Design of an Unmanned Flood Relief Airship

and

Flight Testing of its 1:18 Scale Model

Submitted in Partial Fulfillment of the Requirements for the Degree

of

Bachelor of Technology

in

Aerospace Engineering

by

SATTWIK SUMAN DAS (SC08B108)

SHASHANK S. (SC08B098)

TANVEER ALI (SC08B003)

Department of Aerospace Engineering

Indian Institute of Space Science and Technology

Thiruvananthapuram

May 2012

BONAFIDE CERTIFICATE

This is to certify that this project report entitled “Design of an

Unmanned Flood Relief Airship and Flight Testing of its 1:18 Scale

Model” submitted to Indian Institute of Space Science and Technology,

Thiruvananthapuram, is a bonafide record of work done by SATTWIK

SUMAN DAS, SHASHANK S. and TANVEER ALI under my

supervision from 7/01/2012 to 2/05/2012.

Shri Pankaj Priyadarshi

Adjunct Professor,

Dept. of Aerospace Engineering,

IIST, Thiruvananthapuram

Countersigned by

Dr. K. Kurien Issac

Head of Department,

Dept. of Aerospace Engineering,

IIST, Thiruvananthapuram

Place

Date

Declaration by Authors

This is to declare that this report has been written by us. No part of the

report is plagiarized from other sources. All information included from

other sources have been duly acknowledged. We aver that if any part of the

report is found to be plagiarized, we are shall take full responsibility for it.

Sattwik Suman Das

SC08B108

Shashank S.

SC08B098

Tanveer Ali

SC08B003

Place

Date

Acknowledgement

The authors of this report wish to thank all the people who have contributed to this project.

We would like to begin by thanking our project guide, Prof. Pankaj Priyadarshi whose

continuous support and ideas have been of great help.

We would also like to thank the National Balloon Facility, TIFR Hyderabad, particularly

Prof. Devendra Ojha, Mr. B. Suneel Kumar and Mr. Sakram Korra for their help in

fabricating the envelope for the scale model and the hospitality extended during our stay.

Special thanks to Rahul Raj R., Lab tutor, Flight mechanics lab, IIST for his continued

support and immense help in fabrication of the scale model. The project would not be

completed in time without his help.

We express our thanks to the Purchase department, IIST for processing our purchase orders

in time, the Transport department, IIST particularly Mr. Krishnakumar for entertaining our

repeated vehicle requests to the city and the manufacturing workshop, IIST.

Our thanks also go to the faculty of NIDM and ILDM who have responded to our emails and

given us valuable information on flood relief. Lastly, thanks to our friends and family for

continued support throughout the project.

Abstract

Floods are a major natural disaster in third world countries, affecting millions of people over

the last few years. Current flood relief distribution methods are either slow (boats) or scarce

(helicopters are engaged in human rescue). This project discusses the design of an unmanned,

low cost, dual-gas, multi-chamber airship that can carry a relief payload of 2 tonnes to flood

affected areas per trip. The inside chamber of the airship contains hydrogen whereas the

outside chambers contain helium. The advantages of this configuration are low cost, higher

lifting capacity and enhanced safety and reliability due to multiple chambers. For the

airship’s propulsion system, diesel engines and fuel cells have been considered. Descent and

ascent of the airship by compressing on-board helium is explored along with pre-heating

helium in one chamber to compensate for increased buoyancy as fuel is consumed. To re-

ballast the airship after the payload drop, a novel method that uses flood water is proposed.

The envelope fineness ratio is optimised using ANSYS FluentTM

CFD for least drag and the

speed has been established for quick response and maximum fuel efficiency. Stress analysis

of the airship envelope with loading is also carried out for the optimized speed and shape. A

scaled prototype of this airship has been built and flight tested.

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TABLE OF CONTENTS

Abstract .................................................................................................................ii

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

2. Requirements capture ..................................................................................... 2

2.1 Payload estimation ................................................................................................ 2

2.2 Range ..................................................................................................................... 3

2.3 Altitude .................................................................................................................. 5

2.4 Speed Constraint ................................................................................................... 6

2.5 Mission Profile ...................................................................................................... 7

3. Conceptual configuration and layout .............................................................. 8

3.1 Envelope ................................................................................................................ 8

3.1.1 Shape selection .......................................................................................................... 8

3.1.2 Envelope configuration: .......................................................................................... 10

3.1.3 Envelope material selection: .................................................................................... 11

3.1.4 Inner envelope: ........................................................................................................ 12

3.1.5 Static discharges: ..................................................................................................... 13

3.1.6 Bubble wrap ............................................................................................................. 13

3.1.7 Probability of failure ................................................................................................ 14

3.1.8 Center of buoyancy change ..................................................................................... 15

3.2 Gondola ............................................................................................................... 15

3.3 Empennage Layout .............................................................................................. 16

3.4 Systems ............................................................................................................... 16

3.4.1 Lifting Gas Heating ................................................................................................. 17

