructions from the earth station and also send

essential information such as physical condition and orientation of the projectile to the ground.

This section describes the communication subsystem used and also the antenna of the projectile.

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2.3.3.1. Telemetry, tracking, and command subsystem (TT&C)

The Telemetry, Tracking, and Command subsystem (TT&C) is responsible for data

communications enabling the monitoring and controlling of the projectile from take-off to an

orbit location (Pattan, 1993). Early satellites used very high frequency (VHF) and Super high

frequency (SHF) for TT&C data transmission due to lower power consumption (Pattan,

1993). Figure 6 shows a schematic of the proposed telemetry, tracking & command system.

Figure 6: Schematic of telemetry, tracking, and command (TT&C) system.

The telemetry system collects all the data from the sensors of the projectile including

temperature, pressure in the fuel tank and attitude. The data is usually digitalized and coded

with Frequency Shift Keying or Phase Shift Keying before transmission via an omnidirectional

antenna (Mitra, 2005). Uplink and downlink frequency are usually different to prevent

unwanted interference.

The tracking system is used to determine the orbit of the projectile by calculation from

previously detected positions, using accelerometer and velocity sensor data. This data provides

an estimated range of movement for the projectile. The location of the projectile can then be

cross-checked and corrected by triangulation of multiple earth stations with time delay and earth

elevation angle, similar to a GPS system (Mitra, 2005).

The command system is a system for controlling the projectile to perform specific actions after

processing the received data. For example the system can be used to manually adjust attitude

to protect the projectile from entering other satellite orbits due to unexpected disturbances. To

prevent unauthorized TT&C command, the data transmitted is normally encrypted.

2.3.3.2. Antennas

An antenna is an electrical device that is used as a transmission medium for radio waves. In

satellite communication, they are mainly classified into two types of antennas, earth station

antennas and satellite antennas (Roddy, 1989).

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In designing the antenna it is important that the gain should be balanced with power

consumption to produce a sensible Equivalent Isotropic Radiated Power (EIRP).

2.3.4. Propulsion / Reaction Control System

A control system is to be devised which incorporates the activation of the secondary booster

stage to achieve altitude and velocity requirements for orbital insertion combined with

peripheral thrusters for the purpose of attitude control. It is known that by taking measurements

from the sensors used for guidance and navigation, an algorithm performing Proportional

Integral Differential (PID) control can be used to ensure that the correct trajectory is followed.

2.3.5. High/Low Temperature Electronics Survivability

It is understood that the supergun launch strategy would lead to considerable temperature

extremities due to (P.Cox, 2014), (NASA, NASA, 2014):

Compression of Hydrogen.

Barrel friction.

Shock waves.

Skin friction.

Solar Radiation.

These conditions are problematic due to the operating temperatures of electronics being limited,

typically -55°C to 125°C for military grade electronics integrated circuits and components, and

approximately -30°C to 60°C for batteries (Xilinx, 2014), (Administration, 2014).

To protect the electronics from malfunction and damage, it is necessary to implement a Thermal

Control System (TCS) which can regulate the electronics systems through passive and active

means such as Multi Layer Insulation blankets and radiators, based on temperature sensor

information. The design of thermal control systems such as would be required for this project

is well established, due to developments in electronics for atmospheric entry whereby

temperatures of 1650°C or greater can be observed (Aerospace, 2014).

2.3.6. High g Launch Electronics Survivability

One key difference between a rocket launch and gun based launch is that whereas a rocket

launch typically operates accelerations of around 4g, a gun launch can be in the order of

thousands of g (Page, 1988) (P.Cox, 2014). This is due to the rapid acceleration in the gun

launch method used to achieve the high velocity required to attempt orbital insertion.

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A key design choice in regards to the gun design to reduce the peak g force loading is using a

gun with large calibre and using a discarding sabot. Internally, within the projectile, despite g

force loading being reduced by gun design choices, there will still be a very large g force

loading.

Loading that may impact on internal circuitry can be from a number of sources, including (L3

Communications Interstate Electronics, 1999):

Setback shock from axial acceleration.

Setforward shock at gun barrel exit.

In barrel lateral shock due to imperfections in the barrel.

Centrifugal force from rifling or spinning of the projectile.

Figure 7: Gun shock acceleration vs. time profile a) theoretical b) real (Colibrys, 2012)

To accurately characterise the potentially damaging mechanical properties of the launch

method, it is desirable to analyse the short duration shock pulse in time and frequency as shown

in Figure 7. It can be seen that the real acceleration characteristic has two distinct regimes: long

high magnitude shock as in the ideal case and high frequency shocks due to imperfections in

the gun design (Colibrys, 2012). Such test curves are useful as they allow the projectile designer

to determine potential mechanical resonance damage, and to design for the high axial

acceleration loading.

