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NUS High School of Mathematics and Scienc

Team Name: Per Ardua Ad Astra

Members:

Chia Song Zhi

Jeremias Wong

Le Minh TuNguyen Duy Long

Sanchit Bareja

Tan Je Hon

Mentor:

Mr Wong Chee Leong

PER

ARDUA

AD ASTRASINGAPORE SPACE CHALLENGE 2010REPORT


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ContentsChapter 1- Introduction and Background ..................................................................................................... 4

Introduction .............................................................................................................................................. 4

Mission Overview .................................................................................................................................. 4

Additional Customer Requirements for Space trip ............................................................................... 4

Flight Profile .............................................................................................................................................. 4

Unique Design Concept ............................................................................................................................. 5

Airframe Overview ................................................................................................................................ 5

Climb to 100km ..................................................................................................................................... 6

Why did we use an Aerospike? ............................................................................................................. 6

Live Video-Conferencing ....................................................................................................................... 6

Chapter 2- Design and Analysis ..................................................................................................................... 6

Airframe .................................................................................................................................................... 6

Area Ruled Nose Cone ........................................................................................................................... 6

Variable Geometry ................................................................................................................................ 7

Lifting body ............................................................................................................................................ 8

Canards .................................................................................................................................................. 9

Propulsion ............................................................................................................................................... 10

Stage 1 propulsion ............................................................................................................................... 10

Stage 2 propulsion ............................................................................................................................... 11

Reliability of the system based on previous tests carried out ............................................................ 11

Why do we choose the linear Aerospike engine ................................................................................. 12

Other possibilities we considered ....................................................................................................... 13

Payload Mechanism ................................................................................................................................ 14

Interior..................................................................................................................................................... 14

Avionics ................................................................................................................................................... 15

Materials ................................................................................................................................................. 19

Chapter 3- Safety Considerations ................................................................................................................ 19

Backup plan in case of total failure ............................................................ Error! Bookmark not defined.

Failure mode analysis .............................................................................................................................. 19

Chapter 4- Weight and Cost Breakdown ..................................................................................................... 21

Chapter 5- Conclusion ................................................................................................................................. 22

Appendix ..................................................................................................................................................... 23


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Critical Mission Requirements ............................................................................................................ 23

Engine Appendix 1.1 ............................................................................................................................ 23

Engine Appendix 1.2 ............................................................................................................................ 24

Aerospike Engine Appendix ................................................................................................................. 24

Bibliography ...................................................................................................Error! Bookmark not defined.

References ................................................................................................................................................... 32


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Chapter 1- Introduction and Background

Introduction

The advent of space age has brought tremendous technological advancements and improved the lives of billions. Its th

21st Century and man has already step foot on moon at least a 100 times. However, space access to common man is sti

very limited although recently, there has been a spike in the number of proposals for space tourism. Among them, w

would like to present one we have thought of ourselves from scratch.

Mission Overview

Our mission has to be capable of reaching 100km, do a free fall for two minutes (in order for the tourists to experienc

weightlessness), launch an auxiliary satellite and land back safely at the same base. The two main mission capabilities tha

we must be able to perform are firstly, to climb to 100km and secondly to do a free fall as these are the most importan

guiding criteria. In our design, we have carefully considered different aspects of the flight and chosen an innovative an

efficient design (named Horus) that aims to minimize wave drags by incorporating variable wings. Aiming to maximize th

customers' satisfaction, we have also installed live video conferencing features, which put additional constraints on ou

design. The critical mission requirements are reflected in the appendix.

Additional Customer Requirements for Space tripIn addition to the above mentioned technical requirements, the needs of the space tourist were also considered.1 Howeve

not all needs can be met due to technical constraints. As such, the additional requirements for our group are as follows:

1) To experience weightlessness for a longer time.

2) To maximize viewing area or add more windows.

3) To install equipments like telescope for stargazing.

4) To allow family members to get live video feedback of the flight (both interior to be connected to the customer an

exterior for them to enjoy to the view).

Throughout the designing of the plane, customer satisfaction and safety was kept as a top most priority.

Flight Profile

Our flight profile is straight forward. It consists of a few stages shown below with each stage having a different planform t

optimize flight dynamics.

Phase Description

Stage 1 - Initial Positioning

Preflight Preparation To provide enough power to initiate engine (Engine Control)

Takeoff To take off automatically (Navigation)

Ascent (S1) To monitor the flight characteristics profile (Communications)

Cruise To restart engine in case of engine failure (Emergency System)

Stage 2 - Suborbital Flight

Ascent (S2) To determine and navigate ideal flight path for suborbital transition

(Navigation)

Free Fall To withstand and diverge heat due to reentry (Environment Control)

Stage 3 - Approach & Landing

Approach To navigate flight path for approach (Navigation)

Landing To land in extreme ground and weather conditions (Navigation)

To detect a failed approach and to advice the pilot to abort (Navigation)

Missed Approach To determine the next possible window for landing (Navigation)


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Figure 1: Classification of phases of flight

Figure 2: Flight Path Projection

Unique Design Concept

The space plane we modeled consists of three new design features.

1) It is capable of having a wing sweep from 20 to 70 degrees. The main reason for its implementation is to reduc

wave drag caused by the Mach cone at Mach numbers above 1.

2) The second-stage propulsion consists of a single linear aerospike engine with 4 nozzles to provide thrust-vectorin

and altitude compensation allowing the rocket to perform efficiently over all heights.

3) A video-link on board the plane to provide live video conferencing for space tourists to allow the family members t

keep track of the mission from the ground station and allowing them to get at least a glimpse of the space and t

be attached to their family members at all times.

Many other planes were studied before we finalized with these features to implement. Our space plane is heavi

influenced from the X-15, the Russian Sukhois' and Virgin's SpaceShipTwo. Many of the aerodynamic concepts anmaterials used were borrowed from past planes to ensure sufficient reliability of the plane.

Airframe Overview

The airframe design revolves around 3 main concepts: wave drag reduction, the lifting-body and variable geometry win

planform. The nose section is designed as a Sears-Haack body to reduce bow shock and the resulting wave drag. Canard

are used for pitch control and roll control at high speeds (when the wings are swept back). The mid section is tapered at 2

degrees to the direction of flight, and the integrated wing mount allows the wings to be swept to the same angle, keepin

all airframe surfaces within the shock cone formed at the Mach 3 ascent and re-entry speeds. This prevents further shoc


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strong shock waves from forming, reducing wave drag. In addition, the variable geometry wing planform allows the craft t

be capable of high speeds while still maintaining sufficient low-speed lift. Engine nacelles are integrated into the side of th

fuselage to reduce form drag. The underside of the mid-section is ramped to allow the body to provide lift at high speed

The aft section of the fuselage merges with the aerospike engine.