3.4.2 Compressed Helium Storage ................................................................................... 19

3.4.3 Lifting Gas Venting ................................................................................................. 20

3.4.4 Recommendation ..................................................................................................... 20

3.4.5 Avionics ................................................................................................................... 20

3.4.6 Reballast .................................................................................................................. 21

4. Preliminary Design ....................................................................................... 24

4.1 First Weight Estimation ...................................................................................... 24

4.2 Empennage design .............................................................................................. 25

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4.3 Preliminary weight estimation ............................................................................ 25

4.3.1 Envelope fabric ........................................................................................................ 25

4.3.2 Air lines ................................................................................................................... 26

4.3.3 Catenary ................................................................................................................... 26

4.3.4 Patches and Reinforcements .................................................................................... 26

4.3.5 Nose Reinforcements ............................................................................................... 26

4.3.6 Tail Structure ........................................................................................................... 26

4.3.7 Gondola ................................................................................................................... 26

4.4 Propulsion ........................................................................................................... 28

4.4.1 Fuel .......................................................................................................................... 28

4.4.2 Engine ...................................................................................................................... 29

4.4.4 Speed Optimization ................................................................................................. 32

5. Flow Studies ................................................................................................. 35

6. Stress Analysis .............................................................................................. 38

7. Building indoor airship envelopes ................................................................ 42

7.1 Shape of the envelope ......................................................................................... 42

7.2 Materials used ..................................................................................................... 42

7.3 Fabrication ........................................................................................................... 42

8. Preliminary design of the ALFRA-1 scale model ........................................ 45

8.1 Weight Estimation ............................................................................................... 45

8.2 Envelope .............................................................................................................. 45

8.3 Lifting Gas .......................................................................................................... 47

8.4 Propulsion system ............................................................................................... 47

8.5 Gondola ............................................................................................................... 50

8.6 Empennage design .............................................................................................. 51

8.7 Free Body Diagram of Airship ............................................................................ 52

9. Fabrication of the scale model ...................................................................... 55

9.1 Envelope .............................................................................................................. 55

9.2 Heat sealing machines ......................................................................................... 55

9.3 Stabilizers ............................................................................................................ 58

9.4 Gondola ............................................................................................................... 60

9.5 Revised weight estimation .................................................................................. 61

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10. Flight Testing of Scale Model ................................................................... 62

10.1 Flight Test ......................................................................................................... 62

10.2 Flight parameters ............................................................................................... 63

10.2.1 Turn Rate: .............................................................................................................. 63

10.2.2 Speed: .................................................................................................................... 63

10.2.3 Rate of climb : ..................................................................................................... 64

11. Recommendations ..................................................................................... 66

Appendix 1 .......................................................................................................... 67

Appendix 2 .......................................................................................................... 68

Appendix 3 .......................................................................................................... 69

Appendix 4 .......................................................................................................... 74

References ........................................................................................................... 77

Page | 1

1. Introduction

Floods have been the only major natural disaster in India that has occurred with an unfailing

regularity. Logistics is one of the most important factors in humanitarian aid operations, as

logistics efforts account for 80% of disaster relief [1]

. Currently, flood relief is distributed

through boats and helicopters of the Indian Navy, Air Force and the National/State Disaster

Management Authority. Even though these responses are effective, they face problems of

timely deployment and cost. Also, helicopters are actively engaged in human rescue and

hence relief distribution using helicopters on a large scale is not possible. Aircrafts, which are

the next logical choice for transport by air, suffer from the drawbacks of accurate payload

drop and hovering ability. Thus, airship as a mode of relief and aid distribution during floods

is investigated.

This report gives the results of the design exercise of an autonomous low cost airship;

Autonomous Low cost Flood Relief Airship-ALFRA-1 to airdrop relief materials during

severe floods when access by other means of transport is not available or restricted.

The present B.Tech project was carried out in six phases.

Phase I: The team dealt with requirements capture, configuration of the airship and

weight estimation.

Phase II: The preliminary design of the airship was done and several sub system

options were debated. Several innovative concepts were thought of and their

feasibility was explored.

Phase III: It dealt with CFD simulation and trade-off studies. Final parameters of the

airship were fixed and cost analysis was worked out.

Phase IV: Hands on experience in building small airship models was gained. This

was necessary for fabricating the scale model of the airship.

Phase V: The design parameters were scaled down and the scale model of ALFRA-1

was built.

Phase VI: The scale model was flight tested and various flight parameters were

recorded.

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2. Requirements capture

The requirements of the autonomous airship, ALFRA-1, in terms of range, payload mass,

speed, altitude etc. have been arrived at, based on the public domain data available on the

floods in India over the last ten years.

2.1 Payload estimation

Payload estimation is the most critical factor in the design of ALFRA-1 as it has a direct

implication on the size of the airship. For identifying the relief materials generally distributed

in India, experts from Institute of Land and Disaster Management, Kerala and National

Institute of Disaster Management, New Delhi were contacted and a list of relief materials was

prepared based on the above inputs [2]

. Medecins Sans Frontieres (MSF) reports [3]

on past

floods in India were also referred while creating the list.

