Spiritus

Submission for Royal Aeronautical Society General Aviation Design Competition 2018

Phil Marsh

Contents

1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

1.1 Small jet aircraft 5

1.1.1 Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.1.2 Aircraft requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.2 Review of similar types 7

2 Sizing and mass estimates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.1 Initial sizing 9

2.1.1 Max take off weight calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.1.2 Wing sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.1.3 Tail sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.2 Layout detail 12

2.2.1 Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2.2.2 Wing section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2.3 Mass estimates 13

3 Aerodynamic Estimates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.1 Aerodynamic estimates 15

4 Stability and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

4.1 Stability and Control 19

5 Performance estimates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

5.1 Performance Estimates 23

5.1.1 Take-off distances . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

5.1.2 Landing distances . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

I Part Two

6 Design features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

6.1 Safety features 29

II Part Three

7 Cost Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

7.1 Costings 33

Bibliography . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

Articles 35

Books 35

webpage 36

inproceedings 36

1. Introduction

1.1 Small jet aircraft

Small jet aircraft, sometimes termed micro or mini jets, have been around for a long time. However,

many are based on aircraft originally designed for a piston engine, for instance the BD-5 [13] or

occasional prototype and experimental aircraft such as the Microturbo 200 [6]. The engines utilized

were generally converted auxiliary power units (APUs) which suffered from low power to weight

ratios and high fuel consumption. Somewhat rarer are examples such as Yves Rossy’s jetpack

[7] and the Colomban CriCri MC15 Jet[1] where up scaled RC model aircraft engines have been

utilized. For the private pilot wanting the thrill of flying their own jet the options have thus been

limited. With the advent of new engines in the 100kg – 150kg thrust range, such as the PBS TJ1000

(see panel), it may now be feasible to build practical small jets.

This conceptual design investigates what performance may be possible using these small engines.

The aircraft could be used to explore the possibilities of opening up a new market for small jets that

provides an option for owners/operators that cannot afford or do not require the capability of, what

is generally referred to, as a very light jet (VLJ) or personal jet. VLJs tend to be small scale down

business jets with 4 or more seats.

There does not appear to be any particular standard classification for low power jet aircraft therefore

in this proposal the term small jet is used as described by Sehra and Shin [10]

1 The Czech designed PBS TJ100 turbojet engine was developed for for unmanned aerial

vehicles, target drones and missiles. The TJ100 engine has compact design and has 6.6:1 power-

to-weight ratio. It produces up to 1,300N thrust with low fuel consumption for its category. It

has classic design - radial compressor, radial and axial diffuser, annular combustion chamber,

axial turbine and fixed output nozzle.[1] �

1.1.1 Considerations

The following design describes a small jet aircraft that could be constructed to explore the pos-

sibilities of using engines within the 100-150kg thrust range. Taking advantage of the CAA’s E

Conditions a test flight program would be undertaken. In addition, the program would become

a focal point to generate market research into the viability of opening a new market for small

jets. The aircraft is designed to be flown by a reasonably competent private pilot. The emphasis

will be on being fun to fly with a reasonable range for touring. It is intended that the aircraft

be constructed using composite materials which are more suited to the laminar flow wings and

compound surfaces required. It will be subsonic (although utilising small jet engines Stinton [11]

suggests the possibility of a transonic design using similar thrust!), stable and easy to handle in

all areas of the flight envelope. Equipped with a "glass cockpit" with full authority digital engine

control (FADEC) it will take advantage of light weight instrumentation and a degree of automation

that will reduce pilot workload.

1.1.2 Aircraft requirements

• Aircraft is of a conventional layout

• Max. Operating speed V mo = 300

• Stall speed V stall = 65 knots

• Range 500nm

• Twin PBS TJ1000 turbojet engines

• Twin tandem seating cockpit with canopy

1.2 Review of similar types

Small jet aircraft data and information from various sources was retrieved. The information was

collected in a spreadsheet and used to identify a starting point for the design and as a cross check

for gross errors in estimates.