Failure modes of electronic circuitry can be characterised in the following categories (L3

Communications Interstate Electronics, 1999):

Short and open circuit tracks.

Detached of parts.

Cracks in parts or epoxy.

Change in value for resistors, capacitors, inductors, crystal oscillators.

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The electronic systems of the projectile can be ruggedised in numerous ways including (L3

Communications Interstate Electronics, 1999):

Use of low height surface mount resistors, capacitors, inductors.

Use of Stress-Compensated (SC) crystals.

Use of Ball Grid Array (BGA) integrated circuit packages.

Smaller, thicker and well supported PCBs with metal stiffener supports.

Adhesive and epoxy encapsulation of components.

2.4. Trajectory Simulation

Trajectory optimisation is necessary to ensure economical and environmentally sustainable

launch. The initial estimates for horizontal and vertical delta v were determined for a fully

vertical gun. Table 4 shows the results for a variety of altitudes.

Table 4: Velocity required to achieve a range of orbits

Orbit Type Altitude (km) Theoretical velocity

required (m/s)

Theoretical orbital

velocity (m/s)

Escape Infinity 11200 N/A

LEO 160 1750 7810

LEO 1000 4120 7350

LEO 2000 5460 6900

MEO 10000 8740 4930

GEO 35800 10300 3070

The next stage was to include a drag model; this was done using Newtonian theory with a

density/altitude gradient according to the MSIS-E-90 model and a gradient for the universal

gravitational constant to give the change in the projectile’s acceleration (NASA, n.d.). This was

then used to estimate the trajectory of the projectile’s flight as a three dimensional model

accounting for the Earth’s rotation.

A key method to improve the efficiency of launch is to minimise the drag of the projectile. For

this reason the launch location will be at an altitude greater than 2km. Figure 8 shows the results

given from MATLAB. A list of the estimated parameters may be found in Appendix C. A code

for optimisation of these initial parameters is currently being developed. A pseudo code for this

can be seen in Appendix A.

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Figure 8: Altitude versus time plots comparing an initial altitude of 0 km (left) with one of 2000 km (right).

2.5. Location & Structures

2.5.1. Considerations

In order to determine a suitable launch site for this project it is necessary to consider a number

of variables, such as polar or equatorial launch site, direction of firing, elevation and

infrastructure. It was also necessary to consider what type of payload would be launched and

the type of orbit required.

An easterly launch direction has been chosen in order to take advantage of the earth’s West to

East spin to provide greater angular speed to the projectile as it is launched. In order to take

greatest advantage of this affect, a location close to the equator would be considered favourable

as the surface of the earth is spinning faster at the equator than at the poles. An equatorial launch

has been chosen as opposed to a polar launch. Due to the risk of fallout during launch and the

risk of mission failure leading to the projectile returning to earth, the site should be located on

an eastern coast with a large body of water to the east, in order to allow for easterly firing

without the risk of injury or damage to property.

Locating the launch site above sea level would be beneficial as air density decreases with

altitude, as shown in Table 5, and therefore the drag experienced by the projectile would be

reduced. Locating the launch site at altitude will present many challenges including transporting

building materials, construction and possible adverse weather conditions. A minimum altitude

of 2000m has been chosen to be a suitable compromise.

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Table 5: International standard atmosphere data (Mayhew, 1994).

Altitude (m) ρ/ρ0 ρ

0 1 1.23

500 0.953 1.17

1000 0.908 1.11

1500 0.864 1.06

2000 0.822 1.01

2500 0.781 0.957

3000 0.742 0.909

3500 0.705 0.863

4000 0.669 0.819

4500 0.634 0.777

5000 0.601 0.737

In order to minimise the risk to human life in the event of mission fall out the launch site must

be located a safe distance from populated areas.

The launch site will need to have a sufficient area to house the relevant ground works including

the firing mechanism, recoil absorption system, control centre, fuel storage and staff facilities.

It must also be suitably accessible to support all necessary groundwork, including the

construction of roads and tunnels.

The firing of the gun will cause a significant amount of recoil and so the launch site geology

must be suitable to support a recoil system without causing significant damage to the

surrounding area.

The surface geology at the launch site must be capable of supporting major building work and

be able to withstand the forces associated with recoil from the gun.

The location must not be subject to weather conditions severe enough to greatly limit the

frequency of firing.