Climb to 100km

The main concern regarding the ascent is the plane's thrusting capability and stability. It is not possible to use a single

propulsion system to achieve this objective, and therefore, we have divided the climb into two stages: from 0 to 15km an

15 to 100km. In the first stage we will be using a conventional jet engine while in the second stage we will use a new

developed rocket known as the aerospike rocket, from Lockheed Martin. For information on the jet engine chosen, refe

to the (jet engine section). For information on the aerospike engine, refer to the (aerospike section).

Why did we use an Aerospike?

The possible candidates for the 2nd stage propulsion were the ramjet, scramjet and other various rockets (hybrid, solid-fue

dual-mode or liquid-fuel) and they were chosen based on the thrust they are capable of providing and their level o

technological readiness. The ramjet and scramjet, although attractive choices, do not function in low-air densit

environments. Of the remaining rockets, we chose the aerospike rocket as it matched our mission requirements. It has th

capability to do thrust vectoring and hence solving the problem of stability during flight which other rockets are unable t

accomplish.

Live Video-Conferencing

To implement the live video-conferencing feature, we needed to add a video link capability to our avionics system. This wa

easily taken care of and smoothly implemented. The video link system has been employed on the space shuttles as we

making it a really reliable system.

Chapter 2- Design and Analysis

AirframeThe design for the airframe has been divided into three broad concepts area-ruled nose cone, variable geometry an

lifting body. Below, we explain each concept and its implementation in detail.

Area Ruled Nose Cone

Modern aircraft, both military and civilian, generally operate in the transonic flight regime, with airspeeds of about Mac

0.8. At such speeds, the compressibility effects of air become significant, and shock waves begin to form. One such shoc

wave forms at the bow of the aircraft, and is known as bow shock. When airflow is incident upon the craft, it is deflected

and if this deflection is too great, a detached bow shock is formed. If however the deflection is within certain limits, th

bow shock formed remains attached to the aircraft surface.

The detached bow shock greatly increases the drag on the aircraft. For this reason, modern aircraft are all high

streamlined to reduce the angle of incidence between the surface of the craft and the surrounding airflow in order t

reduce the drag caused by the formation of shock waves (wave drag) including bow shock. This streamlining of the aircra

is known as the area rule.

ConceptIn order to reduce wave drag, the area rule dictates that a body should change cross sectional area as

smoothly as possible. An ideal body that satisfies the area rule is the Sears-Haack Body, which experiences


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the minimum possible amount wave drag for a given set of geometrical conditions (i.e. volume, length,

largest cross-sectional area, etc.). The radius of a Sears-Haack Body is given by:

r(x) Rmax

(4x 4x2 )3/4

In our design, a Sears-Haack Body is used for the nose section of the craft to reduce wave drag.

VariableGeometryVariable Geometry Wing Planforms (colloquially known as swing-wings) have been employed in a number of supersonic

aircraft, such as the F-14 Tomcat, F-111 Aardvark and B-1 Lancer. These aircraft often require high lift at low speeds and

supersonic capability, two traditionally conflicting characteristics. However with the employ of swing-wings, low wing

sweep allows such aircraft to gain adequate lift at low speeds and high wing sweep enables such aircraft to travel at

supersonic speeds without adverse drag.

Concept

Swing-wings benefit our design in two ways.

Firstly, for flights in both high subsonic (transonic) and supersonic regimes, sweeping the wings changes in

the airfoil presented to the airflow, essentially decreasing the thickness to chord ratio. This reduces rate at

which the airflow changes direction, reducing strength of resulting shock waves and thus reducing wave

drag. This allows our craft to fly with less thrust, reducing fuel consumption and hence the resulting weight

of fuel.

Secondly, a shock wave forms at the bow of the aircraft at transonic and supersonic speeds known as bow

shock. This shock wave is formed at an angle to the free stream airflow known as the Mach angle, and is

given by:

sin11

M

where M is the Mach number of the surrounding flow. Behind this shock wave, the airflow is subsonic. By

sweeping the wings to keep them behind the shock wave (i.e. swept at the Mach angle), one ensures that

the flow over the wing is subsonic. This eliminates the need for a supercritical airfoil, which has poor low

speed lift characteristics. This allows our craft to perform well both in the supersonic and subsonic flight

regimes, eliminating the need for exceptionally long runways to accommodate high take-off and landing

speeds.

In addition, this allows our craft to operate in high-traffic airspaces at subsonic speed, whilst remaining

capable of attaining the high speeds of rocket flight.


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Design

Figure 3: Swing-wing mechanism. Several configurations were considered, the final being this space saving design.

Figure 4: Top view of the craft, showing wings at minimum sweep (left) and maximum sweep (right).

Lifting body

Relying on the wings alone for lift would require large wing area, necessitating either extremely long or broad wings. Overl

broad wings (low aspect ratio) are less efficient in subsonic flight and are disfavoured. Overly long wings have problem

with flexure when loaded, which would require the use of highly exotic materials to overcome, increasing production an

maintenance costs. In addition, using such wings would require stronger hydraulics and hinge mechanisms to sweep

drastically increasing the weight of the swing-wing mechanism.

Design

To overcome this, our craft utilizes a lifting body design to offset the lift demands on the wings. The craft has ramped lowe

surface to deflect air at high speeds downwards, providing lift in the manner of a wave-rider. The lifting body also produce

increased drag, due to the induced drag from its lift production, however this is easily overcome with our choice of je

engines and aerospike engine.


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Figure 5: Initial design with non-lifting body. Note the nearly symmetrical pressure distribution.

Figure 6: Revised design with lifting body. Note the asymmetric pressure distribution that provides lift.

Canards

Conventional aircraft utilize tailplanes for pitch control, and in some cases roll control. In such aircraft, a positive pitchin

moment is generated by a downward deflection of the tailplanes, which pushes the aft of the aircraft down, adding to th

weight of the craft. This is problem is exacerbated if the moment arm of the tailplanes (measured from the centre o

mass) is short, which would require a larger down-force to achieve the same amount of pitch. This larger down-forc

translates into a larger speed to compensate for the reduced lift at the high pitch angles during landing and take-off.