Aircraft Engine Thrust (kg) We/Wo Thrust/Weight Max velocity VstallFLS Microjet PBS TJ-100 130 0.42 0.34 278 58Microjet 200 TRS18-46 220 0.57 0.21 250 70Microjet 200B TRS 18 -1 260 0.58 0.20 250 70Sipa 500 Turbomeca Palas 150 0.59 0.17 215 n.c.Le Variviggen TRS18-46 220 0.61 0.25 216 45Petit Canard TRS18-46 220 0.53 0.36 400 74NGT TRS18-46 220 0.60 0.32 280 n.c.Caproni C22J TRS18-46 220 0.63 0.19 286 75Sonex PBS TJ-100 130 0.50 0.29 249 50

Table 1.1: Data for similar types (Velocity in knots)

2. Sizing and mass estimates

2.1 Initial sizing

2.1.1 Max take off weight calculation

Equations 1 — Intial mass calculation. Using the following equation from Jenkinson and

Marchman [5] an estimate of the initial mass of the aircraft and therefore the sizing was made.

MTO =MUL

1− (ME/MTO)− (MF/MTO)(2.1)

Where ME =Empty mass, MTO =Max take o f f mass, MF =Fuel mass, MUL =Use f ul load

Using an empty weight ratio of 0.6 (from analysis of table 1), an initial fuel fraction estimate of

0.23 and a payload weight of 180kg (2 crew + payload) gives an initial take off weight estimate of

1059kg

2.1.2 Wing sizing

A wing loading of 100kg/m2 is used in order to keep the stall speed low. Aspect ratio (AR) is fairly

low at 6. This chosen for practical reasons such as providing enough space for retractable landing

gear and fuel tanks, to make construction light and simple and to give good roll rates. It also makes

it easier to fit in smaller hangers. Further wing sizing data is given in table 2.1.

Parameter Value

Wing reference area weight/wing loading = 10.59m2Wing span

√weight x aspect ratio (AR)= 7.97m

Wing taper ratio 0.6Mean chord 1.36m

Table 2.1: Wing sizing

2.1.3 Tail sizing

Measuring the horizontal and vertical tail moment lengths from the CAD drawings an estimate of

tail sizing using the following equations from Raymer [8] was calculated.

Equations 2 — Horizontal and vertical tail sizing.

SHorizontal =CHT

(cwingSWing

LHT

)(2.2)

SVertical =CV T

(bwingSWing

LV T

)(2.3)

Where;

SHorizontal = horizontal tail area

SVertical = vertical tail area

SWing = wing re f area

cwing = mean aerodynamic chord (MAC)

bwing = wingspan

LHT = horizontal tail moment

LV T = vertical tail moment

CHT = Tail volume coe f f icient (horizontal)

CV T = Tail volume coe f f icient (vertical)

The tail volume coefficients CHT and CV T were given typical values extracted from [9] A 5%

reduction to both vertical and horizontal areas was made due to the expected effectiveness of a

T-tail configuration.

Parameter Value

Horizontal span 3.61mHorizontal mean chord 0.8mHorizontal taper ratio 0.5Vertical span 1.13mVertical mean chord 0.6mVertical taper ratio 0.9

Table 2.2: Tail sizing

Figure 2.1: 3 view drawing

2.2 Layout detail

• Aircraft is of a conventional layout

• 2 x PBS TJ100 engines pylon mounted on rear fuselage

• T-tail mounted tail surfaces to clear jet exhaust

• 2 seat tandem cockpit with clear canopy

• Wing low/mid fuselage mounted

• Wing has no sweep and a 0.6 taper ratio

• Conventional retractable landing gear retracting into wing

• Oxygen system - portable or installed

The twin engines are mounted on pylons. This alleviates the requirement for any structural cutouts

on the fuselage to allow air intake or access for maintenance. Engines mounted on pylon have

the added benefit of allowing easy access for walk round checks and maintenance. In addition, it

presents a future opportunity to fit alternative engines of similar design and thrust without major

redesign. Possible future engines include: jet beetle and AMT Lynx [14][4]. The tandem cockpit

keeps the frontal area and wetted areas to a minimum.