Table 2.1: Food materials airdropped using 3 helicopters in 40

sorties [26]

(Source: UNDP Situation report, Bihar floods 2008)

Sl. No Material Quantity (kg)

1 Chura (Flattened rice) 49620

2 Sattu (Ground cereal) 16540

3 Salt 8270

4 Gurr (Jaggery) 8270

In addition to the food items mentioned in the above table, MSF also states that plastic

sheeting, oral rehydration solutions, water purifying tablets, bleaching powder and buckets

are also provided.

According to the Ramakrishna Mission’s relief division [4]

, 1 ton of relief articles are given

for every 1000 families. Assuming each family to have four members, a conservative

estimate, considering rural India, it means 1 ton of relief is suitable for 4000 people. In case

of a flood, the most severely affected people are marooned in the small villages. In India,

small villages are the Type 4 towns with population of around 8000. [5]

In the 2008 Kosi

floods, one or two Type 4 towns are completely cut-off and transportation by traditional

Page | 3

routes is impossible to achieve [6]

. So, the relief material required for one Tier 4 town comes

to 2 tons and the same is fixed as the payload of the ALFRA-1.

2.2 Range

The airship base of operation should be able to cover the entire flood region and should also

be at a higher elevation. It should be big enough to have adequate relief stock in place.

According to experts from the Institute of Land and Disaster Management (ILDM) consulted

for this project, the District HQ is the key centre for relief operations which are spearheaded

by the District Magistrate [2]

. Hence, the range is calculated as the distance from base to

farthest district HQ.

It is seen from the historical data that 40 million square hectares or 1/8th the geographical

area of India is flood-prone [7]. As the flood-prone area is too large for a detailed analysis,

two case studies are presented here.

Figure 2.1 shows flood zones in Bihar. It is primarily based on the Kosi floods of 2008. From

the figure it is clear that, if two bases were selected at Muzzafarpur (Population: 3,50,000)

and Purnia (Population: 2,80,000), the flood zones in the entire state can be covered. Aerial

distance to the farthest district headquarter is measured and is found to be 120 km.

Figure 2.1: Map showing flood zones in Bihar [26]

(Source: UNDP Situation report, Bihar floods

2008)

Page | 4

Figure 2.2: Bihar district map with airship stations and range [27]

A similar analysis was carried out for the Andhra Pradesh and Karnataka floods of 2009 and

the results are summarised in the below map.

Figure 2.3: Andhra Pradesh map [28]

with airship stations and range

Page | 5

Kurnool and Rajahmundry are chosen as the centres and the airships can be stationed there.

Both these towns have population above 2,00,000. The aerial distance to the farthest district

headquarter was 170 km (Rajahmundry centre to Vizianagram).

The results are summarised in table 2.2.

Table 2.2: Summary of results

State Centre Farthest Affected

District HQ

Distance (km)

Bihar Muzzafarpur Bettiah 120

Andhra Pradesh Rajahmundry Vizianagaram 170

The above two case studies are for two of the largest floods that India has seen in its recent

past in terms of inundated area and people affected. Thus, it was inferred that for any other

state in India, as the inundated area has been historically smaller, two bases are sufficient to

carry out the relief operations and the total maximum distance to be travelled is 340 km.

Taking a margin on this value, the range is fixed at 500km.

2.3 Altitude

From the composite map given in Appendix 2, Andhra Pradesh, Bihar and almost all the

other flood prone states are plain and do not have any mountains. From the map, the

maximum elevation is obtained as 500m above sea level. So, it is assumed that the terrain

will be flat and given the rural demographic, buildings are assumed to be lower than 50 m (15

storeys).

Hence the only structure with significant height is the transmission line towers that are

common across the Indian landscape irrespective of the state. The most commonly used

towers are made by Bajaj Electricals and are 42 metres tall [8]

. Cell phone towers are of lesser

height than these towers.

Taking the above heights of transmission lines and buildings, the cruise altitude is fixed at

600m.

Page | 6

Figure 2.4: Scene from flooded north Bihar (Source: Wikipedia)

2.4 Speed Constraint

The disaster management experts communicated that the relief should reach the affected

people within 24 hours. The flood inundation map is critical for the decision maker to

manage relief operations better. According to NRSC Hyderabad [9]

, “Rapid flood damage

assessment is carried out by integrating the satellite derived flood inundation layer with the

district database such as village boundaries, transport network, land use/land cover etc.

District-wise flood inundated area statistics, crop area submerged, villages marooned, road

and rail network submerged etc. are estimated. All this information is generated within 5

hours after satellite data acquisition”. At this point, officials will take the decision whether

the airship needs to be deployed or not. It is assumed that 12 hours are required from the first

information of flood to the final decision regarding deployment of airship.