In our craft, with the large aerospike engine located at the aft of the craft, the centre of gravity is located far back. Thi

makes using tailplanes for pitch control highly disadvantageous, as it would greatly increase lift, and hence the speed

required for take-off and landing.

Canards on the other hand are located at the fore of the aircraft, and generate a positive pitching moment by producing a

upwards force. This contributes to the lift of the craft, reducing the take-off and landing speeds. Furthermore, in ou


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design, the moment arm of the canards much longer than if we had used tailplanes, reducing the amount of forc

necessary to produce a given pitching moment, reducing the required area of the control surface.

Feathered Re-entry

Our design is such that the aft of the craft produces a larger drag than the fore of the craft. This is achieved by utilizing the

Sears-Haack Body for the nose to reduce the drag at the fore, and using the swept wings and larger fuselage cross-sectiona

area to increase the drag at the aft.

ConceptFeathered re-entry refers to the natural stability of the of the craft on re-entry. This is achieved by increasing the

drag in the aft of the craft in order to control the attitude of the craft in flight, without the need for extensive fly-

by-wire control. This is similar in concept to that of a badminton shuttlecock, which increases the drag in rear to

ensure that the shuttlecock is always pointed in the right direction in flight.

Ramped Air Intakes (Jet)

Ramped air intakes are used in high speed aircraft to maintain the air intake velocity of the jets within operation limits. Thi

is achieved by restricting the amount of air entering the intake using moveable ramps, and expanding them into a large

cavity, hence slowing the air down. Below is an image of the air intakes we modeled.

Propulsion

Stage 1 propulsion

We decided on the stage 1 propulsion based on a simple criterion. Mainly, it has to be able to provide enough thrust t

accelerate the plane to 300m/s (take-off speed) on a 1829m to 2438m (6000ft to 8000ft) assuming a 13000kg plane an

also be able to overcome the drag at any height. Also, FAA requirements dictate that we need at least 2 engines for take-o

(Reference Engine 1).

At take-off, we need a minimum thrust of 38kN.(Engine Appendix 1.1) However, from the simulation we ran at Mach 0.7

our total drag was around 80kN (refer to data table 2). As such, the engine must be capable of providing a thrust of 80kN

From a list of shortlisted engines (Appendix Engine 1.2), we selected the engine based on its capability to sustain flight at a


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stages below 15km. The total excess thrust we have is around 3kN at the time when the plane experiences the maximum

resistance.

The final engine we decided to use is 2*GE CF34-3A which is capable of providing 82 kN.

Stage 2 propulsion

For the 2nd stage propulsion system of the space plane, we have chosen the RS-2200 VentureStar aerospike Engine which

utilizes a gas generator cycle and liquid oxygen (Oxidizer) and hydrogen (Fuel) as its propellants. Below is an image of the

schematics of the aerospike. The specification of a scaled down version of the engine we planned is shown below. Allcalculations can be found inAerospike Engine Appendix.

Figure 7: Schematic of aerospike nozzle (left) and the testing of aerospike (right)2

Engine Specificiations

Mass of Engine

(dry)

1200kg

Mass of Fuel 3000kg

Dimensions 3m X 3m (length and width), 4m perpendicular to the surface

Thurst Range Ideal model engine has a set thrust range of 125kN to 441.45kN

Throttle Range 27.5% to 100% of Combustion chamber pressure

(Pratt and Whitney set target to be 125kN to 1250kN, Average thrust reported is 1,914 kN at sea leve

Chamber Pressure 7.68X10^7 X M, where M is the mass rate flow of the engine

Mixture Ratio 1 liquid oxygen : 2 liquid hydrogen

Fuel Consumption 100 kg/s (Maximum thrust) to 27.5 kg/s (Minimum thrust)

Nozzle Area 0.92 metres square

Area Ratio 173

Isp 447.3/s(at vacuum), 381.6/s(at sea level)

Reliability of the system based on previous tests carried out

Its prototype, the XRS-2200 has been extensively tested (14 times) by its developers for a maximum burn time of 25

seconds. During the period in which it was official under development, it had undergone a total number of 73 tests and a

accumulated burn time of over 4000 seconds. Recently, Lockheed Martin has constructed a rocket with an engine similar i

design to the RS-2200 has been tested three times in flight with two successes and one failure.


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Why do we choose the linear Aerospike engine

We took a look at various propulsion systems and chose the Linear Aerospike to be our engine of choice. We chose thi

engine mainly for its increased efficiency over the conventional bell nozzle. A bell nozzle can only be designed to b

efficient at a particular height. Although the aerospike has reduced efficiency over all bell nozzles at the height they ar

designed for, the aerospike nozzles efficiency exceeds that of the bell nozzle at every other height thus the averag

efficiency of aerospike over all heights is higher than bell nozzle. The aerospike is also more streamlined than

conventional bell nozzle reducing the form drag on the plane.

Figure 8: Comparison of rocket groups

Secondly, the aerospike is designed in such a way that it has thrust vectoring capabilities. This eliminates the need for th

engine to be mounted on a gimbal. This is a reasonable reduction in our weight given the strict weight limit.

Figure 9: Demonstration of thrust vectoring capabilities3

The aeropike is smaller than a bell nozzle of comparable thrust and the engine can also be housed directly inside th

spike of the engine, reducing the materials and resources required to construct and maintain the engine.


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Figure 10: Aerospike and Bell Nozzle Comparison4

Other possibilities we considered

Pulse Detonation Engine

The first alternative we considered was the pulse detonation engine. The engine operates on the by repeatedly detonatin

the propellants instead of combusting it. The shock wave generated by the detonation is of greater pressure than th

expansion of gases through ordinary deflagration in a conventional engine, thus allowing it to theoretically operate mor

efficiently than a normal engine. Due to the nature of the engine, the detonation is self propagating and thus eliminates thneed for turbo-pumps to maintain the detonation, reducing the effective weight of the engine. However, this engine

hypothetical and has not been extensively tested or has an existing model used commercially. For most fuels, a certai

intake of air is required for detonation to take place. As our goal is to reach at height of 100 km, the atmosphere at tha

height is insufficiently dense to provide enough air to perpetuate the detonation.