2.2.1 Structure

The aircraft is designed to be built using composite materials. However, within the weight calcula-

tions no weight savings have been identified that using these materials may offer. Some literature

suggests that savings can be made however, Raymer [8] suggests that in light aircraft this is often

not the case. The cockpit is arranged so that the wing spar passes in front of and below the level of

the rear seat with the fuselage fuel tank below.

2.2.2 Wing section

The wing section used in the preliminary drawings is a 747A315 laminar section. It is deep

enough for the retracting landing gear and fuel tanks, has a low pitching moment and good stall

characteristics. Computer modeling predictions show its best lift/drag ratio at the design cruise

speed.

The rigging/incidence angle is set at 2.4 degrees with 2 degrees washout to ensure that the wing

stalls first at the root.

2.3 Mass estimates

A combination of statistical, historical, component selection and structural analysis methods as

described by Raymer[9] were used to make weight estimations for the various components. An

overview is shown below in table 2.2.

The estimated weight, without fuel, of around 735kg includes an empty weight allowance of 10%.

Using the initial figure of 1059kg MTO suggests that the fuel fraction can be increased up to 0.30

giving greater range. From the given available engine data, SFC 1.0983 at max thrust and anecdotal

operational data of 30-32 gph at max thrust and 15-16 gph at max cruise [12], it is likely that the

initial fuel fraction was set too low to provide adequate range.

Component Mass/kg

Structures 350Propulsion 63Equipment 9410% Weight allowance 50Total empty weight 557Useful load 502MTO 1059

Table 2.3: Mass estimates

Figure 2.2: Concept rendering

Figure 2.3: Concept rendering

3. Aerodynamic Estimates

3.1 Aerodynamic estimates

Measurements were taken from the CAD drawings and used to produce the following table.

Parameter Value

Take off mass MTO 1059 kgMeasured wing reference area (S) 9.9m2

Wing loading 1047 N/m2

Wing aspect ratio (AR) 6.4Wing sweep 0Wetted area (SW ) 41.6m2 calculated using CAD area analysis

Table 3.1: Aerodynamic measurements

The following standard equations taken from Raymer, Stinton and Gudmundson were used to

provide aerodynamic estimates. [3] [9] [11]

Equations 3 Standard aeronautical equations

Induced drag f actor K =(

1πeA

)(3.1)

Oswald′s span e f f iciency f actor e = 1.78(1−0.045AR0.68)−0.64 (3.2)

Assumed skin f riction coe f f icient C f = 0.004 f rom Raymer (3.3)

Pro f ile drag coe f f icient CD0 =CF

(SWS

)(3.4)

Dynamic pressure q = 1/2ρV 2 (ρ air density, V velocity) (3.5)

Li f t =CLqS (3.6)

Drag =CDqS (3.7)

Li f t/drag ratio = L/D =CL/CD (3.8)

Lift and drag parameters were determined for a range of aircraft speed and altitude. Initially

calculations were for the aircraft in clean configuration producing max continuous thrust. Thrust at

10,000 and 20,000 feet was estimated using a graph from [3] showing typical thrust ratio against

sea level thrust for different altitudes. The produced data suggests that the aircraft is capable of a

max speed of around 275 knots.(figure 3.1)

Figure 3.1: The Drag Polar

Figure 3.2: Rate of Climb

Best rate of climb at max takeoff weight is around 2000 fpm which occurs at around 180 knots.

Using the available limited max thrust of 1300N and including the extra drag imposed by having

the gear down gives an engine out climb rate of around 1000 fpm.

Stall speed was estimated at 65 knots using Eq.(3.9)

Equations 4 — Stall speed.