A conservative estimate of six hours is made for time taken to fill the airship and pack the

relief materials into the container. 8 hours are assumed to be required to fix the drop points

and get a team at the affected area ready to receive it. This leaves 4 hours for the airship to

reach the affected area. The farthest distance, according to Table 2.2, is 170 km.

Approximating it to 200 km, the airship needs to cover this distance in under 4 hours. So, the

speed required comes to 50kmph.

Page | 7

Figure 2.5: Estimated emergency response times in case of flood

2.5 Mission Profile

The mission profile is shown in Figure 2.6. The airship, which is moored, is loaded with the

relief payload after the lifting gas has been filled. The engines are started and a sloping ascent

towards the cruising height begins. As the cruising height is reached, the thrust is vectored in

the forward direction. The airship cruises to the drop point. It descends once it has reached

the drop point. The reballast operation takes place and the payload is dropped. The airship

flies back to base.

Figure 2.6: ALFRA-1 mission profile

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3. Conceptual configuration and layout

3.1 Envelope

3.1.1 Shape selection

Envelope is the lifting gas container in the airship, the huge balloon which generates all the

buoyant lift to keep the airship in air. There are three disciplines that directly influence

airship envelope design. They are aerodynamics, structures and weight and balance. There

are broadly three different shapes that are mostly widely used, or have been optimized for

optimal performance, and they are:

NPL shape

GNVR shape

Multidisciplinary optimized shape from IIT, Bombay[10]

NPL shape, with an ellipse at the fore and a parabola at the aft*, was designed by the National

Physical Laboratory in England. The GNVR shape was better optimized for aerodynamics

and structural loading. It is shown in the diagram below. A study conducted at IIT Bombay[10]

improved the GNVR shape by dividing the contour between the ellipse at the fore end and the

parabola at the aft end into larger number of segments, introducing splines and carrying out

an optimization for aerodynamics, structures and manufacturability.

Figure 3.1: The NPL shape

*This shape has been reproduced wrongly in [12] with an ellipse instead of a parabola. This was pointed out to

the author and he has acknowledged it [24]

Page | 9

Equation of the ellipse:

Equation of the parabola:

Figure 3.2: The GNVR shape [10]

The airship hull is known to create the maximum drag in an airship. This is owing to its large

surface area, which increases the skin friction drag and also the large frontal area it projects,

which increases the pressure drag. The National Physical Laboratory’s shape is chosen as the

airship envelope shape due to the possibility of optimization of fineness ratio (length to

diameter ratio) for this shape. The airship fineness ratio is a variable parameter for the

volume and hence, CFD simulation is performed using ANSYS FluentTM

to optimize for the

least drag.

The estimated maximum take-off weight represents the maximum buoyant lift required for

the airship to stay afloat. As the lifting capacity of different gases such as hydrogen and

helium are already known, the weight is indicative of the volume of envelope required to

generate the required lift. Using the above equations, the volume for the body of revolution

was calculated and for a particular choice of fineness ratio the dimensions of the envelope

was obtained.

Page | 10

A detailed study on the material to be used for envelope fabrication was also done. Generally

a three layered fabric is used for making the envelope to cater the weatherability, gas

retention and load carrying needs. The details about the same are discussed in the following

chapters.

For our airship, a novel design of the envelope [11]

was employed to enhance safety and

reliability. A double envelope design with the inner envelope carrying hydrogen and the outer

envelope carrying helium was designed. The helium chamber was further divided into a

number of chambers to increase the reliability. This ensured that any leakage in the hydrogen

envelope results in interaction of hydrogen with helium and not air and hence eliminating a

chance of any accidental hydrogen combustion.

Figure 3.3: Double envelope configuration [11]

3.1.2 Envelope configuration:

As proposed by [11], the multi chamber configuration ensures increased safety of the setup.

The inner chamber is unlikely to get damaged due to the layer of Helium surrounding it. Even

if the inner chamber leaks, it will only mix with helium, an inert gas. Multiple Helium

chambers outside ensure that even if one chamber develops a leak, the overall shape of the

airship will be maintained.

Page | 11

Figure 3.4: Double envelope configuration proposed by [11]

3.1.3 Envelope material selection:

According to [12], materials ideal for non rigid airship envelopes have the properties:

1) High Strength to weight ratio

2) Resistance to environmental degradation and low permeability

3) High tear resistance to give damage tolerance

4) Low creep and easy to join

For the ALFRA-1, cost and ease of maintanence were important factors apart from the ones

listed above that went into deciding the envelope material. An existing envelope material is

used due to simplicity and maturity of technology. Generally, airship materials consist of

three layers, an environmental protection layer, a load bearing layer and a gas retention layer.