2.5.2. Requirements

Coastal site with a large body of water located to the east.

Within 20° latitude from the equator.

Minimum altitude of 2000m above sea level.

Minimum of 10 miles from highly populated areas, with no highly populated areas

located to the east.

Suitable geology and access to support building work.

Must not be subject to severe weather conditions which could severely limit the

frequency of firing.

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2.5.3. Locations Considered

In order to find suitable launch sites the following process was used to assess the suitability of

a number of options:

List countries with an eastern cost to a large body of water that are within 20° latitude

of the equator.

Find maximum altitude of each country to filter those without high points above

2000m.

Check if there are high points near the coast.

Check the locations proximity to populated areas.

Check location geography and geology.

Table 6: Considered launch locations.

Location Maximum Altitude Latitude Suitable?

Hawaii 4,200m – Mauna Kea 19° North Yes

Tanzania 5,900m – Mount Kilimanjaro 3° South No

Kenya 5,200m – Mount Kenya 1° South No

Fiji 1,320m – Mount Tomanivi 17° South No

Philippines 2,950m – Mount Apo 7° North Yes

Papua New Guinea 4,510m – Mount Wilhelm 5° South Yes

Seychelles 905m – Morne Seychellois 4° South No

Somalia 2,460m – Mount Shimbiris 10° North No

Madagascar 2,880m – Maromokotro 14° South No

New Caledonia 1,630m – Mount Panié 20° South No

Tahiti 2,240m – Mount Orohena 17° South No

Marquises Island 1,230m – Mount Oave 9° South No

Table 6 shows the considered launch locations and their latitude in relation to the equator. The

top four countries for the launch location are Hawaii, Tanzania, The Philippines and Papua New

Guinea.

2.5.4. Recoil Absorption System

Once the gun is fired, the momentum carried by the projectile is equal and opposite to the

momentum imparted on the base of the gun; the recoil. The purpose of the recoil mechanism is

to absorb the recoil energy from the launch and dissipate it in a controlled manner. Similar

recoil absorbing mechanisms are used in modern artillery guns in order to facilitate quick firing,

essentially a large spring-damper system. These systems are typically hydraulic or hydro-

pneumatic however they also come in the form of stiff springs as well as muzzle breaks (Singh,

2014). Most recoil absorbing systems work by absorbing the kinetic energy of the barrel by

driving a piston into a chamber filled with a fluid (gas or liquid).

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The compressed fluid is then stored within a pressure vessel and then the energy is released in

a controlled manner such as through a valve to drive a piston or through a turbine to power a

drive shaft.

A recoil absorbing mechanism on a supergun is necessary to limit the effect a launch will have

on the mountings and foundations of the supergun and therefore also limit the effect on the

surrounding environment. To enable a fluid to absorb the energy from launch, the base of the

gun behind the projectile must be able to move, with respect to the piston, which will require

careful design considerations to successfully integrate it with the gun barrel and combustion

chamber. Therefore the recoil absorbing mechanism is assumed to be a structurally significant

part of the gun design.

2.5.5. Launch Site & Ground Work Considerations

Since a gun of this size has not yet been used for satellite launch applications it is not known

what effect firing will have on the ground beneath the launch site and whether or not this could

cause earthquakes or other seismic activity. Also, it would be difficult to predict the landing

site of debris in the event of a mission failure or catastrophic gun failure without the use of

sophisticated computer modelling.

Furthermore, the immediate and long term impact on the environment is not known and it is

not known if planning permission will be granted for the chosen launch site.

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3. Project Risk Management

Risk Risk Description Severity Probability

Risk

Severity

Index

Risk Mitigation

Mission Risks – Terrestrial

Projectile, booster

and/or sabot

fallout

Risk of all or parts of the

projectile, booster or sabot

landing in populated area

5 2 10

Launch direction will be due east. Launch site is on an east

coast with no populated areas within 600 km due east. There

are no densely populated areas within a 50 km radius.

Projectile collision

with air traffic

Risk of collision between the

projectile, sabot or fallout, and

private, commercial or military

air traffic.

5 1 5

The local aviation authority would have to grant permission

before launch and air traffic control would be notified. A

NOTAM (notice to airmen) would be published prior to

launch.

Uncontrolled

explosion at

launch

Risk of explosion of propulsion

products on, before or after

launch in an uncontrolled

manner.

4 2 8

Flammable material must be stored the recommended safe

manor. During launch no one will be located within 5 km of

the launch site

Uncontrolled

destruction of the

projectile after

launch

Risk of inflight destruction of the

projectile and the useful payload

at launch.