VSO =

√2WTOqSCL

(3.9)

WhereWTO = take o f f weight , S = wing re f area and q = dynamic pressure

CL =Coe f f icient o f Li f t (using f lapped measurement)

Figure 3.3: Concept rendering

4. Stability and Control

4.1 Stability and Control

With reference to the CAD layout drawings, the location of components were measured. This

allowed moments to be calculated based on the weight estimates. From this data, the center of

gravity was located. To find the stick free neutral point, the following simplified equations from

Raymer were used.[8]

Figure 4.1: Concept rendering

Equations 5 — Static margin calculation.

Neutral Point : Xnp =C LαXwing−KFuselageterm + KTail term Xtail

CLα +KTail term(4.1)

Where :

C Lα[

cos(sweep)10+18 cos(sweep)/A

]Xwing = location o f quarter chord o f wing MAC

Xtail = location o f quarter chord o f horizontal tail MAC

KFuselage term = (0.2125L2ratio−0.0675Lratio +0.011)W 2f uselageL f uselage

Sre f

W 2f uselage = max f uselage width;

L f uselage f uselage length

Lratio = distance f rom f ront o f f uselage to 25% o f wing root chord/L f uselage

KTail term = 0.95 StailSwing CLα tailKdownwash[1− Selevator2Stail

]Pitch (longitudinal) stability checked by f inding the static margin

Static margin; SM =Xnp−Xcg

C(4.2)

Where; C =(

23

)Croot

1+λ +λ 2

1+λ

Using the above equations and extracting data from the CAD drawings the static margin was

calculated at approximately 12%. This should provide a reasonably stable and easy to fly aircraft:

however, it is at the sportier end of static margin so should have responsive handling. The figure

4.2 shows Static Margin (S.M.) against weight. Flying single seat and any changes in fuel load

does not have any dramatic effect on stability. Calculations for CG at the extremes suggests static

margin would stay within 12% -18%

Figure 4.2: Loading effect on static margin

5. Performance estimates

5.1 Performance Estimates

5.1.1 Take-off distances

Take off roll and over a 50ft obstacle were calculated using equations from Raymer [9] and

Gudmundsson [3]

Equations 6 — Take off distance.

Ground roll distance (SG) =(

12gKA

)Ln(

KT +KAVf 2

KT +KAVi2

)(5.1)

Where; KT =(

TW

)−µ (T hrust terms)

KA =ρ

2(W/S)(µCL−CD0−KC2L

)(Aerodynamic terms)

µ = rolling f riction coe f f icient (hard runway)

Vf = Final velocity (takeo f f velocity)

Vi = Intial velocity (0)

SR = Rotation time

Climb angle θclimb = sin−1(

TW− 1

L/D

); Where T = T hrust (5.2)

Transition radius R = 0.2156 V 2S1; Where VS1 = stalling speed without f laps (5.3)

Transition distance ST R = 0.2156 V 2S1

(TW− 1

L/D

)(5.4)

Transition height hT R = R(1− cosθclimb) (5.5)

Climb distance SC =hobst −hT Rtanθclimb

(5.6)

Ground roll distance was calculated at 374m (1228ft) with take off over a 50 foot obstacle at 635m

(2082ft).

5.1.2 Landing distances

Landing distance is found in a similar way using equation (5.1) taking 1.1 x Stall speed as the initial

velocity and 0 as the final velocity to calculate braking distance SBR . To this, the distance covered

before applying brakes SFR is estimated using an application delay time of 1 second. This gives a

total distance of 223 meters (733 ft).

The distance necessary to land and come to a complete stop over a 50ft obstacle was calculated

using the following equations from Gudmundsson.[3]

Equations 7 — Landing distance over 50ft obstacle.