Table3.1: Comparison of envelope material characteristics

Airship ZPG 3W Goodyear GZ 20 Skyship 600 Zeppelin NT

Type Non Rigid Non Rigid Non Rigid Semi Rigid

Volume (m3) 42500 5380 7600 8425

Envelope material 2 ply

neoprene

coated

polyester

fabric

2 ply neoprene

coated polyester

fabric

Single ply

Polyether grade

polyurethane

coated polyester

fabric with Saran

on inner surface

Polyester fabric

coated with

TedlarTM

and

Polyurethane

for gas

retention

Page | 12

Material Breaking

strength (kN/50mm)

2.8 1.45 1.85 1.425

Material weight

(g/m2)

560

370 380 250

Finally, Polyester fabric coated with TedlarTM

for weatherability and Polyurethane as the gas

retention layer was chosen as the ALFRA 1 envelope material. This material has already been

tested on the Zeppelin NT with good results.The primary factor that influenced the choice of

this material was the low material weight (250 g/sqm).

TedlarTM

(Polyvinylfluoride film) marketed by Dupont has been proven to be resistant to a

wide range of acids, alkalis and solvents at ambient temperatures.[12]

Figure 3.5: Cross section of the envelope material

3.1.4 Inner envelope:

The only criteria important for the inner envelope is Hydrogen permeability. As it is not

exposed to the atmosphere, there is no need for the weathering protection layer. Similarly, it

is not necessary to have a woven load bearing layer. Polyurethame is selected as the material

for the inner envelope due to its durability and proven gas retention capability.

Page | 13

3.1.5 Static discharges:

Friction during flight between the atmosphere and the outer material creates a lot of static

electricity. As the envelope and the stabilisers are not conducting surfaces, a large number of

static dischargers are required to be fixed on them.

3.1.6 Bubble wrap

As there is no constraint on the amount of Helium in the outer chambers, an idea inspired by

bubble wrap is discussed below.

Bubble wrap is the trademarked name for a packing material consisting of two plastic sheets

laminated together in a way that traps air bubbles in small, uniform pockets. This plastic

sheet assembly is used as a flexible cushion to protect fragile objects during storage or

shipping. Cushioning laminate is primarily made of plastic film or a thin sheet formed from

resins such as polyethylene and polypropylene. These resins are widely used because they

perform well and are relatively inexpensive. They can be cast into strong, flexible films,

which have the ability to hold air without leaking.

In earlier discussions it was seen how no constraint can be put on the size of the inner

hydrogen envelope. The more the hydrogen used the better it is in terms of lift generated as

well as cost (hydrogen is inexpensive). The weight penalty for a large inner envelope is more

Figure 3.6: Bubble-wrap manufacturing (Source: Madehow.com)

Page | 14

than compensated for by the increase of lift. So, the possibility of helium filled bubble-wrap

to be used in place of the outer helium envelope was explored.

Polyethylene has good gas retention properties and is widely used to make sounding balloons.

Manufacturing bubble-wraps in a helium environment will trap helium in the bubbles. Two of

these sheets can be further laminated, with the blistered sides facing each other, suitably

offset such that the bubbles get closely packed. This laminate can be then be used as the

hydrogen retention layer in the three layered envelope fabric.

Figure 3.7: Cross section of proposed envelope with bubble wrap

This configuration of the airship envelope material has several advantages. This will reduce

the amount of helium being used, thus making the operation of the airship more cost

effective. The arrangement ensures that there is a layer of helium around the hydrogen and

thus no single leakage will lead to direct mixing of hydrogen and air. There being a large

number of helium chambers (bubbles) around hydrogen, it increases the reliability of the

whole system. If any bubble gets punctured, it has a very small impact on the whole system.

A disadvantage in using the bubble-wrap based material is that the helium can diffuse out

gradually from the bubbles. If this happens, it will be impossible to refill the bubbles. Thus

either the whole material has to be changed or the layer of bubble wrap has to be replaced.

3.1.7 Probability of failure

Let the probability of an airship envelope becoming punctured be x

Then, for a six chamber envelope, the probability of a puncture in one envelope = 1/6 x

The probability of failure due to hydrogen chamber puncture in the already punctured helium

chamber = 1/36 x

Page | 15

Therefore,

Thus it is clear that as the number of chambers increases the reliability factor goes up. In case

we use the bubble wrap method, the reliability factor increases many times.

3.1.8 Center of buoyancy change

Suppose the inner envelope has a volume of 4000 m3 and develops a leak into one of the six

helium chambers. Each helium chamber has a volume of 666 m3. The mixing of the two

gases takes place until the concentrations are equal. The ratio of the gases in the mixture will

be in proportion to the intital volume. Thus, it will be a 1:6 ratio of helium to hydrogen. This

mixture will shift the center of buoyancy and hence, may cause stability problems. An

increase in the number of chambers will bring the shift to manageable levels. In case the

bubble wrap idea is implemented, there will be no such problem.

3.2 Gondola

The gondola houses the propulsion system, the fuel tanks and the avionics bay. In the

preliminary design phase, different possible propulsion layouts for the airship were explored.

