Submersible Aircraft1

Submersible Aircraft

MECN3005

Joseph Thomas 0710343e

10/21/2011

Declatation

Executive Summary

The following are details on the design development of a submersible aircraft. The method to enter the water is to first land on the surface, undergo the necessary reconfigurations, and dive. The mission specifications are to have a range of 800 nm in the air, 50 nm on water and 20 minutes underwater. Three concepts were generated. Concept 1 was chosen to undergo design development on account of its simplicity and pilot friendliness. Analysis was done primarily on the structures required in the underwater phase. A decoupled power transition was used to power the underwater phase with an electric engine and batteries.

Table of ContentsDeclatation............................................................................................................................................. i

Executive Summary............................................................................................................................... ii

Nomenclature.......................................................................................................................................vi

List of Figures.......................................................................................................................................vii

List of Tables.......................................................................................................................................viii

Introduction...........................................................................................................................................1

Literature Review..................................................................................................................................2

Previous Work...................................................................................................................................2

Power plants......................................................................................................................................5

Shrenk’s Approximation....................................................................................................................7

Conclusion.........................................................................................................................................8

Product Requirement Specification.......................................................................................................9

Requirements....................................................................................................................................9

Constraints........................................................................................................................................9

Criteria...............................................................................................................................................9

Concepts..............................................................................................................................................10

Concept 1........................................................................................................................................10

Advantages..................................................................................................................................10

Disadvantages..............................................................................................................................10

Concept 2........................................................................................................................................12

Advantages..................................................................................................................................12

Disadvantage...............................................................................................................................12

Concept 3........................................................................................................................................14

Advantages..................................................................................................................................14

Disadvantages..............................................................................................................................14

Selection..........................................................................................................................................14

Design Development...........................................................................................................................16

Wing................................................................................................................................................16

Empennage Sizing............................................................................................................................17

Fuselage Sizing.................................................................................................................................17

Drag.................................................................................................................................................18

Engine..............................................................................................................................................19

Wing Load Distribution....................................................................................................................21

Load Factors....................................................................................................................................23

Take-off condition for the hull and main float.............................................................................23

Hull and float pressures...............................................................................................................24

Structures........................................................................................................................................26

Wing Structure................................................................................................................................26

Spar Position................................................................................................................................26

Rib Position..................................................................................................................................26

Shear Flow Analysis.....................................................................................................................27

Tail Structure...................................................................................................................................30

Spar Position................................................................................................................................30

Rib Position..................................................................................................................................30

Shear Flow Analysis.....................................................................................................................30

Hull Structure..................................................................................................................................32

Longerons....................................................................................................................................32

Ribs..............................................................................................................................................33

Seals.................................................................................................................................................34

Mechanical..................................................................................................................................34

Electric.........................................................................................................................................34

Propeller..........................................................................................................................................35

Secondary Power Source.................................................................................................................37

Electric Engine.............................................................................................................................37

Fitting...............................................................................................................................................39

Performance........................................................................................................................................40

Field Performance...........................................................................................................................40

Underwater Speed...........................................................................................................................41

Velocity Minimum Unstick...............................................................................................................42

Water Take Off................................................................................................................................43

Drag Polar........................................................................................................................................44

Mission Profile.....................................................................................................................................45

References...........................................................................................................................................46

Appendix A..........................................................................................................................................47

Appendix B..........................................................................................................................................51

Appendix C..........................................................................................................................................52

Appendix D..........................................................................................................................................55

NomenclatureSymbol Name Unit/ValueMTOW Maximum take-off weight KgTOW Take-off weight KgCOG Centre of gravityg acceleration due to gravity 9.81 m/s2

HP Unit of Horse powerc, MAC Mean aerodynamic chord mc(x) local chord at span wise position mδ(x) Local twist angle rad Angle of attack radCl Lift curve slopeCl∞ Ideal lift curve slopeV Velocity m/sρ Density kg/m3

b span mλ Taper ratiok CurvatureSref Wing area m2

Swet Wetted area m2

dA/dx Shrenk’s lift distribution N/mCl Lift CoefficientCd Crag CoefficientCdo Parasite DragCdi Induced Drage Wing Oswald’s factorAR Aspect RatioQ Shear Flow NmCr Root chord mCt Tip chord mt/c Thickness to chord tatio

List of Figures

Figure 1: Folding wing: US Patent 2444332...........................................................................................3Figure 2: Doors for the wings................................................................................................................3Figure 3: PB2Y Coronado [6]..................................................................................................................4Figure 4: Concept 1..............................................................................................................................11Figure 5: Concept 2..............................................................................................................................13Figure 6: Concept 3..............................................................................................................................15Figure 7: Top view, Tri Seating.............................................................................................................18Figure 8: Drag Bucket..........................................................................................................................19Figure 9: Thrust Curve.........................................................................................................................20Figure 10: Thrust and Drag at FL200....................................................................................................20Figure 11: Shrenk Load Distribution....................................................................................................21Figure 12: Shrenk Shear Diagram........................................................................................................21Figure 13: Shrenk Bending Moment....................................................................................................22Figure 14: Wing simplification.............................................................................................................27Figure 15: Statistically determinant shear flows..................................................................................28Figure 16: Compatibility Flows............................................................................................................28Figure 17: Skin Thickness Iteration for Aluminium..............................................................................31Figure 18: Idealised Stress in Fuselage................................................................................................32Figure 19: Ball Cock.............................................................................................................................34Figure 20: Fitting..................................................................................................................................39Figure 21: Drag in Water......................................................................................................................41Figure 22: VMU....................................................................................................................................42Figure 23: Combined Resistance with Thrust and Trim Angle.............................................................43Figure 24: Water Drag.........................................................................................................................43Figure 25: Cruise Drag Polar................................................................................................................44Figure 26: Mission Profile....................................................................................................................45

List of Tables

Table 1: Specifications of oxygen tank [10]...........................................................................................5Table 2: Data Input................................................................................................................................6Table 3: Horsepower Calculations.........................................................................................................6Table 4: Speed & Power Calculations....................................................................................................7Table 5: Propeller Size...........................................................................................................................7Table 6: Wing Characteristics..............................................................................................................16Table 7: Empennage Characteristics....................................................................................................17Table 8: Engine Specification...............................................................................................................20Table 9: Water Load Factors................................................................................................................25Table 10: Empennage Skin Thickness..................................................................................................30Table 11: Propeller Data Input.............................................................................................................35Table 12: Propeller output [14]...........................................................................................................35Table 13: Propeller Output [17]...........................................................................................................35Table 14: Electric Engine specifications...............................................................................................37Table 15: Battery and inverter specifications......................................................................................38Table 16: Field Performance................................................................................................................40Table 17: Drag Table............................................................................................................................47Table 18: Diameter Data......................................................................................................................47Table 19: Thrust Curve Data................................................................................................................48Table 20: Drag Data.............................................................................................................................49Table 21: Weight Estimation................................................................................................................49Table 22: Field Performance................................................................................................................50Table 23: Water Drag Data..................................................................................................................50Table 24: Load Facor Data...................................................................................................................51

Introduction

An aircraft is truly the master of its surroundings. It can travel in the thirds degree and makes intercontinental travel convenient. In our time of rising sea levels and disappearing habitat, there is an emerging market to explore the seas. A submersible aircraft would be perfect for such an application. It can be used from runway to reef in the fastest time (excluding the health and safety brief).

To make this a reality, there are some criteria that must be discussed. The type of engine is constrained to one that cannot be ‘blown out’. Ships and submarines use propellers for propulsion for good reason. The torque is transmitted from an internal power plant. This power plant must be air independent, or extra oxidiser must be carried.

The operating depth will be limited to 10 meters. This is to avoid crushing pressures, including the weight of the structure to protect from it. Secondly, beyond a certain depth, the visibility becomes limited and the novelty of the experience would wear out. The major design considerations are

The skin thickness needed to resist the external loads. The seals for the vents and ports (pitot, air ducts, general nacelles) Control and stability underwater Propulsion and power source underwater Take off on the surface Submerging from the surface General dynamics of an aircraft

Task as GivenStudents are to propose a topic on some form of aircraft. Each student will go into detailed design on their idea and present a report.

Task as UnderstoodFor my topic I chose a submersible aircraft. The requirements for the device are to have the capability of controlled flight in the air and under water. The following report details the is the design process for such an aircraft.

Literature Review

To have a vehicle with the ability to operate in any medium is highly attractive. Travel on land, air and water has been conquered separately. There have been numerous experimental aircraft attempting this, with limited success. An aircraft that can explore the skies and the depths of the seas is the focus of the following document.

The complexity of this problem is best described by a grant offered by DARPA on a ‘Submergible Aircraft’ [1]. A submersible aircraft would combine the key capabilities of three different platforms: 1) The speed and range of an aircraft 2) The loiter capabilities of a boat 3) The stealth of a submarine

Efforts in this field have had limited success due to the requirements of a submarine and an aircraft being largely different. Submarines are heavy with have thick skin due to the high pressures involved while diving. The engines of submarines are heavy and are independent of air. In contrast an aircraft has to be as light as possible for efficiency in the air. Therefor it has thin skin and engines with high power to weight ratios. The engines depend on air with their internal combustion engines. Flow conditions for the two mediums are different due to the density of water being higher by an order of magnitude. Submarines have small appendages designed to be submerged even when the sub surfaces. Amphibian aircraft have high mounted wing to stay clear of the water, in the ‘clean air’. The geometry of the lifting surfaces is therefore diametrically opposed.