Ajak engine (Magneto-Hydro Dynamic generator)

Another hypothetical concept we considered is the Ajak model which had reportedly being tested by Russia. It uses

particle beam in its nose to ionize the air in front of the plane, creating a field of plasma around the plane which supposed

reduces drag of the hypersonic aircraft. It uses a MHD (Magneto-hydro dynamic) generator to derive energy from the flo

of plasma. Part of the energy is used to reform the fuel by cracking a heavy hydrocarbon to maximize the energy derived

from the combustion, part of the energy is used to accelerate the flow of the exhaust in the end. As the engine is quit

complex and a large magnetic field is required in order to operate the drive. A huge amount of resources are required fo

the construction and maintenance of the engine. Also, the project is shrouded in relative secrecy, thus the reliability of th

engine is questionable. For the above two models, the added complexity in understanding the propulsion system limits ou

ability to model the system for our use. Thus, other less complex systems were chosen.

Rocket assisted ramjet/scramjet

The next probability we considered is the (Rocket and fan assisted) ramjet/scramjet engine. These types of engines us

atmospheric pressure to compress or ram the air before combusting the fuel. This form of combustion is highly efficien

and it also provides high thrust, with a hypothetical limit of mach 25. One of the major drawbacks of using such an engine

the high speeds in which it must travel in order to obtain sufficient pressure to compress the gases. Another concern is th

low air density at high altitudes, where there may be insufficient air to compress the fuel for combustion. Lastly, the engin

is ultimately an air-breathing engine, which is largely inapplicable at high altitudes where a sufficient amount air is no

present.

Conventional Rocket Engine

Lastly, we also considered the conventional rocket engine which is largely similar to our choice of the aerospike engine. Th

only difference in their design is the nozzle which the former uses the bell nozzle. As mentioned in the above paragraphs,

the aerospike is more efficient than a bell nozzle at all other heights than the height it was designed for. In order to


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capitalize on the strengths of the conventional rocket engine, a multi-stage engine would be preferable. However, as this is

a RLV (Reusable launch vehicle) using a multi-stage rocket would be inefficient.

Payload Mechanism

We have chosen the ACU 2624 as our payload launch system as it has the capacity to deliver a 10000-kilogram S/C even

with a centre of gravity higher than 5m given modest design of about 4.48m in diameter. The ACU2624 works basically like

a spring pushing the payload out of the space plane and at the same time, releasing the payload using pyrotechnical bolts.

A diagram of the mechanism is shown below.

Figure 11: Mechanism of ACU26245

This extends the limit of the size and weight of satellites that our spaceship can bring into orbit and hence allowing for

future uses for our space plane. Extensive test programs including Separation Tests, Random Vibration Test, Static Load

Tests, Release Tests, Static Load Test to Rupture and Friction Tests have been carried out on the ACU 2624 and it has

performed well overall. Furthermore, ACU 2624 has undergone various successful missions. It has been used twice on

Ariane 5 for the launch of PPF/Envisat-1 and XMM in 2000 and in Proton for the launch of INTEGRAL.

Interior

Figure 12: Interior Design of spacecraft


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Avionics

There are many things that the aircraft avionics system has to do. This can be further subdivided into these categories,

which will be described later.

Navigation

Electrical Power

Engine Control

Flight Control

Emergency System Environment Control

Communication and Telemetry System

The objectives of the communications and telemetry system are stated below.

To establish high gain communication with ground stations(s)

To establish communications with payload and assist in payload deployment

To provide emergency communications for internal and external problems

Different frequencies have different characteristics that are geared towards different aspects of communications. In

general, the higher the frequency, the more data the signal can carry but the larger the atmospheric attenuation. We have

chosen HF/VHF, S band and X band communications as a basis for the communications due to their characteristics.

Frequency SelectionFrequency

Band

Spectrum width Spectrum Characteristics

Ku band 12 18 GHz High bandwidth data rate communications

Severely limited by rain and atmospheric attenuation

X band 8 12 GHz Medium-High bandwidth data rate communications

Suitable for telemetry communications

S band 2 4 GHz Medium bandwidth data rate communications

Suitable for payload communicationsUHF 0.3 3 GHz Backup data /audio communications

Suitable for backup telemetry communications

VHF 30 300 MHz Reliable audio communications, universal reception

Suitable for ATC / Mission Control audio traffic during takeoff/approach

HF 3 30 MHz Low data transfer rate, universal reception

Suitable for audio and emergency radio traffic during suborbital flight

A design is of X-band based wide-angle antenna is suggested. According to literature, this antenna is capable of transmittin

data at 1Gb/s experimentally. This will provide for essential for ground monitored, spacecraft controlled payloa

deployment, video conferencing capabilities and scientific research. Furthermore, we propose the following telemetry lin

budget. Initially, we would also incorporate the S band and HF/VHF band communications using a dish antenna and tune

antenna wire respectively.

After calculations (appendix), we found out that the carrier to noise (C/N) ratio for HF/VHF communications is lower tha

9dB/Hz which is the minimum specification for effective radio communications. Also, we determined that S ban

communications could be optimized for uplink communications due to its high uplink C/N ratio.


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Type of DataFrequency

Selection

Uplink band

(MHz)

Downlink

Band

(MHz)

Bandwidth (Hz, kbps, Modulation)

Voice S band 2025 2050 2075-2100 500 kbps (PSK)

Video Conferencing

X band 8500 -9000 8500 -9500

10 - 20 Mbps (PSK)

Housekeeping &

Mission Critical

Data

50 - 100 Mbps (PSK)

Payload Data S band 2050 - 2075 2100-2120 20 - 30 Mbps (PSK)Figure __ - Final Communications Budget

Flight Control System

The flight control system for this spacecraft would follow the recommendations for the update of the space shuttle avionic

system. Fail operational/fail safe (FO/FS) system architecture, which deploys 4 systems for redundancy, is utilized. A FO/F

system will allow the flight control system to be capable of performing the operational mission after 1 failure and the saf

return of the spacecraft after 2 failures. Furthermore, built in test equipments (BITE) will be deployed to ensure th

operational integrity of the equipments and spacecraft.