1. Two propellers for forward motion and one propeller to give a vertical thrust, each

propeller powered by independent engines: The advantage was that the vertical and

horizontal components of thrust could be independently controlled and would make

the control system simpler.

2. Two propellers with thrust vectoring (for pitch up and pitch down): The advantage

here is there was no extra engine as in the previous case making it lighter as well as

minimizing the fuel consumption. The drawback is the control is not discrete and the

angular variation needs to be finely controlled by the control system, which tends to

make it complicated.

Yawing in both the above cases was to be achieved by control surfaces present on the tail.

However at low speeds, particularly during mooring or payload drop, the velocity of the

airship would be too low for control surfaces to be effective. So, in those cases, differential

thrust was to be used to give yawing moment.

Page | 16

The second configuration with thrust vectoring was chosen as it can allow weight saving and

is fairly simple to implement

The payload comprises of 2 ton of relief material for distribution in flood affected areas. For

carrying the payload, a 2 ton container was required. This container requires to be housed in

the gondola.

The avionics system comprises of a navigation, guidance and control module and is primarily

GPS based.

The fuel system in ALFRA consists of two fuel tanks-located aft and fore of the main engine.

A fuel pump is installed for c.g control.

3.3 Empennage Layout

There are three different empennage designs that could be used as shown in the diagram

below. The three fin configuration had the advantage of lower structural load on the

empennage. However, the four-fin plus configuration was chosen as it keeps the yaw and roll

controls independent and thus would be convenient in terms of the control system design

Figure 3.8: Empennage configurations

3.4 Systems

Systems form an integral part of every vehicle. They are the physical means of achieving a

designed function [12]

. Some of the major systems in an airship are covered in this section.

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3.4.1 Lifting Gas Heating

Buoyancy control during flight is one of the prime challenges faced by airship designers. The

prime change in buoyancy is due to the consumption of fuel. Many methods to control

buoyancy have been proposed and tested out on airships. They are listed below.

Table3.2: Methods for buoyancy control in airships

Sl.

No.

Method Comments

1 Using Dynamic Lift Requires a hybrid airship design

2 Reducing buoyancy by

venting lifting gas

Helium is too expensive to vent out and venting out

hydrogen is a safety concern

3 Changing the density of

lifting gas by heating or

cooling

Explored here

4 Use of thrust vectoring Additional fuel needs to be carried

5 Water extraction from

exhaust

Bulky and heavy apparatus leading to weight penalty.

Not suitable for airships of this class.

The concept of lifting gas heating was first proposed by Burgess in his book “Airship

Design” [13]

. With the materials available then, he calculated the rate of heat loss and

concluded that such a system would be impractical unless the convective heat loss can be

reduced. In 1987, Ray Maurice [14]

proposed a simple heat transfer model for heated helium

airships in which he concluded that a 30% increase in gross lift could be achieved by utilizing

the engine exhaust for heating helium.

For ALFRA-1, the following buoyancy control model was explored:

1) One of the chambers of the airship is insulated with a layer of insulation material. The

insulation material thickness is decided by the thermal conductivity required to

Page | 18

provide a heat loss rate which exactly compensates the reduction in weight due to the

fuel consumed.

2) Electric heating elements driven from the generators attached to the engines are

affixed to the helium chamber.

3) Just before take-off, the airship is made heavier than air and heating elements increase

the temperature of helium in the chamber to make the airship lighter than air.

4) As the airship cruises, fuel weight decreases which increases net buoyancy. But, the

lifting gas also cools down and hence, decreases buoyancy leading to a state of neutral

buoyancy. Finer control is achieved through the auxiliary fans.

Calculations:

Using ideal gas law, we get (

)

From [1], the typical specific fuel consumption of a diesel engine is 30 kg/hr.

For an airship of 8000 m3, assuming that one helium chamber volume is 700 m

3, the required

temperature that the helium should be heated for a three-hour flight is calculated to be 471.4

K.

Energy required to heat helium,

This is an extremely high temperature and the high specific heat of helium (5188 J/kgK)

makes the necessary energy required to heat it impractical to supply. Increasing the chamber

volume on the other hand will bring down the temperature, but will increase the effective cost

of the airship.

Hence, the heating of helium for this kind of an airship is not a viable option.

Page | 19

3.4.2 Compressed Helium Storage

For the descent stage, the helium inside a chamber is compressed into a storage tank in the

gondola, thereby decreasing lift. After the payload drop and reballast, the helium is pumped

back into the chamber providing lift and thus enabling ascent.

Calculations:

Acceleration,

TAR[15]

specifies the maximum descent rate of 7 m/s. Assuming that this descent rate is

achieved at the ground, we get

Applying force balance in the z-direction, we get

As the helium is being compressed and stored on-board, the mass is conserved and thus, the

weight remains the same.

Using the first estimate of mass,

Reduction in lift should be, 1960 N.

The lifting capacity of helium is 10.32 N/m3.