Previous Work

There are limited accounts on working submergible aircraft. The discussions are limited seaplanes and the improvement that can be made to aircraft for them to submerge. The problems, outlined in the introduction, are mainly related to the weight and power. An additional issue is the drag that is experienced underwater being order of magnitudes larger than in the air.

There are accounts of submergible aircraft with folding wings. The aircraft will have locking wings in the extended and retracted position. The advantage in underwater drag is immediately apparent. The main control surfaces will be the vertical and horizontal stabilisers. Figure 1 illustrates a model with wings retracting inside the hull. An electric motor is proposed to operate the rotation of the wings. Figure 2 shows the doors of the wings. They are originally proposed to be controlled with springs and solenoids. A hydraulically actuated system will have to be investigated to provide the required operation and reliability. A variable geometry wing will have a structural weight penalty, but will be favourable in terms of the drag reduction under water. [2]

Figure 1: Folding wing: US Patent 2444332

Figure 2: Doors for the wings

Ekranoplans use the ground effect to glide just over the water’s surface at high speed and low drag. The proximity of the wings to the ground is to reduce the lift dependent drag. The ground blocks the trailing edge vortices and reduces the amount of downwash from the wing. The effectively increases the angle of attack and the wing creates more lift and less drag that it would otherwise at altitude. At an altitude of 1/10 wingspan, the drag is decreased by half [3].

As the aircraft approaches the ground, the cushion of air under the wing can cause the lift induced drag to decrease. As induced drag is a prime cause of drag at sea level (where there is high lift coefficient) the aircrafts performance dramatically increases. The theory of ground effect is summarised in the following formula.

∅=1−2eπ2ln(1+( πb8h )

2)Cd=Cdo+∅Cdi

In the above equations h represents the height of the wing from the ground. As the ratio of b/h increases, the reduction of the induced drag is multiplied. Ground effect can contribute to an increase in range and speed with a reduction in fuel flow.

To artificially create lift in the initial stage, the thrust from the engines can be channelled over the wing to lower the pressure. This can give sufficient lift to ‘take off’. The advantages of wing in ground effect (WIG) include longer range, higher speed, lower cost and higher transport efficiency than surface vessels. [4]

Wings that work in ground effect include:

Inverted delta Biplane: using main wings on shoulder and sponsons on the hull. Canard: Directing airflow under the main wind, initiating ground effect at lower speeds.

Between 1923 and 1962 Convair published an account of their seaplane programs [5]. It features commercial and military seaplanes and one account of a submergible aircraft. Their research has encountered and solved problems relating to aerodynamics, stability and control and general configuration.

Early designs, such as the TW-3 and NY husky models, incorporated floats on the wing tips and the fuselage. These were biplanes with the engine mounted on the nose, therefor significant clearance was needed. The extended floats ensured that the entire body of the aircraft was clear from the water. The single propeller engines ranges from a liquid cooled 180-hp to air cooled 240-hp manufactured by Wright.

The first mono-seaplane was the XPY-1 created in 1929. Nicknamed the Admiral it featured a U-shaped displacement hull and floats on the wings. It would land the main fuselage on the water, becoming a flying boat instead of an aircraft on stilts. It had a wing attached above with truss work and two 425-hp engines mounted between the fuselage and wing. The empennage had a twin tail configuration. Redesigned for the PB2Y Coronado shown in Figure 3, the twin tail alleviated problems with directional stability.

Figure 3: PB2Y Coronado [6]

Alternatives for the landing gear position were investigated. To save space in the hull, the landing gear was house in external, streamlined pods in the XP3Y Catalina. These pods were mounted between wing struts, allowing the landing gear to retract upwards and outboard. A slot on the side of the aircraft was proposed for stowing the aft landing gear for the Model 28 and 31. This saved space and eliminated the possible leakage issues associated with doors on the hull.

In December 1962 a submersible aircraft was proposed. It featured 3 turbojet engines, one for cruise and the additional two for take-off. Various water take-off methods were investigated including vertical and ground effect. The various underwater phase propulsion systems included propeller, pump jet, cyclic underwater gliding and buoyancy propulsion. To minimise the pressure vessel, all non-critical areas were flooded. A retractable hydro ski was proposed to facilitate take-off and landing. However the F2Y Sea Dart demonstrated that a twin ski was a more viable design.

Tilt wings where designed for various models in Convair. This was to provide clearance of the propellers, high lift (upward thrust) and slipstream effect over the entire wing. These variable incidence wings could tilt up to 30 degrees. This could be effective when submerged and an emergency surfacing is needed. The extra weight of structure may not be justified.

Power plants Submergible aircraft have the issue of being cut off from air. Internal combustion (IC) engines depend on air to burn fuel to produce power. There are two options to go forward: carry an oxidiser on board or have an alternate power source independent of air. Alternate power includes nuclear, fuel cell, electric and pump jet.

The power plants used in submarines are heavy. These include nuclear, internal IC engines and battery power. The IC engine runs the propellers and charges the battery on the surface. The battery can be weigh up to 400 tonnes. The sub must be in contact with the atmosphere by surfacing or use a snorkel to run the engines. Once the batteries are charges the sub can submerge, using electric motors to run drive the propellers. [7]

The underwater phase of the mission can be run using an electric motor. During flight, it can charge and provide drive for the motors and power of the electronics while submerged. The weight of an AC 150 electric motor is just over 80 kg. This provides a maximum of 200Hp, 19 Nm torque with a charger from 200 to 20000 Watts. [8]

Oxygen tanks are needed to breathing and continuous operation of IC engines under water. The recommended rate of oxygen supply is 3 litres per minute [9]. The oxygen requirements of a propeller engines depends on the SFC and the stoichiometric relationship between the air and fuel mixture. A standard oxygen tank has the specifications shown in Table 1. This will be more than enough for the journey for two passengers.

Weight full 15 kgHeight 71 cmDiameter 18.2 cmCapacity 2122 litres

Table 1: Specifications of oxygen tank [10]

Fuel cells convert energy from chemical to electrical. Unlike batteries it depends on a constant supply of ‘fuel’ and therefore does not lose charge over time. It provides a DC voltage to power the motors or electrics. The most common fuel cell converts hydrogen and oxygen into water. The drawbacks of fuel cells are that they take up large amounts of space for the energy requirements. Despite lightweight materials used, the weight is an issue considering there are two separate ‘fuels’ needed. [11]

Nuclear power is a form of clean, efficient energy. It produces a huge amount of energy with a small amount of fuel. Using uranium, water is heated to stream which intern runs a turbine. The drawback is that shielding is needed from the radioactive material in the core of the reactor. [12]

Nuclear adapted aircraft have been investigated by Convair. These aircraft used nuclear adapted turboprop and jet engines. A circulating fuel reactor was used as a heat source for the jet engine. The navy was directed to use a direct air cycle. This is where air is heated and pumped to drive the turboprop engine. There are not many technical details for these aircraft except the technical drawings. The diameter a select reactor, used in the Model 23, is about one meter and two meters in length. [5]

French born Gerard Thevenot has dedicated his life to hydrogen power. He was first honoured in 2006 when he won the Shell Eco Marathon. He set a record of over 4000 km s corresponding to an equivalent gasoline fuel consumption of 0.02343 litres per 100 km s. The was more recently awarded with the E-Flight award at Aero 2009 for his aircraft with an electric motor powered by a hydrogen cell. No battery is needed in this design, contributing to the aircraft’s weight of 130 kg s including the pilot. The 7KW motor weighs 55 kg and has a 5 litre tank of hydrogen. The fuel flow at 100m above sea level is 550g per flight hour with the emissions being pure water vapour. [13]

A pump just uses a centrifugal pump to create a jet of water for propulsion. A rudder placed in the path of the flow can control the direction of motion. Alternately an empennage can achieve the same effect. Torpedoes use similar technology (pump jets), with liquid fuelled engines.

The main propulsive power of modern boats and submarines comes from propellers. The following information in Table 3, Table 4 and Table 5 are calculations done for propeller specifications given hypothetical inputs in Table 2. The calculation is done on full RPM. In practise only a percentage is used to avoid burning out the motor. These values vary from 70 to 90 per cent of the maximum. [14]

Table 2: Data Input

Waterline length in feet: 30 feet

Beam at the waterline in feet: 10 feet

Hull draft in feet (excluding keel): 10 feet

Vessel weight in pounds: 2000 lbs

Engine Horsepower: 200 HP

Number of engines: 1

Total Engine Horsepower: 200 HP

Engine R.P.M. (max): 5000 RPM

Gear Ratio: 3:1

Shaft R.P.M. (max): 1667 RPM

Desired speed in Knots: 20 knots

Table 3: Horsepower Calculations

Total available horsepower at the engine(s): 200 HP

Total available torque ft/lbs at the engine(s): 210 ft/lbs

Horsepower loss of 3% per gearbox: - 6.0 HP

Horsepower loss of 1.5% per shaft bearing: - 6.0 HP

Total horsepower available at the propeller(s): 188.0 HP

Total torque ft/lbs available at the propeller(s): 592 ft/lbs

Table 4: Speed & Power Calculations

Basic displacement speed and horsepower required

Displacement hull speed (1.34 X sqrt of waterline length): 7.34 Knots

Minimum horsepower required at propeller(s) for Hull speed: 4.0 HP

Calculations based on desired speed and available HP

HP required at propeller(s) for desired 20 knots speed: 80 HP

Estimated speed with existing 200 horsepower:This is the speed we will use for the propeller size.