Data between the various components will be channeled through 3 command data buses (L, C and R), each with its own se

of signal interfaces. A list of common signal interfaces is described in the appendix. The various data buses are responsib

for different areas of the aircraft. (mainly, the left wing, the right wing and the central body)

Figure ___ - Schematic for flight control system

Engine Control SystemIn the engine control system, parameters are obtained from the data bus and the engine output is managed. Th

parameters are described in the appendix. Hydrazine powered auxiliary power unit (APU) manages the activation of th

aerospike engine and the turbofan. This APU provides power to critical components in the case of a power failure and t

start the engines. We based the activation of aerospike and turbofan engines on the space shuttle engine. A mechanica

minimum of 30,000 rpm is required for the starting of the engines. (refer to appendix)


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Figure ___ - Schematic for engine control system

Hydraulic Systems

For the hydraulic systems, we would be adopting the electro-hydrostatic actuators that are used in many modern aircrafts

due to the low weight and energy requirements of the modules. The hydraulic system is split up into 3 components, namel

the Primary Flight Control, the Secondary Flight Control and the Utility System. The various components are summarized in

the table below.

Power Source Primary FlightControls

Secondary FlightControls

Utility Systems

Electro-Hydrostatic

Actuator (EHA)Electrical System

Rudders

Flaperons

Carnards

Wing Sweep

Airbrakes

Air Intake Ramp

Undercarriage

Wheelbrakes

Anti-skid

Parking Brake

Parking Brake

Nosewheel SteeringFigure ___ - Hydraulic Circuit Design

Electrical Systems

Power System

The turbofan is used to turn an alternator, which serves as the primary source of power. The system is maintained at 270

and 27 V which is the industry standard. Power is regulated in AC and DC to avoid interferences and the systm is protecte

from being underpowered by an electrical load management system (ELMS) which manages the power distribution aroun

the aircraft. A power budget analysis is provided in the appendix.


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Emergency Power Unit

The other source of power is derived from the various APUs, which are usually also emergency power units (EPUs). W

propose 3 EPUs powered by hydrazine, hydrogen fuel cell and lithium ion polymer to be used to provide electrical power i

case of engine failure. This is because the hydrazine EPU is based on a spontaneous reaction between hydrazine and wate

which also provides for a instantaneous high torque for engine start. The batteries are designed to keep the system runnin

for 15 minutes after power failure.

Electromagnetic Compatibility

Electromagnetic Compatibility (EMC) addresses the categories of electromagnetic emissions. They include electromagnet

interference (EMI), radio frequency interference (RFI), electrostatic discharge (ESD) and electromagnetic pulse (EMP).

multi prong solution is required to prevent the radiated emissions from critically influencing the systems integrity. (Refer t

appendix)

Electrical Harness Design

The electrical harness system would utilize the various aspects of the electrical system to tackle EMC issues. Firstly, a single

point multi point ground system will be used. The singe point ground system is used for power distribution while the

multi point ground system is used to facilitate digital signal transfers. Effective EMC is reduced by adequate partitioning an

the separation of noisy and quiet signals. Signal matching will be conducted in driver circuits to prevent noise and

unwanted harmonics. Coaxial cabling will be used for relaying high frequency signals and optical cabling technology will be

implemented to reduce EMC caused and induced by high frequency signals. Lastly, low gauge (large width; below 18 AWG)

wires are to be used to carry signals and for power distribution.

Navigation Systems

There are 4 types of navigation systems that are used in

Dead Reckoning (INS)

Radio Navigation (Ground Based Radio, GPS)


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Celestial Navigation

Map-Matching Navigation

We plan to use both dead reckoning and radio navigation as the main forms of navigation as it is the most accurate and ha

the highest update frequency. Accurate positioning can be determined using celestial navigation with reference to th

position of the sun, and other objects of interest. To improve the accuracy of the GPS navigation system, we propose

modification to the current GPS system and the utilization of interferometry of GPS signals.

Materials

Wings-Ti-6Al-4V

Ti-6Al-4V is widely used in the Aerospace for its wings due to its excellent combination of light weight, high strength, hig

toughness, and good corrosion resistance especially at cryogenic temperatures.6 Given the high loads that the wings mus

hold, the low temperatures that it will encounter during the mission and the tropical weather of Singapore, Ti-6Al-4V is

suitable choice for the wings of our plane. The cost for Ti-6Al-4V is about $398/kg.7

Canopy- Monolithic Polycarbonate

Monolithic polycarbonate has the similar optical clarity of glass is a lot tougher. Monolithic Polycarbonate is widely used t

make windows for both commercial and military airplanes and as such, we will be using it for the Canopy of our plane a

well.

Fuselage- Carbon fiber-reinforced epoxy

Carbon fiber-reinforced epoxy possesses high tensile strength, high stiffness, high fracture toughness and good resistanc

against corrosion, abrasion and puncture at a relatively low cost.8 It has been used extensively in aerospace engineerin

especially to manufacture load-bearing aerospace structures9 and hence, we have selected it as the material to be used

building the fuselage. The cost for Carbon Fiber-reinforced epoxy is about $788/kg.

Pivots- Boron Composite

Wing pivots require materials that have very high tensile strength and good resistance against corrosion and abrasion

Hence, there is a competition between carbon fiber composites and boron fiber composites. Though there are severa

types of carbon fiber composites in the market which have tensile modulus or strength exceed that of boron fibe

composites, boron fiber composites, however, have a blend of tensile and compressive properties that no carbon fibe

composites can match and hence, it is a better material to use for the pivots in the wing mount.

Aerospike-Molybdenum

The aerospike nozzle requires a material with an extremely high melting point and strength. Molybdenum was the only

choice of metal to use as it has been widely tested and been used in the nozzles of the Polaris missiles. 10

Chapter 3- Safety Considerations

Failure mode analysis

We applied the FMEA method to the different stages of flight to analyze the potential causes and effects of various

subsystem failures. Failure on the component level was not considered due to the high complexity of the systems. The

results are presented below.