Flare height hF = 0.1512 V 2SO(1− cos θapp) (5.7)

Where; θapp = glide path in degrees,VSO = stall speed in landing con f iguration

Approach distance SA =hobst −hFtan θapp

(5.8)

where; hobst = obstacle height(50 f t)

Distance covered during f lare SF = 0.1512 V 2SO sin θapp (5.9)

where; hobst = obstacle height(50 f t)

Calculating these distances and adding to the distance traveled during free roll SFR and braking roll

SBR gives a total landing distance SLDG over a 50 ft obstacle of 510 metres (1680 ft).

Predicted take-off and landing performance is comparable to high performance single and light

twin piston engined aircraft.

Range and endurance

The manufacturer quote the specific fuel consumption (SFC) for max thrust as 0.112 kg/N/h for the

PBS TJ100. Further information in the form of a graph [2] plotting thrust and SFC against engine

speed was obtained. Using this data, the weight of fuel that can be carried and the multiplying SFC

x hours x thrust [11] allows an estimation of range and endurance to be calculated. Graphs were

plotted for range of speeds at S.L., 10,000 and 20,000 feet. As a reality check, real life consumption

figures, obtained from a flight test [12], show 30gph at max thrust and 15gph at max cruise. This is

in line with the calculated values. From range and endurance graphs, it can be seen that the best

range can be obtained at a cruise speed between 175 and 200 knots. With a calculated fuel capacity

of 324kg (390 Litres), this would give a theoretical range of over 750 nautical miles at 20,000 feet.

Figure 5.1: Range at different altitudes

Figure 5.2: Endurance at different altitudes

I

6 Design features . . . . . . . . . . . . . . . . . . . . . . 296.1 Safety features

Part Two

6. Design features

6.1 Safety features

Although not ideal, the fuselage fuel tank is located below the main spar and underneath the rear

seat. Between the tank and the cockpit floor there is a firewall constructed of fire resistant fiberglass.

Safety is further enhanced by the bulge shape of the fuselage which allows the addition of a strong

keel to protect the aircraft in the event of a wheels up landing.

The engines are mounted on pylons which keeps these hot parts well away from fuel tanks. If

they become detached in a crash landing situation they are also unlikely to penetrate the fuselage.

Canopy frames are reinforced to provide roll over protection and a cross beam situated behind the

front seat serves to strengthen fuselage cutout and provide attachment for 5 point harness.

For pilot comfort, the seats are semi reclined at an angle of 25 degrees with a 109 degree trunk/thigh

angle. In order to absorb impact forces in heavy landings or a crash, the seat cushions are made of

a sandwich of hard grade low resilience foam on the bottom of the seat with a 12mm medium grade

on top for comfort. A seatbelt airbag system is incorporated into the design.

Forward of the cockpit the composite skin has an aluminum substructure with v shaped folds design

to initiate crumpling. The cockpit itself is formed of carbon composite for stiffness and deformation

resistance.

II

7 Cost Analysis . . . . . . . . . . . . . . . . . . . . . . . . 337.1 Costings

Bibliography . . . . . . . . . . . . . . . . . . . . . . . . 35

Articles

Books

webpage

inproceedings

Part Three

7. Cost Analysis

7.1 Costings

Although difficult to accurately cost a new design concept, reasonable estimates for a one off

aircraft can be made by looking at similar aircraft. The costs of major components and equipment

such as the engines and avionics can be obtained from manufacturers. A reasonable assumption

would be around £400-£450k. Taken in context the cheapest VLJ are in the low million pounds

mark.

Component Cost

PBS TJ100 $55,900 ea.Airframe £250,000Avionics £50,000Oxygen system £1000Seatbelt airbags £5000

Table 7.1: Representative costings

Figure 7.1:

Figure 7.2:

Figure 7.3:

Bibliography

Articles

[7] James Randerson. “Jet man ready for cross-Channel attempt after ’awesome’ test flight”.

English. In: The Guardian (Aug. 2008) (cited on page 5).

Books

[3] Snorri Gudmundsson. General Aviation Aircraft Design : Applied Methods and Procedures,

, . Available from: ProQuest Ebook Central. [14 February 2018]. English. Saint Louis:

Elsevier Science, 2013. ISBN: 9780123973290 (cited on pages 15, 16, 23, 24).