26.52 Knots

Table 5: Propeller Size

Number of blades Diameter (inches) Pitch (inches)

2 Blade 22.1 X 25.6

3 Blade 21.0 X 25.3

4 Blade 19.8 X 24.8

Shrenk’s Approximation

A theoretical method to predict the span wise load distribution for a wing is provided, given its characteristics. The assumption that is made is that the real distribution is somewhere between an ideal, elliptical, distribution, independent on wing shape and function that is dependent on the wing shape. The formula that is used is as follows: [16]

dAdx

=14ρV 2Clα

α [c ( x )+c ( 4π )√1−( 2xb )2]+ 14 ρV 2C lα∞

δ (x ) c (x)

The values in the formula are mostly constants and the speed is for cruise conditions. In particular the ideal lift curve slope is determined using the graph in Appendix D from [15]. The following relationship is important:

C lα=

Clα∞

1+C lα∞( F

π b2 ) (1+k )

The curvature is found using the graph in [16], dependant on the taper ratio.

Conclusion

There are key boat design features that must be incorporated into the design to make it efficient. This is flair and a step. A flair is the angled base of the hull. The shape is to ensure that the water is swept away from the main body. The higher the angle the faster the boat can go. For aircraft applications there is a limit for the angle for the loads. The step is a sharp discontinuity just aft of the COG. Landing on the step breaks the water and makes for a smooth, long touchdown.

There are numerous design issues in designing a submergible aircraft. The design parameters that must be strictly controlled are depth and duration of the dive. Once the aircraft has landed and stopped on the surface of the water it must reconfigure to prepare for its underwater phase. Design analysis will confirm the appropriate actions to take in this regard. These actions include:

Sealing all intakes and ports Flooding all non-essential areas Switch to auxiliary power Turn on oxygen supply for occupants Adjust lifting surfaces to minimise drag and produce down force

Once the aircraft surfaces, the reconfiguration repeats for flight mode. Emphasis is placed on the ability of the aircraft to submerge and not the range, payload or speed in any medium.

Product Requirement Specification

RequirementsTransport two people who have a height of 180 cm and a combined weight of 200 kg, without baggage.

Land and take off on water, as well as conventional runways.

The transportation system must utilise an on-board energy source/s that provide propulsion in air and water.

UnderwaterTotally submerge and maintain control.

Attain maximum speed of 5 knots

A loiter time of 20 minutes.

Airborne The range must be 800 nm

SurfaceAttain a minimum speed of 10 knots

The range must be a minimum of 50 nm

Constraints

UnderwaterDepth underwater of 10 m .

Aircraft may not refuel in the middle of the mission

The hull must be pressurised.

CriteriaLow fuel consumption as possible

Increase visibility underwater

Concepts

Three concepts were developed that all have the required power and performance characteristics to successfully adhere to the PRS. Each concept is explained followed by the drawing.

Concept 1

A low wing, single engine aircraft that is capable of taking advantage of ground effect on the surface of the water. The engine is positioned on fore of the vertical tail. Step landings are a preferred form for this concept.

Advantages Ground effect increases the range that the aircraft can cover on the water surface. The low wing will help in submerging. As the aircraft partially submerges, the flaps are

partially submerged and can be used to generate down force in the initial stages of submerging.

The single engine is positioned high above the water, with the fuselage partially protecting it from the spray.

The elevator will be sensitive with the engine blowing directly onto it.

Disadvantages Successful landing is dependent on pilot skill as there are additional variables to take

account of. Ground effect can make it difficult to land on conventional airstrips as the aircraft rides on a

‘cushion of air’. The performance of the single engine is critical importance. There is no alternative on the

occurrence of engine failure. There is a high power requirement for the main engine on water take off for low wing.

The three view of concept 1 is shown in Figure 4.

Figure 4: Concept 1

Concept 2

A high wing, two engine aircraft capable is high speeds on the surface of the water. The hull is capable of cruising on the surface like a power boat. The aircraft can handle step landing and hull and stern landings.

Advantages There is a large amount of power available for a high rate of climb and sail speed. The diversity of landing capabilities makes it easier for pilots The aircraft can sustain flight with one engine inoperative. A higher stall angle will be achieved with the position of the engine.

Disadvantage The wing mounted engines will contribute to a higher parasite drag than concept 1. Dual engines will have higher fuel consumption than a single, more powerful engine. There is additional weight associated with an extra engine. The wing loads will be higher and contribute to higher increase in skin thickness than the tail

in concept 1.

The three view of concept two is shown in Figure 5.

Figure 5: Concept 2

Concept 3

A low wing pusher with canards contributes to effectively one of the most difficult aircraft to land on and take off from water. With sufficient skill, then it is in the water it will be in its element. The design is effectively a modified marlin. It can take advantage of ground effect when close to the surface. The canards appear to be going through the fuselage. The pilot’s seat is fashioned directly on top of this. Step landing is recommended.

Advantages Highly streamlined design will operate will underwater and prove to have a lower true drag

coefficient than what is calculated. Ground effect increases the water surface tactical radius. Low wing helps in submerging. The width can allow for a longer fuselage section with a higher number of passengers.

Disadvantages A low angle of attach must be maintained during take-off to prevent propeller strike. This

contributes to a high SGA. Hull landings can severely damage the canards. The cockpit space will be constricted due to the ‘structural seat’. Novice pilots will struggle with operating the aircraft to its full potential.

The three view of concept 3 is shown in Figure 6.

Selection

With all the parameters considered, concept 1 is chosen for further development. Concept three is ruled out due to the skill required in safely landing and taking off. Concept 2 will prove to have a higher SFC and parasite drag. These factors will affect the range and cost of operation for the aircraft.

Figure 6: Concept 3

Design Development

WingThe wing was iterated with wing loading and aspect ratio to find the configuration with the lowest drag. Each iteration was run through the following process to find the drag.

Re f=ρσv l¿

μRec=313713( l

¿

k )1.067

C f=¿GraphQwing=1+Ltc+100( tc )

4

Swet1 st=2Sref−(CR×FuselageWidt hassumed )Cdo1 st=3 (C f ×Qwing×Swet )+15%

Sref

CD=Cdo+

C l2

πARe

D= 12 g

ρoσ v2Sref CD

The optimal wing was found with the following characteristics in Table 6:

Table 6: Wing Characteristics

Wing Loading 222S_ref 20.27027027Cl_cruise 0.353199194V_stall 47.14090905AR (assumed) 8b 12.7342908Taper (ass) 0.7Cr 1.872689823Ct 1.310882876MAC 1.608310084Sweep 0.34906585t/c 0.12NACA 4412It must be noted that the sweep is listed for quarter chord and is in radians. The aerofoils is chosen because it gives good characteristics for light, slow aircraft and allows space for structure.

Empennage SizingFor the horizontal and vertical tails:

SVT=V VT Sref b

lVTSHT=

V HT Sref MAC

lHT

The volume coefficient for the horizontal tail (VHT) and vertical tail (VVT) are 0.8 and 0.08 respectively. The denominator in both equations is the distance from 0.3 MAC to 0.3 MAC of the respective tail section. A similar process was the wing was followed to size the empennage. Listed in

Table 7: Empennage Characteristics

The aerofoil selected for the vertical tail is so thick to allow space for the structure needed to hold the engine. Both aerofoils are symmetrical, as per theory for empennages.

Fuselage SizingConsidering the aircraft will be entering the water, it must me as long and sleek as possible. For this reason, the seating position will be tandem and not side by side. Using standard seating data, the width is 0.48m and the pitch is 0.81 m.

Using tandem seating arrangement, the width of the fuselage is:

W fuselage=W seat×1.2W fuselage=0.58m

This arrangement will limit the number of people travelling in the aircraft to its already conservative 2. To be able to have a contingency plan to provide for more passengers, a ‘tri’ seating plan can be devised shown in . There will be an aisle width of 0.3 m.

H Tail V Tail

V ht 0.8 0.08

S_ref_ht

2.60807 2.06502

AR ass 3.3 3

Cr 2.73539 1.185232

Ct 0.820617 0.474093

b 1.466853 2.488988

MAC 1.949842 0.880458

sweep 0.487461 0.220115

t/c 0.12 0.15

NACA 0012 0015

S_wet 5.216141 4.13004

Figure 7: Top view, Tri Seating

W fuselage=(2∗W seat+W aisle )×1.2W fuselage=1.32m

This fuselage ensures that there is a thin leading edge, capacity for improvement and good visibility for the pilot. To compensate for the extra width, the fuselage bust be lengthened, to maintain a healthy length to diameter ratio in the water. If the length of the aircraft is capped at 10m, the length to diameter ration is within acceptable range.

DragThe drag for the wing, empennage and horizontal tail was done similarly. The fuselage and the engine were calculated with a different form of Q value. To estimate the Cdo the, D/q for each component is found, summed and divided by the wing area.