Thus, Volume of Helium that needs to be compressed,

.

In terms of weight, this turns out to be 33.82 kg.

Z

X

Page | 20

If the storage of helium is done in long cylindrical tubes, then the minimum weight of the

tubing, irrespective of its dimensions or pressure ratio, is given approximately by [12]

Where

The bare weight of the cylindrical tubing assuming a density of 1800 kg/m3 and working

tensile strength of 3x108 N/m

2 would thus be,

The weight of the compressor system and the required piping would bring the total weight of

the system to well over 300 kg. The weight factor makes this system undesirable for

application in the airship unless composite materials become affordable.

3.4.3 Lifting Gas Venting

The idea of venting lifting gas has been used in airships such as the R-101, Hindenberg etc.

Helium, if vented, becomes expensive to replenish. The option of venting hydrogen is not

without safety concerns, mostly due to ignition by a lightning strike.

3.4.4 Recommendation

Considering the various buoyant control options available in sections 3.4.1 to 3.4.3, the

method of thrust vectoring is recommended, mainly due to its simplicity in implementation

and the extensive use in airships of similar class.

3.4.5 Avionics

From a safety point of view, airship regulations [16]

forbid the presence of humans in an

airship which uses hydrogen as its lifting gas. To conform to these regulations, ALFRA is

made unmanned. It is capable of delivering the relief payload to the location using only the

GPS co-ordinates as its input. This requires a sophisticated navigation, guidance and control

system. A detailed discussion of this system is out of the scope of this project.

Page | 21

CS Jin [17]

has created an autonomous GPS navigation module for an indoor airship with

variable line of propulsion. An improvement of this model which accounts for wind

compensation is applied to ALFRA-1.

Figure3.9: Block diagram of the algorithm

In order to reduce the power consumption of the processor inside the airship, the less

intensive implicit guidance algorithm is used.

Apart from this system, ALFRA-1 avionics also contains a pressure detection system in each

chamber. In the event of a leak in one chamber, the pressure drop is detected and auxiliary

power is increased to counter it without the interference of the NGC system. This is required

because of the use of implicit guidance algorithm, which fails if the airship strays too far

from the pre-determined flight path.

3.4.6 Reballast

The ‘payload drop’ problem is a unique problem to airships, which has in a way hindered the

large scale use of airships for cargo transportation. The problem is stated as “after the payload

has been dropped, the airship’s net buoyancy becomes very high and the return flight cannot

be carried out”.

There is no fool-proof method currently available for this problem. Proposed methods

involve compressing lifting gas, thrust vectoring etc. But, the economics of using such a

system for compensating a weight difference of 2 tons is questionable.

A system of pumps and pipes was carried for pumping in flood water weighing 2 ton into

water tanks carried on board. The payload was going to be dropped in flood water and thus it

was essential to ensure that it does not sink but stays afloat till the relief workers in the area

Target GPS Co-

ordinates

Navigation

Module

Guidance

Module

Control Module

GPS and

Attitude

Sensors

Engines and

actuators

Clock

Page | 22

reach it and collect it for distribution. For this purpose an inflatable buoy was attached to the

payload containers.

The detailed manoeuvre of payload drop is explained below

Airship approaches the region where payload can be dropped.

Descents towards the surface

Buoy attached to payload containers inflated

Pipes deployed and pumps started to fill the tanks

Payload is dropped

Airship takes off and returns to base

As the buoy is designed to carry the weight of the containers, i.e., 2 ton, it can carry the

weight of the airship when it is being filled with water. This is because, the payload being

still attached to the airship,

The net weight carried by the buoy = the weight of the airship + the weight of the water in

the tanks (that would be a maximum of 2 ton) - the lift generated by the envelope (which

equals to the weight of the airship).

So, effectively, the buoy has to carry a maximum weight of 2 tons and not more.

Page | 23

Figure 3.10: Figure detailing airship manoeuvres

A standard two ton freight container was considered for an approximate size and weight

estimation of the payload bay. The average dimensions for a container with a load carrying

capacity of 2 ton would be around 2m x2.5mx 1.5m and weighed around 550kg. These

containers are typically made of weathering steel, to withstand the high density cargo.

However, in our case wooden containers can also be considered.

Similar study for standard water tanks available in market showed that a 2000litre capacity of

water tank would weigh close to120 kg.

Page | 24

4. Preliminary Design

4.1 First Weight Estimation

Statistical weight estimation was done in order to have an estimate of the maximum take-off

weight for the given payload. Statistical data was collected for various non-rigid airships and

a graph was plotted between the payload weight and the take-off weight for these airships.

This graph gave us a value of the maximum take-off weight for our airship to start with.

The table below shows the values of the payload and take-off weight for airships in the class

of the ALFRA-1.