[5] Lloyd Jenkinson and Marchman. Aircraft Design Projects for Engineering Students. Butter-

worth Heinemann, 2003. ISBN: 0-7506-5772-3 (cited on page 9).

[8] Daniel P. Raymer. Simplified Aircraft Design For Home Builders. English. 1st. Los Angeles,

CA.: Design Dimension Publishing, 2007. ISBN: 978-09722397-0-7 (cited on pages 10, 12,

19).

[9] Daniel P. Raymer. Aircraft Design: A Conceptual Approach. 5th. Reston, VA: AIAA Educa-

tion Series, 2012. ISBN: 978-1-60086-911-2 (cited on pages 10, 13, 15, 23).

[11] Darrol Stinton. The Design of the Airplane. English. Oxford: BSP Professional Books, 1983.

ISBN: 0-632-018771 (cited on pages 6, 15, 25).

Websites

[1] Colomban MC-15J CriCri Jet. URL: https://minijets.org/fr/0-100/amt-olympus/

colomban-mc-15j-cricri-jet/ (cited on pages 5, 6).

[2] desertaerospace | TJ-100. en. URL: https://www.desertaerospace.com/tj-100 (cited

on page 25).

[4] Horace Ho. Jetbeetle–Affordable Micro/Mini/Small Jet Engines. URL: http://jetbeetle.

com/ (cited on page 12).

[6] MICROJET 200 et 200 B. French. URL: https : / / minijets . org / en / 100 - 150 /

microturbo-trs18/microjet-200-200b/ (cited on page 5).

[12] SubSonex Sport Jet JSX–2. Text. May 2015. URL: https://www.aopa.org/news-and-

media/all-news/2015/july/pilot/f_sonex (cited on pages 13, 25).

[13] The FLS Microjet. English. 2017. URL: https://www.bd-micro.com/ (cited on page 5).

[14] Bennie Van De Goor. AMT Netherlands Lynx. URL: http://www.amtjets.com/lynx.php

(cited on page 12).

Conference papers

[10] Arun K. Sehra and Jaiwon Shin. “Revolutionary Propulsion Systems for 21st Century

Aviation”. English. In: National Aeronautics and Space Administration, 2003. URL: https:

//ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040001042.pdf (cited on

page 6).

https://minijets.org/fr/0-100/amt-olympus/colomban-mc-15j-cricri-jet/https://minijets.org/fr/0-100/amt-olympus/colomban-mc-15j-cricri-jet/https://www.desertaerospace.com/tj-100http://jetbeetle.com/http://jetbeetle.com/https://minijets.org/en/100-150/microturbo-trs18/microjet-200-200b/https://minijets.org/en/100-150/microturbo-trs18/microjet-200-200b/https://www.aopa.org/news-and-media/all-news/2015/july/pilot/f_sonexhttps://www.aopa.org/news-and-media/all-news/2015/july/pilot/f_sonexhttps://www.bd-micro.com/http://www.amtjets.com/lynx.phphttps://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040001042.pdfhttps://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040001042.pdf

1 Introduction1.1 Small jet aircraft1.1.1 Considerations1.1.2 Aircraft requirements

1.2 Review of similar types

2 Sizing and mass estimates2.1 Initial sizing2.1.1 Max take off weight calculation2.1.2 Wing sizing2.1.3 Tail sizing

2.2 Layout detail2.2.1 Structure2.2.2 Wing section

2.3 Mass estimates

3 Aerodynamic Estimates3.1 Aerodynamic estimates

4 Stability and Control4.1 Stability and Control

5 Performance estimates5.1 Performance Estimates5.1.1 Take-off distances5.1.2 Landing distances

Part I — Part Two6 Design features6.1 Safety features

Part II — Part Three7 Cost Analysis7.1 Costings

BibliographyArticlesBookswebpageinproceedings