Dq

=C fQ Swet

Cdo=

ΣDq

Sℜ f wing

Iterating twice with wing area, the parasite drag was found to be:

Cdo=0.0196

The Reynolds cut-off number is lower that the Reynolds flight number in every case. This is independent of density. Therefor the parasite drag in air would be the same as for water and plasticine. There will be a significant drag rise in other media due to the effect of density. The drag table is shown in Appendix A, Table 17.

EngineThe engine will be operating in flight and on the surface of the water. The design point of the engine is at cruise. The cruise speed will be 1.15Vd. Vd is the speed at which the drag is lowest. Figure 8 shows the drag of the aircraft with respect to speed.

0 50 100 150 200 250 3000.0

100.0

200.0

300.0

400.0

500.0

600.0

700.0

800.0

900.0

1000.0

Series2

Figure 8: Drag Bucket

The lowest drag was found at 110m/s. Therefor the cruise speed will be 126.5 m/s, approximately M0.4. At cruise, with an initial assumed efficiency (np) of 0.7, the required HP was found using the following procedure:

T=D

H Pcruise=Tvnp75

H Pground=H Pcruise

σcruise

To find the diameter of the propeller the following equations were iterated to find the value where the assumed efficiency equals the efficiency from the standard propeller efficiency chart [25]. The data of the diameter sizing is shown in Appendix A, Table 18.

C p=SHP (75)gρoσ n

3d5

J= vnd

np=¿Graph

The diameter of the propeller was found to be 2 m and the engine power was found to be 1111 HP at sea level. The thrust curve is as follows for sea level. The raw data is available in Appendix A, .

10 30 50 70 90 110 130 150 170 1900

200

400

600

800

1000

1200

1400

1600

Sea Level

FL200

Figure 9: Thrust Curve

The selected engine that meets the requirements is the Garrett TPE331-14GR. The specifications are shown in Table 8Error: Reference source not found.

Table 8: Engine Specification

Power SHP 1120Length (m) 1.33Diameter (m) 0.53SFC (kg/HP/hr) 0.228Weight (kg) 195.1

Overlaying the thrust and drag curves, the maximum speed can be extrapolated by inspection (or curve fitting.

0 20 40 60 80 100 120 140 160 180 2000.0

200.0

400.0

600.0

800.0

1000.0

1200.0

1400.0

1600.0 Drag

Thrust FL200

Thrust SL

Drag SL

Figure 10: Thrust and Drag at FL200

From Figure 10, the maximum speed it found to be 160 at FL200 and 150 at sea level. The raw data is in Appendix A, Table 20

Wing Load DistributionUsing Shrenk’s approximation a load distribution was found for the wing shown in Figure 11. Numerically integrating this produced the shear force diagram shown in Figure 12. The second numerical integration would produce the bending moment diagram shown in Figure 13

0 100 200 300 400 500 600 7000

500

1000

1500

2000

2500

3000

3500

L(y) e

L(y) w

L(y)

y(mm.)

L(y)

(kg/

m)

Figure 11: Shrenk Load Distribution

0 100 200 300 400 500 600 700

-4000

-3500

-3000

-2500

-2000

-1500

-1000

-500

0

y(mm.)

Sh

ear(

kg)

Figure 12: Shrenk Shear Diagram

0 100 200 300 400 500 600 7000

2000

4000

6000

8000

10000

12000

y(mm)

Ben

din

g M

om

ent

(kg

.m)

Figure 13: Shrenk Bending Moment

In the preceding diagrams the highest load was found to be at the root. These loads correspond to the distributed, shear bending moments of 3142.383 kg/m, -3372.11kg and 9578.125 kg/m respectively. This incorporates a safety factor of 1.5.

Load Factors

The following analysis is taken from [17] and is for take-off and landing on water. The TOW from water is equal to the initial MTOW with fuel reduced from the airstrip to landing.

The assumptions are:

There is no fuel used during the submerged phase of the mission. The aerodynamic lift for the wing is equal to zero. No slip stream effect Landing is symmetrical-loads applied at the keel to the centre of gravity. Loads are directed perpendicular to the keel Upward and side component: Fr (0.75)(tan) and Fr (0.25)(tan) respectively. Where Fr is the

resultant load.

The limit load factor for water reaction loads is associated with the type of landing case. The two cases are step landing and bow landing. The step is the sharp discontinuity in the hull. This allows the aircraft to clearly break free from the water during take-off.

The most common form of landing on water is the step landing. This is where the aircraft touches down at the position of the step in the hull. This breaks the water suction and leads to a smooth landing. A bow landing is common for larger lea planes. This is where the bow of the aircraft touches down first. This results in a rougher but shorter landing.

For the step landing case:

nw=C1V so

2

( tan β )23( W4.448 )

13

For the bow and stern landing case:

nw=C1V so

2

( tan β )23( W4.448 )

13

×K1

(1+rx2 )23

Take-off condition for the hull and main float

The assumption is:

The stall speed for take-off is the same as the stall speed for landing

A downward inertial load is linked with a load factor as follows:

ni=C ¿V s1

2

( tan β )23 ( W4.448 )

13

Hull and float pressures


The loads are symmetrical. The hull is flared . The area over which the pressure is applied do not need to extend over an area that would

induce critical stress in the frames or in the overall structure. The uniform symmetrical pressure is applied over the entire hull. Transferred to the sidewall

structure, it is not transmitted as shear or bending loads in the fore and aft directions.

The bottom pressures that the hull and floats will experience can be calculated as follows. The design of the structural elements will use these loads. There are numerous advantages to flared hulls, discussed in the literature review. Flared bottom hulls will be subjected to the following pressure at the keel and the chine:

Pk=C2K2V s1

2

tan β×6.895Pch=

C3K 2V s12

tan β×6.895A uniform symmetrical pressure distribution is as

follows:

Ps=C3K 2V s0

2

tan β×6.895

Auxiliary Float Loads


The loads are distributed over the float bottom. For step loading, the water load must be applied to a point 3/4th of the distance from bow to

step. For bow loading, the water load must be applied to a point 1/4th of the distance from bow to

step. The loads must be perpendicular to the keel. L should not exceed three times the weight of the displaced water when the float is

completely submerged. should not be less than 15°

The resulting water loads from auxiliary floats are:

L=4.448×C5V so

2( W4.448 )

23

( tan β )23 (1+r y

2 )23

Immersed float conditions


The resultant load must act through the centroid of the float cross section. The load must act at 1/3rd the length from bow to step.

The components of the load are:

vertical=ρgV aft=C x ρV f

23 (K V so )2

2side=

C y ρV f

23 (K V so )2

2

The tabulated load factors are shown in

Table 9: Water Load Factors

L Limit Load (N) for float 1872.038Aft Float component 3052.003side Float component 2432.219nw Water load -(Nw/W) step landing 0.547

nw Water load -(Nw/W) bow and stern landing 0.419

nw Water load -(Nw/W) twin float landing 0.442

ni Inertial load factor for take off 0.013

Pk Pressure at the keel (KPa) 22.256

Pch Pressure at the chine (Kpa) 16.718

Ps Symmetrical pressure distribution 9.780

The breakdown of the empirical factors for the components is shown in Appendix B, Table 24.

Structures

Wing Structure

There are different forms of wing structure. As the aircraft will be in a highly pressurised, buckling of the wing is of prime concern. The forms of the wing structures are as follows: [lian-3e.pdf]

Single spar Double spar with stiffeners Triple spar

The single spar cannot be considered as viable option as the torsional stiffness of the wing will be compromised. The double spar with the stiffeners will be the best in theory as it offers a compromise between stiffness and weight. The triple spar system will be too heavy for an efficient design, despite it providing the best stiffness figures.

Spar Position

Typically the front spar is position between 10% and 25% MAC and the rear is positioned between 60% and 70% MAC. An increase in torsional stiffness can be achieved by moving the rear spar further aft, creating a larger area in the middle. As the spar moves further back, it must decrease in height, affecting the support of the flaps. The final position with these limits in mind, is 20% and 65% MAC illustrated in

.

Rib PositionProviding stiffness in the chord wise direction, the ribs react to the shear force caused lift. They aid in preventing bucking, but don’t provide additional stiffness in the span wise direction. The rib pitch can be approximated by the following formula:

L=0.55×√D

Where l is the pitch of the ribs, D is the maximum height of the aerofoil and the factor of 0.55 has been extrapolated from test data. With the maximum thickness at the root being 0.3 m, the pitch of the ribs along the span of the wing will be:

L=0.55×√0.3L=0.301m

With the half span of the wing being 6 m excluding the fuselage, there will be 20 ribs on each wing.

Shear Flow AnalysisThe maximum force on the wind will be at the root. The load distribution of the wing can be found using Shrenk’s approximation.

The method that is to be carried out for the wing structure for a wing converted to the model shown in

Figure 14: Wing simplification

Assume that the torsional rigidity at the wing tip is 1X108 Nmm/rad.

T s=TLθc=

Mmax

T s

This value of c will be the required specific torsional angle for the wing. With shear flows and torque formulae in each cell, the thickness of the skin can be found.

T=∑ 2Qi A i2cAG=∮ Qdst

The integral around a closed loop represents the loop of each cell in the wing. Simultaneously solving for the four equations, for the required value of c, the thickness is found for the skin. The thickness of the spars is an assumed value.

Similar analysis is done for titanium instead of aluminium. All values are kept the same except G. GTi is 48 GPa. The thickness was found to be less, however the mass of the structure will be greater due to the higher density of titanium. The weight penalty is not worth the reduction in thickness.