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20

Scenario 1: Take-off Scenario 2: Failure before Ascent phase 2


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21

Chapter 4- Weight and Cost Breakdowns) No.

of

Units

Per

Unit

Surface

Area

(m^3)

Surface

Thickness

(m)

Per Unit

Material

volume

(m^3)

Density

(kg/m^3)

Per

Unit

Weight

(kg)

Subtotal

Weight*

(kg)

Material Rounded Per Volume

Cost

(SGD/kg)

Per Unit

Cost

Subotal

Cost

me

elage

Cone"

1 53.57 0.0010 0.05 1620.00 86.78 86.78 graphiteepoxy sheet

85.00 $950,000.00 $50,891.50 $50,891.50

Cone

py

1 8.06 0.0050 0.04 2457.60 99.04 99.04 Monolithic

Polycarbonate

100.00 $10,170.62 $409.88 $409.88

Cone

ture

1 - - 0.14 4400.00 600.00 Various 100.00 $0.00

Mount

"

1 466.66 0.0010 0.30 1620.00 486.00 486.00 graphite

epoxy sheet

750.00 $950,000.00 $285,000.00 $285,000.00

Mount

ture

1 - - - - 1800.00 Various 300.00 $0.00

$0.00

$0.00

rol

ces

$0.00

s 2 67.89 0.0010 0.07 1620.00 109.99 219.98 graphite

epoxy sheet

220.00 $950,000.00 $64,499.49 $128,998.97

rds 2 12.07 0.0010 0.01 1620.00 19.55 39.11 graphite

epoxy sheet

40.00 $950,000.00 $11,466.50 $22,933.00

cal

liser

1 18.19 0.0010 0.02 1620.00 29.47 29.47 graphite

epoxy sheet

30.00 $950,000.00 $17,280.50 $17,280.50

tures

- - - - - 50.00 50.00 50.00

$0.00

ulsion

spike

e 4 1.07 0.0100 0.01 2023.29 21.65 86.60 3D carbon

fiber in

carbon matrix

85.00

1 22.50 0.0100 0.23 2023.29 455.24 455.24 3D carbon

fiber in

carbon matrix

450.00

zer

opump

1 - - - - 260.00 260.00 - 260.00

opump

1 - - - - 350.00 350.00 - 350.00

bustion

mbers

4 - - 0.00 2023.29 3.00 12.00 3D carbon

fiber in

carbon matrix

15.00

ellaneous(Piping,

es, Injectors)

50.00 50.00 50.00

ngine 2 - - - - 737.09 1474.18 - 1450.00

nics

r to

ics

on

pter 2-

1 - - - - 771.00 571.00 - 750.00 $500,000.00


22/33

22

nics)

r Body

Mount 4 - - 0.00 2380.00 7.16 28.62 Boron

Composite

30.00

ators2 - - - - 100.00 200.00 Boron

Composite200.00

Pistons 2 - - - - 50.00 100.00 Boron

Composite

100.00

rds

os

2 - - - - 20.00 40.00 - 40.00

n

ors

- - - - - - 200.00 - 200.00

oad

h

anism

1 - - - - 100.00 100.00 - 100.00

ng 1 - - - - 500.00 500.00 - 500.00

0.00

0.00

uel - - - - - - 2914.78 2900.00

d

ogen

- - - - - 176.47 176.47 Liquid

Hydrogen

180.00

d

en

- - - - - 2823.53 2823.53 Liquid Oxgen 2850.00

Gas - - - - - 50.00 50.00 Inert Gas 50.00

oad 1 - - - - 200.00 200.00 200.00

ans 2 - - - - 100 200.00 200.00

Dry 8038.01 8.038010965 6455.00 Cost $1,005,513.85

Wet 14002.79 14.00279453 12435.00

Chapter 5- Conclusion


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23

Appendix

Critical Mission Requirements

The mission critical requirements are as follows (taken from the competition guidelines):

1) The maximum take-off weight of the spaceplane must not exceed 12,500 kg.

2) The lifespan of the spaceplane has to exceed 10 years.

3) The design cost of the spaceplane must not exceed S$500 million over 10 years.

4) The spaceplane must be able to support a minimum of 2 passengers (inclusive of Pilot), or a maximum of 3passengers (inclusive of Pilot).

5) The spaceplane must sustain a total mission time of no less than 45 mins (+5 mins).

6) The spaceplane must takeoff from a runway length of between 6,000 to 8,000 feet.

7) The spaceplane must attain a minimum altitude of 100km and a maximum altitude ceiling of 102km.

8) The spaceplane must be designed with a payload delivery system to deploy a microsatellite not weighing more tha

200kg.

9) The spaceplane must land with at least 1 engine on: Powered landing.

10)The spaceplane is designed to take-off from a commercial airport. Hence considerations of busy civilian/military ai

traffic must be taken into account. This might determine the optimum time of departure for the spaceplane.

11)The spaceplane should have a realistic flight profile that includes: a) Take-off, b) Initial Atmospheric Climb, c)

Rocket burn trajectory to suborbital position, d) Re-entry, e) Landing.

12)Maintenance, repair and overhaul (MRO) of the spaceplane must be able to be carried out by existing MRO players

in Singapore.

13)The spaceplane structure and integrity must be able to sustain the rigors of suborbital flight (e.g. stress patterns,

heat, pressure, radiation etc).

Engine Appendix 1.1

The problem of selecting the first stage engine can be simplified as follows.

1. Thrust required for take-off assuming take-off velocity of 300km/h = 84m/s.

2.

Runway length = 6000ft to 8000ft = 1829m to 2438m

Writing our horizontal force equation during take-off,

, = = In this case, m is changing as fuel is being consumed at a variable rate and f is also changing with time as lift changes the

Normal force acting on the wheels. Solving this equation fully is not entirely possible as the exact function to model and is not known. However, we can easily get an overestimation by letting = = 13000and = 2 . Solvingthe equation with these assumptions,

=300

=84

= 1122 = 1829 = 112

2 = 2438 = 222

2

=> = 2 = 1.447 1.5 = 1 = 1.929 2.0Solving for ,

, = 37622 , = 50154


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24

Hence, the criterion to select the engine is:

> 38000Also, we want the combined weight of the engine and fuel to be as low as possible assuming a 1 hour 15 minutes

operational time. We chose the operation time to be 1 hour 15 minutes as it is a safe estimate of the upper limit of the

engine run time as our actual planned flight requires a run time of only 50 minutes.

However, this is not the only thrust requirement. Through the CFD analysis (data table 2), our maximum drag is 82kN. As

such, our engine must be capable of providing a thrust of 82kN.

Referring to data table 1, the best engine based on fuel consumed and thrust is 2*GE CF34-3A.

Aerospike Engine Appendix

For the modeling of the prospective aerospike engine, several assumptions were made beforehand:

No energy is lost as heat or through friction throughout the engine.

Assuming the efficiency of the engine is 0.9.