5 2 10

The projectile will be designed with a suitable safety factor to

compensate for the expected aerodynamic pressures and

resonating frequencies at launch.

Uneven expansion

or deflection of

the gun barrel

Risk of the barrel deforming due

to temperature, wear or structural

effects, resulting in a failed

launch.

2 3 6

The gun barrel will be designed with insulation to mitigate the

effects of uneven temperature. A suitable support structure

will be designed to ensure the structural strength and accuracy

of the barrel

Interruption of

Communications

Risk of loss or interruption of

communications during launch

due to plasma and ionising gas

effects.

2 5 10

Antennas will be shielded from ionising gases to allow

communications. The projectile will be capable of automated

flight and trajectory control

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Adverse Tectonic

Activity

Tectonic activity in the local area

with the potential of disrupting a

launch.

4 4 16

The launch site facilities and foundations will be designed

using a suitable safety factor, to survive an earthquake of

typical regional magnitude.

Mission Risk – In Orbit

Collision with

debris or other

satellites in orbit

Risk of a collision between the

projectile or useful payload, and

in orbit satellites or debris

5 2 10

The trajectory of launch will be set before launch to avoid in

orbit bodies. Any significant deviation from this predicted

course will result in a mission abort.

Solar Radiation

Risk of disruptions to the mission

from solar radiation, solar flares

and coronal mass ejections

3 4 12

The activity of the sun at launch will be checked to ensure a

low solar activity. The projectile will include basic shielding

to prevent significant disruption from solar radiation

Orbital Decay

Risk of losing altitude in orbit

due to small but not insignificant

aerodynamic drag forces.

1 5 5 The decay of an orbit due to drag can be predicted therefore

necessary thrust input can be scheduled during the mission.

Project Risks

Launch Schedule

risks

Risk of causing disruption or

cancelation of a launch due to

project/construction delay

4 3 12 Agreed project management process and use of a Gantt chart

will be used to assure on time delivery of launch

Financial Risk

Risk of the cost of construction

and launch being greater than

predicted resulting in financial

loss

4 3 12 Agreed project management process and constant assessment

of cost will be used to assure on budget delivery of launch

Human Risk

General risks caused by poor

workforce methods during

construction and launch

5 3 15

A strict adherence to local/accepted health and safety laws to

ensure a safe construction and launch process. Project specific

methods for human work and the use of machines will

minimize human errors.

Construction

Risks

Risk of delivery and construction

of a not to specification launch

system

4 2 8

During manufacturing and construction of the gun, set

tolerances of parts will be followed and measured to assure a

within specification launch system.

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4. Project Organisation & Management

The project is broken down in to smaller sections according to the strength of group members

which come from different engineering disciplines. Table 7 shows the responsibility of group

members.

Table 7: Group member allocations

Group member Responsibilities

Aidan Cookson Projectile design.

Jack Feakins Control system and actuators.

Kin Man Benjamin Lee Sensors and communication system.

Sophie McElhill Launch systems design.

Alex Oses Trajectory simulation, MATLAB analysis.

Eleri Williams Launch location analysis, launch site design and access.

Formal and informal meetings are held regularly to update the project progress with formal

meetings taking place every Tuesday with three academics. A meeting agenda is provided

before the formal meeting and minutes are produced within 48 hours after the meeting.

Individual and group weekly goals are setup and reviewed after the formal meeting in order to

keep the project on track. Current Gantt chart with the latest progress is shown in Figure 9.

Supergun Group 1

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Figure 9: Gantt chart of the project.

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eet.pdf

Appendices

Appendix A

Declare constants.

Initialise variables

Run first loop from simulation start time to end time

Get max(Altitude)

For max(Altitude) > (1+X)*RequiredAltitude && max(Altitude) > (1-X)*RequiredAltitude

If max(Altitude) > (1+X)*RequiredAltitude

Initialise variables again

Reduce muzzle velocity by Y%

Run loop

Get max(Altitude)

end

If max(Altitude) < (1+X)*RequiredAltitude

Initialise variables again

Increase muzzle velocity by Y%

Run loop

Get max(Altitude)

end

end

Get muzzle velocity

Get total expended energy and cost

Calculate discrepancy between maximum altitude and the required maximum altitude

Supergun Group 1

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Appendix B

Appendix C

Estimates used for MATLAB model

Variable Estimated Value

Frontal Cross-sectional Area, S 0.05 m2

Coefficient of Drag, CD 0.6

Gun Length 300 m

Total Mass 150 kg

Barrel Attitude from Vertical 0®

Muzzle Velocity 8340 m/s