The shear flows in the MAC is to be found with the applied loads. The method is as follows:

Find statically determinant shear flows of open cells Add a shear flow (Q) to each cell for compatibility when the cells are closed Find the shear centre Find the torque due to the applied load Add a shear flow (Q’) to each cell to compensate for pure torque.

Cell 3Cell 2Cell 1

Assumptions

The simplification in Figure 14 holds true for the following analysis. The stiffeners all have a cross sectional area (CSA) of 100 mm2. The Ix value of the structure takes account of stiffeners that are equidistant from the

horizontal neutral axis.

The equations to be used are as follows:

For statistically determinant shear flows qk in each cell with an applied load Sz, a CSA ci, a vertical distance z, the following equation applies for the structure in Figure 15

qk=−S z

I x∑i=1

k

c i zi

Figure 15: Statistically determinant shear flows

The double strike in Figure 15 represents cut cells, with the boxes stiffeners and the arrows shear flows. After the shear flows are found, the sum of the forces must be as follows.

∑ F x=0∑ F y=Sz

Where

F=qds

Once the cells are closed, an additional Q must be added to each cell as in Figure 16.

cA=∮ Qds2>¿=0

¿

Figure 16: Compatibility Flows

The shear centre, x, will be needed to find the applied torque, T, of the force, F. It is found using the sum of moments about a convenient point. The bottom left of the structure is chosen.

Sz x=qi x ids

T=Sz(x−xapplied)

Cell 3Cell 2Cell 1

Q 3Q2Q 1

With the applied torque, an additional flow Q’ is added to each cell. Solving simultaneously, a new value of c is found.

cA=∮ Q'ds2>¿

¿T=∑ 2Qi A i

Finally a check must be done for the force balance.

∑ F x=0∑ F y=Sz

Due to an unforeseen error with the structures spreadsheet, the details of the wing structure cannot be presented. This is partly due to poor time management. The average skin thickness is idealized at 4 mm. This can be rationalized by:

The skin loads are purely shear. There are low maneuvering capabilities and requirements. The speed in water is restricted.

Tail StructureThe method of analysing the tail structure is the same as the method outlined for the wing structure. The variables that change are the geometry of the wing, the position of the spars and the load applied.

Spar PositionThe vertical tail is different to the wing in that a higher percentage of the area and span is dedicated to its flap, the rudder. The position of the tail must be such that it provides support to the rudder and stiffness of the overall structure. The final position with these limits in mind is 25% and 60% MAC.

Rib PositionA similar analysis to the wing is done for the rib positions for the tail. Both the vertical and horizontal tail has a thickness to chord of 0.1 and a span (half span for the horizontal tail) of 1.4m. The rib pitch

L=0.55×√DL=0.55×√.14L=0.206m

This amounts to 7 ribs for the vertical tail and 14 for the horizontal tail.

Shear Flow AnalysisThe maximum bending moment that the tail experiences is restricted to 3213 N.m. This is achieved with low speed and deflection in water. Using the formulae and analysis detailed in wing structure, the required thickness of the skin is found for three materials. The detailed data is provided in tables in Appendix C. The summary is detailed in Table 10 and Figure 17.

Table 10: Empennage Skin Thickness

Aluminium Titanium SteelSkin t (mm) 2.16 2 3 1 1.37 0.8Q1 462.5009 453.6935 467.7741 434.7082 447.6074 433.8134Q2 615.1329 619.9106 612.1894 629.7277 623.1231 630.1766Q3 405.4786 398.2258 410.0344 383.8307 393.443 383.1876c 0.00018 0.000194 0.00013 0.000386 0.00018 0.00018

1 1.5 2 2.5 3 3.50

0.00005

0.0001

0.00015

0.0002

0.00025

0.0003

0.00035

0.0004

0.00045 Required c

Linear (Required c)

C vs t

Polynomial (C vs t)

Figure 17: Skin Thickness Iteration for Aluminium

Table 10 suggests that using materials like steel and titanium can help in reducing the overall skin thickness. This comes with a weight penalty as the densities are higher. A check for the effectiveness is to find the mass of an elemental section of skin.

Mass= ρt

MassAl=(2700 ) (2.16×10−3 )MassAl=5.832kg

MassSt=(7850 ) (0.8×10−3 )MassSt=6.28kg

MassTi=(4506 ) (1.37×10−3 )MassTi=6.17 kg

The above calculations show that for this application, aluminium is the best material with the lowers elemental mass. All the skin will therefore be made from aluminium.

Hull Structure

The hull structure needs to withstand the external pressure of water at maximum depth. This is the design point for the fuselage. As the vessel will have to be pressurised, the pressure that the hull sees will be the differential pressure. . With a maximum depth of 10 meters is needed. A maximum operating pressure of 150 KPa is used with and incorporated safety factor. The differential pressure then will be 60 KPa with the hull pressurised to a comfortable level.

To effectively submerge, part of the main hull will be flooded. This will reduce the pressurised section and the differential pressure. Using thin plate analysis with an effective thickness will provide a conservative base of the stresses that the skin, longerons and ribs will have to take. In an elemental section of skin, there will be hoop stresses and longitudinal stresses. As the section of skin is thin, it is assumed that there are negligible radial stresses [28].

Figure 18: Idealised Stress in Fuselage

The respective formulas for longitudinal and hoop stresses are:

σ 2=Pr2t

σ 1=Prt

The above formulas are for r is radius, t is effective skin thickness and P is differential pressure . The length of the pressurised section will be approximately half the fuselage length. This is to account for the possibility of technical failures with the motor in the submerged stage. The compartment with the secondary power source can be accessed.

LongeronsThe assumptions that are made are that the longerons take axial load only and the skin takes shear stresses only. The longitudinal stress will be taken purely by the longerons. From Figure 18, a different form of the longitudinal stress can be formulated:

pπ r2=σ 1 A

As all the load is taken be the longerons, the area is simple the product the cross sectional area (CSA) and number of longerons. The longeron cross sectional area will need to be 200 mm2. The yield stress of 414 MPa of aluminium will have a safety factor of 1.5. The number of longerons needed will be:

Hoop StressLongitudinal Stress

pπ r2=σ 1nCSA

n= pπ r2

σ1CSAn=19.2

This is a very high number; an alternative material must be investigated. Using titanium, the yield strength is 924 MPa. The number of longerons will then be:

n=8.6

This is by far the better solution. The weight penalty of having a higher density material is overshadowed by the strength. Therefore there will be 9 titanium longerons.

RibsThe ribs of the structure will have to carry hoop stresses. With the length of the fuselage 10 m, there will be two sections, pressurised and non-pressurised. The non-pressurised section will have an internal pressure of 60KPa, and a differential pressure of 90 KPa. Using the same logic as the longeron design, an alternative form of the hoop stress formula is presented below:

pA=σ2 A ribs

p2 rdx=σ22nCSAdx

Separating the two sections, the number of ribs in each section can be determined. Titanium will be used with a CSA of 200 m2. The non-pressurised section will be divided into two, fore and aft of the cockpit. The longest unpressurised section will therefore be 3 m near the tail where the average radius is approximately half of fuselage.

nCSA= prσ2

nCSA=0.548

This is the effective thickness for an elemental section of the fuselage. For the entire 3 meters (L), the number of ribs required is:

n=0.548× LCSA

n=8 ribs

This gives a rib pitch of 0.375 m. Similarly the number of ribs for the 5 m pressurised section will be:

n=18 ribs

This gives a rib pitch of 0.278 m. This pressurised section has a shorter pitch despite the lower differential pressure. The radius has a significant effect.

Seals

To ensure that the aircraft is air worthy after submerging, proper sealing is needed for every component that is exposed to the air in flight. These components include:

Air intakes and exhausts for the engine. Static pilot probes All doors including landing gear and hull

The option of exiting the aircraft and adjusting the parameters is not always a viable option for a pilot and crew. This is especially true in a stealth mission. There can be dangerous conditions at sea and the stability of the aircraft on the water would be compromised if there is significant movement of its passengers. To have the seals working remotely, automated locking devices must be used.

MechanicalA device similar to a ball cock can be used to close a valve. As the water level rises, the inflated ball rises and through a connecting rod closes a valve. A modified version of a device shown in Figure 19 will seal the required ports [30].

Figure 19: Ball Cock

This method will affect the performance if the aircraft and the ports it is sealing. If incorporated, in flight the losses due to the resistance of the device will give increased uncertainties. An increase in Cdo is inevitable with mechanically operated seals.

ElectricAn electrical system will help in automation without elevated performance effects. The seals must be engages when the craft hits the water to prevent water damage or humidity effects to equipment. Hydraulic, pneumatic or electrically driven solenoids can act as a power source to seal the required vent. The trigger will be the water itself. A strategically positioned cell/switch can be closed on direct contact with water. This switch can intern power the drive system to seal the necessary ports. It is of utmost importance for ports to be sealed properly maintain proper performance of the respective components.

In the case of static ports, the tube can be ‘flushed’ out with air to get rid of residual water drops on resurfacing.

Propeller

A second propeller is needed for the water section of the mission. This has a power source, decoupled from the main engine. The following is the design specification for the secondary propeller.