The reaction between hydrogen and oxygen is not in equilibrium and is complete, implying that the reaction will

always proceed in one direction and not halfway or backwards. (This prevents the formation of additional

compounds which are difficult to model and account for.) The area expansion ratio of the nozzle is 173 (Exit nozzle and aerospike nozzle), which is the same as that of the RS

2200 engine. (This allows to find the thrust provided by the nozzle without accounting for more of the complex

effects and the size of the flow field.)

The flow throughout the engine is laminar and no compression waves are formed. (It greatly simplifies calculationsin finding the thrust due to the aerospikes Spike.)

The expansion of the gas is adiabatic.

The gas (water) behaves like an ideal gas.

The two streams of exhaust converge at some distance from the truncated end of the aerospike.


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25

1) Engine model (The aerospike engine is circled in red)


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26

2) Calculations

Legend:

: Mass rate flow of the engine: Gas constant (8.314) : Molar mass of water (18 g mol1) : Heat capacity of gas at constant pressure of water (2.0785): Ratio of heat capacities at constant pressure and volume of water (9

7)

0: Gibbs free energy for formation of water (228.582 kJ mol1) : Temperature of combustion chamber0: Temperature of reactants before combustion : Pressure of combustion chamber : Pressure at the spike of the aerospike : Pressure at the throat of the exit nozzle : Pressure at the truncated base of the aerospike0: Pressure of reactants before combustion : Pressure of gas at exhaust : Ambient pressure

: Exhaust velocity of exit nozzle

Sound : Speed of soundn : Speed of gas at part of the engine: Mach number of exhaust : Area of exit nozzle (0.922) : Area of the truncated base : Cross-sectional area at part of the engine : Thrust


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27

0 1000000 = ( 0)By assuming no loss in energy,

2 = ( 0)Resulting in the following expression after taking into consideration that the expansion of as through the nozzle is

adiabatic.

2 = 2 1 (1 ()1 )

Consequently, the expression for the cross sectional area of the nozzle is:

= 2 1 ()

2(1 ()1 )

The ratio of the pressure at the throat of the exit nozzle to the combustion chamber pressure is:

= 21 + 1

Consolidating everything, the thrust of a general bell nozzle rocket can be expressed as:

= +( )The bell nozzle rocket is most efficient at only one particular height in which it was defined for, where = . The thrustequation for the aerospike model can be expressed through modifying the thrust equation for the bell nozzle rocket, where is the angle the exit nozzle makes with the axis which is normal to the plane made by the truncated base.

= + cos+(Centerbody ) + ( ) Of which the base pressure can be approximated by:

= (M)( 21 +

1)(0.05 +

0.967

1 + 1

22)

Which was determined empirically by Fick, M. and Schmucker, R. H. in their journal entry "Performance Aspects of Plug

Cluster Nozzles" for a circular base, which is used to find the lower limit of thrust by rectangular truncated base. (The

following graphs are assuming the mass rate flow at maximum.)

With the assumption that the flowlines converge the original expression can be written as:

= + cos+(Centerbody

)+ M( 21 + 1)(0.05 + 0.967

1 + 1

22)

The force on the center-body can be found through calculating the change in momentum of the exhaust gases by the

aerospike. After applying the assumptions stated, the respective thrust for both a bell nozzle designed for maximum

efficiency at 15 and the aerospike can be calculated.


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28

As the space plane climbs, the ambient pressure will drop rapidly, causing the aerospike to rapidly outperform the bell

nozzle as can be seen from the graph.


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29

Diagrams of engine cycles. a) Gas generator cycle, b)Staged-combustion cycle, c) Expander cycle, d) Pressure-fed

cycle. Taken fromhttp://www.aero.org/publications/crosslink/winter2004/03_sidebar3.html

Engine cycle

Similar to the SSME, it will also utilize a gas generator cycle for the combustion of the fuel and oxidizer. This cycle was

chosen over the other cycles, the expander cycle, pressure-fed cycle and staged-combustion cycle as it is relatively simpler

compared to the others except the pressure-fed cycle. The pressure- fed cycle would be the simplest and most reliable as it

only involves using inert gas to pump the fuel and oxidant into the combustion chamber. However, one large drawback

would be that the pressure of the inert gas has to exceed the pressure of the combustion chamber which would be

immense, limiting the use of this cycle to propellants which generate a low combustion chamber pressure. Due to the

expected high combustion chamber pressures of the propellants used (Hydrogen and Oxygen), this cycle was not selected.

Among the others, the gas generator cycle is the only open cycle, which means that it discharges some form of working

fluid (Combustion product) is discharged overboard. Both the expander and staged combustion cycle are more efficient

than the gas generator cycle as they do not dump any propellant overboard, making the engine more efficient. However,

there is a theoretical limit to how much thrust an expander cycle engine can provide caused by the geometrical constraint

of the nozzle as it uses waste heat from the nozzle to power the turbine. Open cycle engines are less complicated than

closed engine cycles as it does not have to deal with the pressure when ejecting the working fluid into the combustion

chamber, allowing the use of thinner combustion chambers and eliminating the need for more elaborate piping to deal

with the hot fluid, resulting in an engine which is lighter compared to that if other engine cycles were used. This is an

important consideration given the strict weight limit which was set.

a) b)

c) d)
http://www.aero.org/publications/crosslink/winter2004/03_sidebar3.htmlhttp://www.aero.org/publications/crosslink/winter2004/03_sidebar3.htmlhttp://www.aero.org/publications/crosslink/winter2004/03_sidebar3.htmlhttp://www.aero.org/publications/crosslink/winter2004/03_sidebar3.html


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30

Composition of engine

As this is an engine which utilizes oxygen and hydrogen as its oxidizer and fuel respectively, it would be safest to use parts

of the Space Shuttle Main Engine (SSME) as components of the engine as it has already been tested and used in real space

missions. The SSME is similar to our engine as it also utilizes a Hydrogen and Oxygen engine, thus its components have bee

designed specifically for the combustion. Other than the aerospikes Spike, the interior of the engine would be relatively

similar to the interior of the SSME. Thus, using the components from the SSME would ensure the reliability of the engine

combustion process. In addition, parts from the SSME can be obtained from NASA as the space shuttle program is due to b

cancelled in 2010 with the completion of the International Space Station at a reduced cost, eliminating the need to

outsource the manufacturing of components as they can be obtained directly from NASA.

a) b)

The engine consists of the following components:

(1) Combustion chamber / Igniter

There are four combustion chambers in the entire engine, each connected to a nozzle. As the combustion of hydrogen and

oxygen is highly exothermic, the combustion chamber is required to withstand both high pressures and temperatures. Thecombustion chamber is modeled after the combustion found in the Booster stage for the H II launcher developed byMitsubishi Heavy Industries (Characteristic chamber length: 0.77671) as it also utilizes Hydrogen and oxygen as itspropellants to ensure sufficient volume (Total chamber volume: 0.0427 m3) for the fuel and oxidizer to sufficiently mix in

order for a smooth combustion to occur. The igniter is connected to the fuel injector.