Due to the nature of the task, the propeller calculation for the water was done using a combination of imperial data from spread sheets from propeller manufacturers for boats. Two sources were used to find both, the correlation and the worst case scenario. The basic inputs are presented in Table 2. The assumption that was made for the hull draft (height of the waterline) is that the water line exceeds the height of the aircraft for underwater travel.

Table 11: Propeller Data Input

Waterline length in feet: 30 feet

Beam at the waterline in feet: 6 feet

Hull draft in feet (excluding keel): 10 feet

Vessel weight in pounds: 11000 lbs

Engine Horsepower: 24 HP

Number of engines: 1

Total Engine Horsepower: 24 HP

Engine R.P.M. (max): 1000 RPM

Gear Ratio: 1.5:1

Shaft R.P.M. (max): 667 RPM

Number of shaft bearings (per shaft): 1

Desired speed in Knots: 5 knots

Table 12: Propeller output [14]

Total horsepower available at the propeller(s): 22.9 HP

Total torque ft/lbs available at the propeller(s): 181 ft/lbs

HP required at propeller(s) for desired 5 knots speed: 14 HP

Estimated speed with existing 24 horsepower:This is the speed we will use for the propeller size.

5.89 Knots

Propeller SizeNumber of blades Diameter (inches) Pitch (inches)

2 Blade 25.1 X 20.4

3 Blade 23.9 X 20.2

4 Blade 22.5 X 19.8

Table 13: Propeller Output [19]

Shaft Horsepower required at propeller 18

Pounds per shaft horsepower required. 1,213

Approximate maximum speed attainable 5.3

Shaft HP required for speed required at top of page. 21

Required prop pitch for top speed. (inches) 20

3 blade diameter (inches) 25

4 blade diameter (inches) 24

It is important to note that all of the calculations above are based on full RPM and HP. Most engines are rated to run at a percentage of their full RPM. This is what will determine your maximum cruising speed. The propeller sizing calculations below are based on 90% of full RPM, which allows the engine to develop its maximum power without overloading.

There is close correlation between the two methods. The four blade propeller will be selected to have the minimum diameter. This is critical as during flight the propeller needs to be stowed, to reduce drag. The required motor will have to be at least 18 HP at the propeller. With losses included, a 24 HP motor would be perfect for the required speed of 5 knots.

Secondary Power Source

The decision was made early in the design to separate the propulsion drives for over and underwater. The decision to be made is which form of power source would work to power the secondary propeller for the underwater part of the mission. The power requirements are constant at the equivalent minimum of 24 HP.

The prospect of any form of internal combustion engine is to be ignored due to complications. These include extra oxygen requirements and the safe exhaust of waste gasses. The power source must be completely independent of harmful exhausts. In the pressurised vessel this could be lethal, if not handled properly.

As nuclear power plants in an aircraft have limited information, it will not be considered as a realistic option for now. As explained in the literature review it has been done before by Convair, but these aircraft were much bigger than the concept at being considered.

The propeller engine can be sealed and supplied with oxygen to work under water. There are some issues with this concept. The efficiency of the engine goes to the dogs when it is run cold. In the air if it is cold the efficiency jumps due to the higher density. However the heat loss could adversely affect the performance. The amount of oxygen needed for the current engine is 5.5 kg of air per second. It has a air to fuel ratio of 50:1. The extra oxygen for 20 minutes will amount to 6545 kg. [31]

Electric Engine

With electric engines the torque curve is flat with respect to RPM. This is unlike diesel engines that have decreasing functions of torque and horsepower. Undersized props are used in diesel engines, to prevent stalling the engine at low RPM. There is no minimum idle speed for electric motors. Even with large three bladed propellers, electric motors and apply full torque at zero RPM.

The propeller can be used to charge the batteries. If the speed of the vessel increases the prop is forced to turn faster due to the rushing water past it. This may be due to change in currents or can be forced. By using the thrust of the main engine on the surface of the water, the battery can be recharged by allowing it to 'freewheel'. This can recharge the batteries and increase the loiter time in the water.

An engine that fit the description was found in [20] to have the following specifications in

Table 14: Electric Engine specifications

Power 1.8 KW ≈ 24 HPLength 0.432 mWidth 0.165Weight 76.7 kgMax RPM 1000Input 144 V

An electric motor needs a battery to run from when the internal power is run out. A combination of a rechargeable battery and an inverter is needed to form a UPS (uninterrupted power source). The inverter is needed to convert from AC to DC or vice versa and to keep the power supply constant. The electric engine will run off DC power while components, like lights, in the aircraft would operate more efficiently with AC power, hence the use of the inverter.

There is a huge demand for power in the 144 V motor. Multiple batteries will be used to achieve the voltage requirements. The operation time is not critical for the selection of the battery. The battery will have a larger capacity that to just power the motor as auxiliary components will draw from the battery. However, the higher the Watt-hour (W-h) rating of the battery pack, the heavier and larger the battery will have to be. The following equations are to calculate the battery requirements [24].

HP×746=Watts24×746=1.8kWkW=Voltage× Amp−hrUsable kW=kW (0.8 ) (0.55 )

kW needed=1.8 kW

(0.8 ) (0.55 )kW needed=4.1kW

The empirical formula for usable kW takes into account the power loss due to acid leakage (0.55) and other effects. The value of 4.1 kW is the power needed to run the motor alone for an hour. The power requirements of the aircraft systems will be order of magnitudes less than the draw from the motor. As a conservative estimate, 4.1 kW will be the battery requirements.

Using the information in Table 14 , [23]and [22] the respective battery and inverter specification are as follows in Table 15. It should be noted that to achieve the 144 V input requirements, 11 13.5V batteries have been combined. Table 15 shows the specifications of the Nortek battery pack and the Dimensions 3500D. The arrangement of the batteries has been chosen to be 3X2X2.

Table 15: Battery and inverter specifications

Unit Inverter Battery (one) Battery (11 pack) Battery (one) Battery (12 pack)Length (m) 0.40 0.120 0.48 0.196 0.588Width (m) 0.43 0.190 0.38 0.13 0.26Height (m) 0.254 0.130 0.26 0.175 0.35Weight (kg) 29.1 4 44 10.9 130.8Output 3500 W 540 W-h 5940 W-h 396 W-h 4752 W-h

The 11 pack battery [21] is found to be much more efficient and more powerful than the 12 pack, suggested by the manufacturers of the motor [23]. There will be excess power with less space and weight of the components. It seems like the manufacturers get commission by promoting the inferior batteries.

Fitting

The fitting of all the components is illustrated in Figure 20. The addition of water tanks is to aid in submerging the aircraft. The rotor in the tail is connected via a universal joint and is deployed as the aircraft comes to a stop and is ready to submerge.

Figure 20: Fitting

Floor line

Water Tank

Seat with Oxygen tank

Electric engine

Battery

Inverter

Sealed Wall

Performance

Field PerformanceThe field performance criteria states that existing runways must be used. The conventional take off conditions will follow the following procedure.

Sgr=mv lo

2

g ( (T o−Do )+ (T lo−Dlo ))

Do=μm

Dlo=12g

ρσV 2SCdo

v lo=1.1 vstall

Where To and Tlo are values found from the thrust curve in Figure 9 at sea level.

Sga=2W (h+ vobs−v lo

2

2g )T obs+T lo−D obs−Dlo

vobs=1.2vstall

∆Cl=dCldδ

×δ×b f

bw

∆Cd=1.1 (sin (δ ) )2×S f

SwCl lo

' =Cllo−∆ClCd lo=Cdo+∆Cd+Cl lo

' 2

πARe

Table 16: Field Performance

Symbol Valueδ (°) 20dCldδ

4.25

b f

bw

0.4

S f

Sw

0.18

h (m) 10.67

Using the values in Table 16, the field performance is found to be:

SGR=465.78mSGA=396.28m

This is in line with existing runway requirements. The expanded calculation is tabulated in Appendix A, Table 22.

Underwater Speed

The electric engine produces a small amount of torque compared to the drag forces that the aircraft can experience underwater. The original specification was for the craft to do 10 knots underwater. This would prove to be very tricky as the drag appears to rise exponentially with speed. Figure 21illustrates this relationship as the thrust remains at a relatively constant rate.

0 5 10 15 20 250.0

1000.0

2000.0

3000.0

4000.0

5000.0

6000.0

7000.0

8000.0

9000.0

10000.0

Water dragengine limit

m/s

Drag

(kg)

Figure 21: Drag in Water

Velocity Minimum Unstick

The minimum unstick velocity (VMU) will be the same in air regardless on the terrain. This is the velocity at which the aircraft is free to rotate. This corresponds to the moment balance about the real landing gear (RLG) equalling zero. The free body diagram is shown in Figure 22. [25]

Figure 22: VMU

Using idealisations, L1 can be found:

L1+L2≈Lfuselage

2L1≈0.06 L2

With Lfuselage equal to 10 m, combining the two equations gives

L1=0.3mL2=4.7m

The governing equation with lift of the wing excluded, is as follows,

∑M RLG=0M L1=LT L2

The lift on the horizontal tail can be approximated relating it to the wing. LT=12 g

ρσ SHTV mu2ClT

ClT=ClαHT δClαHT=ClαW×0.4×δ

With ClW as 5.8 /rad, δ as 15°, Sht as 2.6 m2 and conditions at sea level, the Vmu is found to be:

V mu=√ 2 gM L1×180L2ρσ SHTClαW×0.4×δ×π

V mu=49.2m /s

L1 L2

m

LT

Water Take Off

There is no criterion specifying the length of the water take off. However there is a maximum drag that must be overcome on the water during take-off. The drag from the water is shown in Figure 23[27].

Figure 23: Combined Resistance with Thrust and Trim Angle

It can be seen that at approximately 0.5 of the unstick speed (VUS), there is a peak in drag. The unstick speed on water is the same as on land. This figure provides drag characteristics during take-off on water. Combining the data in Figure 23 with the aircraft specifications, the drag due to water is shown in Figure 24. The data is tabulated in Appendix A, Table 23.

0 20 40 60 80 100 120 140 160 180 2000

200

400

600

800

1000

1200

1400

1600Drag on take off

Series4

Thrust SL

Figure 24: Water Drag

It is to be noted that at the peak of drag on the water, there is not enough power in the engine to power past it. As the data is empirical, the hull must be further analysed to reduce the peak drag. This can be done by increasing the flair angle, or introducing multiple steps in the hull. Selecting a new engine on the basis of hull design will have a negative effect on the fuel flow of the mission.

Drag Polar

The cruise drag polar is shown in , for the clean configuration and flaps extended. The trend of the drag polar will be the same for conditions at sea level and in water.

0.000 0.050 0.100 0.150 0.200 0.250 0.300 0.350 0.4000.000

0.500

1.000

1.500

2.000

2.500

3.000

3.500

CruiseTop SpeedFlaps-20 deg

Cd

Cl

Figure 25: Cruise Drag Polar

The under water conditions show an exponential decay of lift coefficient with speed. This is due to the density of the water being an order of magnitude higher than air. This contributes to the drag coefficient being primarily reliant on parasite drag and not induced drag. The highest lift coefficient in water is as the craft starts moving, with a lift coefficient of 0.48 at 3 knots. The discussed is presented in Appendix A, Table 20.

The top speed is at 160 knots at altitude, and it shows the cut-off point of the graph. There are two points for the top speed; the higher one is for the flaps extended.

Mission Profile

The mission profile was explained in the PRS. The range is 800nm in the air, 50 on water and 20 minutes submerged at a depth of 10 m. The cruise altitude is chosen to be 15000 feet. The mission profile is illustrated in Figure 26.

Figure 26: Mission Profile

In Figure 26, the numbered items correspond to the stage of the mission.

1. Take off from an airfield, and climb to FL150 for cruise. The maximum half cruise is 400 nm, at a speed of 126 m/s.

2. Land on the surface of a water body, significantly deeper than 10 m. The main engines can be run, or the auxiliary power can be used for this stage, depending on required speed or ambiance. Before, entering stage three, all ports, vents and hatches must be sealed. Flood tanks must be filled and the main engine must be disengaged.

3. The secondary power source is used for this stage. The loiter time depends on the oxygen supply and battery charge, 20 minutes is recommended.

4. Resurfacing, the flood tanks are pumped dry, and seals are opened. The take off happens under the power of the main engine, as the secondary propeller is stowed.

5. Climb to FL150 for return journey.

12

34

5

References

[1] https://www.fbo.gov/?

s=opportunity&mode=form&id=825bc0a788e33d95e9f4dafa467e0e5a&tab=core&_cview=0

[2] http://www.freepatentsonline.com/2444332.pdf

[3] http://www.aerospaceweb.org/question/aerodynamics/q0130.shtml

[4] http://cr4.globalspec.com/thread/39112/Why-is-DOD-not-considering-Unmanned-Surface-

Effect-Vehicles

[5] Bradley, R. E., Canvair Advanced Designs, Specialty Press 2010

[6] http://www.aviastar.org/pictures/usa/cons_coronado.gif

[7] http://science.howstuffworks.com/transport/engines-equipment/question286.htm

[8] http://engineering.sdsu.edu/~hev/motor.html

[9] http://www.mounteverest.net/expguide/oztech.htm

[10]http://www.airproducts.co.uk/homecare/health_authorities/homeOxygenService/

OurService_Equipment-cylinders.htm

[11]http://auto.howstuffworks.com/fuel-efficiency/alternative-fuels/fuel-cell.htm

[12]http://science.howstuffworks.com/nuclear-power1.htm

[13]http://www.aero-expo.com/aero-en/press/press-releases.php?obj_id=128&sMode=detail

[14]http://www.vicprop.com/calculator.htm

[15]IRA H. Abbott, “Theory of Wing Sections”, page 488[16]O. Shrenk, NACA Technical Memorandum No. 948, Luftwissen, Vol. 7, No. 4, April 1940.[17]ASTM International, F2245-10, Standard specification for Design and Performance of Light

Sport Airplane http://www.astm.org/DATABASE.CART/HISTORICAL/F2245-10.htm[18]Flared vs unflared http://www.maasboats.com/design.htm[19]www.alberg30.org/maintenance/MechanicalPropulsion/Propeller/ [20]http://www.electricmarinepropulsion.org/Pages/Technology_hptorque.html [21]http://www.nortekusa.com/en/products/batteries-and-battery-cans [22]http://www.electricmarinepropulsion.org/Pages/Components_Inverter.html [23]http://www.electricmarinepropulsion.org/Graphics/Batteryspecs/batteryspecs_AGM.pdf [24]http://www.ev-propulsion.com/EV-calculations.html [25] MECN 3005 Aircraft Design, course notes, University of Witwatersrand[26]http://www.engr.colostate.edu/~dga/mech325/handouts/pressure_vessels.pdf [27]Darrol Stinson, The Anatomy of the Airplane, second edition[28]http://www.engr.colostate.edu/~dga/mech325/handouts/pressure_vessels.pdf [29] R.C. Hibbler, Mechanics of Materials, seventh edition, page 436[30] http://www.thefreedictionary.com/ballcock[31]http://www.aerocompinc.com/TPE331%20specifications.htm

http://www.thefreedictionary.com/ballcock

http://www.engr.colostate.edu/~dga/mech325/handouts/pressure_vessels.pdf

http://www.engr.colostate.edu/~dga/mech325/handouts/pressure_vessels.pdf

http://www.ev-propulsion.com/EV-calculations.html

http://www.electricmarinepropulsion.org/Graphics/Batteryspecs/batteryspecs_AGM.pdf

http://www.electricmarinepropulsion.org/Pages/Components_Inverter.html

http://www.nortekusa.com/en/products/batteries-and-battery-cans

http://www.electricmarinepropulsion.org/Pages/Technology_hptorque.html

http://www.alberg30.org/maintenance/MechanicalPropulsion/Propeller/

http://www.maasboats.com/design.htm

http://www.astm.org/DATABASE.CART/HISTORICAL/F2245-10.htm

http://www.vicprop.com/calculator.htm

http://www.aero-expo.com/aero-en/press/press-releases.php?obj_id=128&sMode=detail

http://science.howstuffworks.com/nuclear-power1.htm

http://auto.howstuffworks.com/fuel-efficiency/alternative-fuels/fuel-cell.htm

http://www.airproducts.co.uk/homecare/health_authorities/homeOxygenService/OurService_Equipment-cylinders.htm

http://www.airproducts.co.uk/homecare/health_authorities/homeOxygenService/OurService_Equipment-cylinders.htm

http://www.mounteverest.net/expguide/oztech.htm

http://engineering.sdsu.edu/~hev/motor.html

http://science.howstuffworks.com/transport/engines-equipment/question286.htm

http://www.aviastar.org/pictures/usa/cons_coronado.gif

http://cr4.globalspec.com/thread/39112/Why-is-DOD-not-considering-Unmanned-Surface-Effect-Vehicles

http://cr4.globalspec.com/thread/39112/Why-is-DOD-not-considering-Unmanned-Surface-Effect-Vehicles

http://www.aerospaceweb.org/question/aerodynamics/q0130.shtml

http://www.freepatentsonline.com/2444332.pdf

https://www.fbo.gov/?s=opportunity&mode=form&id=825bc0a788e33d95e9f4dafa467e0e5a&tab=core&_cview=0

https://www.fbo.gov/?s=opportunity&mode=form&id=825bc0a788e33d95e9f4dafa467e0e5a&tab=core&_cview=0

All websites last accessed on 21th August 2011

Appendix A

Table 17: Drag Table

Cr / Diameter

Char Length

S_wet Re_f Re_c t/c or l/d

Q Cf D/q

Wing 1.873 1.608 38.481 9.937E+06 3.430E+06 0.105 1.222 0.0033 0.155H Tail 2.735 1.950 5.216 1.205E+07 4.212E+06 0.100 1.210 0.0038 0.024V Tail 1.325 0.984 5.163 6.082E+06 2.031E+06 0.100 1.210 0.0039 0.024Fuselage 1.100 10.000 32.970 6.179E+07 2.410E+07 9.091 1.103 0.0025 0.091Engine 0.530 1.330 1.772 8.218E+06 2.801E+06 2.509 4.803 0.0039 0.033Interference (8%) 0.028

Excrescences (5%) 0.018

Total 0.397

C_do 0.0196Drag At cruise

324.12

Table 18: Diameter Data

D Cp J n T1.8 0.957713 1.405436 0.4 475.93285761.9 0.730853 1.331466 0.55 654.4076792

2 0.56552 1.264893 0.67 797.18753652.1 0.4431 1.20466 0.725 862.6283045

Table 19: Thrust Curve Data

Cp sl 0.301309 Cp alt 0.56552

V np J T sl np alt T alt0 0 0 0 0

10 0.15 0.095238 2257.515 0.1 1505.0120 0.2 0.190476 1505.01 0.125 940.631430 0.275 0.285714 1379.593 0.175 877.922640 0.35 0.380952 1316.884 0.225 846.568250 0.425 0.47619 1279.259 0.275 827.755660 0.5 0.571429 1254.175 0.325 815.213870 0.55 0.666667 1182.508 0.375 806.255480 0.625 0.761905 1175.789 0.425 799.536690 0.7 0.857143 1170.563 0.475 794.3109

100 0.75 0.952381 1128.758 0.525 790.1303110 0.8 1.047619 1094.553 0.575 786.7099120 0.85 1.142857 1066.049 0.625 783.8595130 0.9 1.238095 1041.93 0.675 781.4476140 0.9 1.333333 967.5065 0.72 774.0052150 0.9 1.428571 903.0061 0.74 742.4717160 0.9 1.52381 846.5682 0.76 714.8798

Table 20: Drag Data

ALT Water Sea LevelV Cl Cd Drag Cl Cd Drag Cl Cd Drag

3 627.89218454.8

69132263.