(2) Piping

Similar to the combustion chamber, it would have to carry gases at very high temperatures and pressures.

(3) Valves (For thrust vectoring and controlling irregularities in flow)

(4) Injector

(5) Turbine / Preburner

(6) Turbopump (s)

The above components are adapted from the SSME. Refer to SPACE SHUTTLE MAIN ENGINE: THE FIRST TEN YEARS, by

Robert E. Biggs for more details.

(7) Nozzle

a) Rough sketch of Aerospike engine not including details. b) Rough 3D model of aerospike engine.


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31

Due to the design of the aircraft, a total exit nozzle area of0.92 m2 was used to obtain a high exhaust velocity of the

propellants.

8) Aerospike Ramp

As the physical specifications of the aerospike were not released, an aerospike was approximately optimized by modeling

an arc from the nozzle to the end of the ramp to prevent any one point from being heated more than other points.

Graph of the curvature of Spike at full length, the straight line indicates where the Spike is truncated.

Simulation of Aerospike during burn.

Thrust vectoring

Through the use of valves, the cross-sectional area of certain pipes can be restricted or altered. By assuming that mass rate

flow is conserved and the thrust from each nozzle is directly proportional to the amount of fuel and oxidant it receives, the

thrust vectoring capability of the engine can be estimated:

11 = 22Torque for roll and yaw axis: 0.75 + Sin

Maximum value: 97900 N m

Torque for pitch axis: 7.83 + SinMaximum Value: 1022100 N m


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32

Engine material

List of available material for engine material:

ATJ Modern

Graphite

Pyrolytic

Graphite

Three-dimensional

Carbon Fibers in a

Carbon Matrix

Carbon Cloth

Phenolic

Silica Clo

Phenolic

Density (kg m3) 1539.00244 2186.7121 1716.15 to 1992.95 1467.0347 1716.15Thermal Conductivity

(W m1) 49.84512 2.0353424 0.830753 to 8.7229 91.3828 45.691Thermal Expansion (m1) 0.000127 to

0.0001778

0.000036576 2.54 108 to2.286 107 2.03708 107 1.930 10

Modulus of elasticity (psi) 1.5 106 4.5 106 35 to 80 106 2.86 106 3.17 1Shear Modulus (psi) - 2 1 05 - 8.1 105 8 1 0

Erosion Rate (m s1) 0.0001016 to0.0001524

0.0000254 to

0.0000508

0.0000127 to

0.0000254

0.000127 to

0.000254

0.00025

0.0005

The material chosen for the composition of the engine are the Three-dimensional Carbon Fibers in a Carbon Matrix as it ha

the lowest erosion rate, reasonable density low thermal conductivity and high modulus of elasticity. The low erosion rate

increases the operational lifespan of the engine and given the strict weight limit, having a low density and strength of

material are important criterion in the selection of the material. The low thermal expansion allows the engine to retain its

shape and geometry under high temperatures, reducing the changes in thrust during burn time. Lastly, the low thermal

conductivity lowers the heat lost through conduction, allow a more efficient transfer of energy.

Data Tables

Data Table 1 (jet engines)nufacturer Model Application(s) Thrust SFC Fan Length Width/ Dry Fuel Weight Weight of

engine

Total

(dry) (dry) Diameter Diameter Weight Required 2*engine Weight W

[lbf] [lb/lbf

hr]

[in] [in] [in] [lb] [lb] [lb] [lb]

CF34-1A

Challenger 601-1A 8,650 0.36 44 103 49 1,625 6480 3250 9730 44

CF34-3A Challenger 601-3A 9,220 0.357 44 103 49 1,625 6426 3250 9676 438

CF34-3A1

CRJ100/200 9,220 0.357 44 103 49 1,655 6426 3310 9736 441

CF34-3B

CRJ100/200,Challenger 604,

Challenger 800

9,220 0.346 44 103 49 1,670 6228 3340 9568 433

CF34-3B1

CRJ100/200/200ER/200LR,Challenger 604,Challenger 800

9,220 0.346 44 103 49 1,670 6228 3340 9568 433

ls-Royce Spey Jr.RB.183-2Mk.555-15

F28 Mk.1000/Mk.1000C/Mk.2000

9,850 0.75 32.5 97 2,257 13500 4514 18014 817


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References

Every effort is made to identify the source of the information. All images and diagrams that are not referenced are self

generated.

1Webber, D. (2006). Designing the Orbital Space Tourism Experience. Space Technology and Applications International Forum (pp. 1041

1048). American Institute of Physics.2 http://green.myninjaplease.com/wp-content/uploads/2007/03/twin_linear_aerospike.jpg3

Adapted from Rocketdyne, 19994

Adapted from http://www.aerospaceweb.org/design/aerospike/x33.shtml, assessed 29th

April 20105Image taken from: Proceedings of a European Conference held at Braunschweig, Germany, 4-6 November 1998. Paris: European Spac

Agency (ESA), ESA-SP, Vol. 428, 1999, ISBN: 9290927127., p.136 and 1386Adapted from http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MTP64, assessed 23

rdMarch 2010

7Adapted from http://www.arcam.com/CommonResources/Files/www.arcam.com/Documents/EBM%20Materi als/Arcam-Ti6Al4V

Titanium-Alloy.pdf, assessed 25th

March 20108Adapted from http://www.substech.com/dokuwiki/doku.php?id=polymer_matrix_composites_introduction , assessed 8

thFebrua

20109Adapted from http://www.solarnavigator.net/composites/glass_fibre_reinforced_plastic.htm, assessed 23

rdMarch 2010

10Adapted from http://www.springerlink.com/content/p01l414425423565/fulltext.pdf?page=1, assessed 25

thApril 2010