1 0.48396 0.0287 327.2 395.0697306.15

7 83220.1

5 226.0412391.76

6 47615.0 0.174226 0.0192 606.8 142.225 946.893 29959.710 56.510 149.502 11905.1 0.043556 0.0178 2258.8 35.556 59.197 7492.020 14.128 9.361 2981.6 0.010889 0.0178 8993.1 8.889 3.716 1881.430 6.279 1.863 1335.4 0.00484 0.0178 20229.4 3.951 0.748 852.4

47.1 2.543 0.320 567.1 0.00196 0.0178 49947.5 1.600 0.138 387.050 2.260 0.257 511.5 0.001742 0.0178 56189.7 1.422 0.112 355.860 1.570 0.133 381.6 0.00121 0.0178 80912.8 0.988 0.063 289.070 1.153 0.080 312.2 0.000889 0.0178 110131.2 0.726 0.042 263.080 0.883 0.054 276.5 0.000681 0.0178 143844.6 0.556 0.032 260.990 0.698 0.041 261.5 0.000538 0.0178 182053.3 0.439 0.027 274.5

100 0.565 0.033 260.5 0.000436 0.0178 224757.1 0.356 0.024 299.7110 0.467 0.028 269.5 0.00036 0.0178 271956.0 0.294 0.022 333.9120 0.392 0.025 286.3 0.000302 0.0178 323650.1 0.247 0.021 375.7130 0.334 0.023 309.4 0.000258 0.0178 379839.4 0.210 0.020 424.2140 0.288 0.022 337.9 0.000222 0.0178 440523.8 0.181 0.019 478.7150 0.251 0.021 371.1 0.000194 0.0178 505703.3 0.158 0.019 539.0160 0.221 0.020 408.5 0.00017 0.0178 575377.9 0.139 0.019 604.6170 0.196 0.020 449.9 0.000151 0.0178 649547.7 0.123 0.018 675.5180 0.174 0.019 494.9 0.000134 0.0178 728212.7 0.110 0.018 751.3190 0.157 0.019 543.5 0.000121 0.0178 811372.8 0.098 0.018 832.1200 0.141 0.019 595.4 0.000109 0.0178 899028.0 0.089 0.018 917.8

Table 21: Weight Estimation

Table 22: Field Performance

Mass Nickolailb kg

Wing 933.3069884 423.3409h Tail 89.20648312 40.46338V Tail 107.2781175 48.66054

Fuselage 1782.654409 808.5984Engine 955.02 434.1

Landing Gear 212.1495877 96.22943Electric Engine 500 226.7962

Battery 97.00339536 44Furnishing 554.6757934 251.5967

Air conditioning

18.49431714 8.388881

Miscellaneous 220.4622622 100

Total 2482.174

Field performanceDo 112.5Δ Cl max 0.593111 b f/b w 0.4Δ Cd 0.023139 Sf/Sw 0.18Vs 40.26499 20D lo 92.86188 T 0 #REF!SGR 465.7834

Cl lo 1.219377 Cl obs 0.929883

Cd lo 0.107039 Cd obs 0.077899

Dlo 265.7539 D obs 193.4056

SGA 369.2795 tobs =tlo

To (curve) 1187.406

Table 23: Water Drag Data

V/Vmu V D/W D0.1 4.92 0.04 1800.2 9.84 0.09 4050.3 14.76 0.12 5400.4 19.68 0.16 7200.5 24.6 0.18 8100.6 29.52 0.15 6750.7 34.44 0.14 6300.8 39.36 0.14 6300.9 44.28 0.13 585

1 49.2 0.1 450

Appendix B

Table 24: Load Facor Data

C1 Empirical seaplane operations factor 0.012




CTO Empirical seaplane operations factor 0.004

C5 Empirical seaplane operations factor 0.0053Vso Stall speed, with flaps extended at landing position (kts) 24.26672575

Vs1 Stall speed, with flaps extended at take-off position (kts) 24.26672575

Angle of dead rise (rad) 0.436332313W Design landing weight 4059.989461K1 Empirical hull station weight factor 1.1

K2 Empirical hull station weight factor 1.2K Empirical hull station weight factor 0.8rx Ratio lx : radius of gyration(pitch) 0.85

ry Ratio ly : radius of gyration(roll) 0.5

lx Longitudinal distance from COG to position of where nw is applied m

ly Lateral distance from COG to plane of symmetry of float m

VF Volume of float (m2) 1.5

ρ Density (kg/m3) 1000Cx Coefficient of drag force 0.01236

Cy Coefficient of side force 0.00985

Appendix C

The following shows details of the structural analysis of the horizontal tail. The maximum applied load will be 3213 kgm. Using the equations detailed in Wing Structure, the shear flows are detailed in

Applied moment M max (Kgm) 3213.2Max Moment M max (Nm) 31521492Torsional stiffness T Nmm/rad 100000000Length of section L (mm) 1750Angle of rotation θ (rad) 1Torsional Regidity Ts (Nmm^2) 1.75E+11Required c (rad/mm) 0.000180123

AssumeG (Mpa) 28000t web (mm) 4 L web (mm) 60

Cell 1 Cell 2 Cell 3A 4000 18000 7000S 210 600 410

Q1 Q2 Q3 c ansT 8000 36000 14000 0.00E+00 3.15E+07Cell 1 126.997690

5-30 0 -2.24E+08 0.00E+00

Cell2 -30 337.1362587

-30 -1.01E+09 0.00E+00

Cell3 0 -30 219.3764434

-3.92E+08 0.00E+00

Aluminium Titanium SteelSkin t (mm) 2.16 2 3 1 1.37 0.8Q1 462.5009 453.6935 467.7741 434.7082 447.6074 433.8134Q2 615.1329 619.9106 612.1894 629.7277 623.1231 630.1766Q3 405.4786 398.2258 410.0344 383.8307 393.443 383.1876c 0.00018 0.000194 0.00013 0.000386 0.00018 0.00018

3030

(1) 100 150 150 200

150

Finding Shear centre Moments about (1)F applied/ S z (N) 13000qx ds 4563229.725

Shear centre x 351.0176712Applied Torque T 2.61E+06

Compatability Q1 Q2 Q3 ansCell 1 123.3 -15 0 -3611.11 Q1 -41.7685Cell 2 -15 280 15 -27083.33 Q2 -102.689Cell 3 0 -15 440 32138.88 Q3 69.54216

Q'1 Q'2 Q'3 c ansQ' for pure torque T 12000 36000 24000 0.00E+00 2.61E+06

Cell 1 81.66666667 -15 0 -3.36E+08 0.00E+00Cell2 -15 280 -15 -1.01E+09 0.00E+00Cell3 0 -15 206.6666667 -6.72E+08 0.00E+00

Stat det Flows I x 900000S z 13000Number CSA z q x F x F y q compatibility Fx F y Final q Fx Fy

Cell 1 1 100 -30 43.33 100 4333.33 1.56 156.48 45.46 4546.39cell 2 2 100 -30 86.67 150 13000.00 -16.02 -2403.39 20.58 3086.30cell 2 3 100 -30 130.00 150 19500.00 27.31 4096.61 63.91 9586.30cell 3 4 100 -30 173.33 200 34666.67 242.88 48575.10 274.91 54982.72cell 3 5 100 -30 216.67 60 13000 286.21 318.25cell 3 6 100 30 173.33 200 -

34666.67 242.88 -48575.10 274.91 -54982.72cell 2 7 100 30 130.00 150 -

19500.00 27.31 -4096.61 63.91 -9586.30cell 2 8 100 30 86.67 150 -

13000.00 -16.02 2403.39 20.58 -3086.30cell 1 9 100 30 43.33 100 -4333.33 1.56 -156.48 45.46 -4546.39cell 1 10 100 30 0.00 60 0.00 -41.77 2.13

60 vert q1 41.77 2506.11 -2.13 -127.8360 vert q2 60.92 3655.24 68.22 4093.3160 vert q3

-172.23-

10333.89 -167.67 -10060.3060 vert q4 286.21 17172.53 318.25 19094.82

F x (N) 0F y (N) 1300

0as applied

Appendix D

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