Table 4.1: Comparison of airships in the ALFRA-1 class

Sl. No Name Payload MTOW Volume

1 Zepellin NT 1900 8040 (semi rigid airship)

2 Skyship 600 13 people= 1 tonne 7100 6667 cubic meters

3 ZPG 3W 26 people= 2 tonne 10150 42475 cubic meters

4 ABC lightship

spectator A

1026 3808

5 ABC lightship

spectator B

1546 4394

6 WDL-1B 1180 5100

7 Augur MD 900 3950 8630

Page | 25

Figure4.1: 2nd

order polynomial fit of MTOW vs. Payload weight

From the above graph it was seen that an airship designed to carry a payload weight of 2000

Kg would have a maximum take-off weight of approximately 8808 Kg.

4.2 Empennage design

Owing to the length of the airship hull, the boundary layer is thick at the stern. A formula for

the thickness of boundary layer at any section of the envelope, given by [12]

S=0.02 x for x<xm

S=0.02 x (d/dx) for x>xm

Where, d is the maximum body diameter occurring at x=xm and s is the boundary layer

thickness. A conservative value of fin dimension was obtained here so that the weight of the

tail structure could be estimated. The fin length was taken as four times the length of the

boundary layer and was computed at 1.12m.

4.3 Preliminary weight estimation

4.3.1 Envelope fabric

The fabric weight is the surface area of the fabric multiplied by the surface density of the

material used. The surface area of the envelope for the lifting the earlier estimated maximum

take-off weight was calculated by integration of the equation of the profile of the envelope.

The net area density of the material with allowances for seam and patches should be in the

Page | 26

range of 0.35-0.52kg/m2.

According to the table given in [12], for the calculated volume, the

surface density of the material was obtained to be 0.35kg/m2.

The net volume of the envelope required to carry the maximum weight estimate of 8808 kg is

8350m3. From this value of and the dimensions of the envelope were computed and then

integration was carried out to obtain the surface area of the envelope, which turns out to be

2432.7m3.

The weight of the envelope fabric=surface density of envelope material x surface area of the

envelope=852Kg.

4.3.2 Air lines

The air lines were to be provided to allow filling inner chambers as well as the outer helium

envelopes. The weight of the airlines was estimated to be 3% of the fabric weight [1] and was

computed to 30kg.

4.3.3 Catenary

Catenary screens are provided to suspend the Gondola. As the ALFRA had a twin envelope

design and the catenary screens provided would have complexity in design, the catenary

weight was estimated towards the higher side.

The computed value of catenary is 15% of envelope fabric weight=128Kg.

4.3.4 Patches and Reinforcements

From the empirical formula given in [12], the weight of the patches and reinforcements were

estimated to be 5% of envelope fabric weight= 43Kg.

4.3.5 Nose Reinforcements

Nose reinforcements include the nose cone, mooring probe etc. fixed at the nose of the

envelope. The allowance that is to be given for nose reinforcements, according to [12] is

computed as 180Kg.

4.3.6 Tail Structure

According to [12] a survey over a period of two decades gave a structural weight density for

fins=6Kg/m2. This was including control surfaces. So for the four fins the weight of the

weight was estimated=781Kg.

4.3.7 Gondola

The gondola had several components like the container for holding the payload, the water

tanks, the engines, the fuel tanks, the transmission system, the propellers and ducts, the fuel

Page | 27

tanks and the structural framework or a chassis to which all things were mounted. The weight

of the containers and water tank were based on the weight of the standard freight containers

and water tanks available in the market. The engine, transmission and fuel weight were

estimated using empirical formulas given in [12]. The values of all these components have

been tabulated in table 4.2.

Table 4.2: Estimated weight distribution

Component Weight (Kg)

Envelope fabric 851.5

Catenary 128

Airlines 30

Patches reinforcements 43

Nose reinforcements 180

Suspension 170

Tail structure 781

Container 550

Water tank 120

Engine 450

Fuel 650

Transmission 100

Propellers and ducts 200

Pipes and pumps 20

Structural frame 300

Lifting gas 1450

Payload 2000

TOTAL 8024

Page | 28

Figure 4.2: Estimated weight distribution

4.4 Propulsion

The propulsion system consists of three components viz. the fuel, engine and the propulsor.

The range of airship operation varies from using 20% of engine power at hovering to 80% of

engine power at full speed. Thus, gas turbine engines which are economical to operate at

close to 90% cannot be used for airships.

4.4.1 Fuel

Considering that gas turbines are not used, the options for the fuel are aviation gasoline,

mogas (petrol), diesel and kerosene (jet fuel).

Aviation gasoline: This is the most widely used fuel for piston driven aircraft engine. Many

ultralights and general aviation aircraft have engines which use this fuel. The main drawback

of this fuel is its skeptical availability in the near future as the reserves are running out. The

fuel is also expensive in most parts of the world.

Petrol: This is the automotive form of aviation gasoline, which is used on certain aircrafts. It

has the advantage of being easily available and being relatively cheaper compared to avgas.

The FAA does not issue certification for aircrafts which use this fuel and take passengers for

hire.

18%

12%

25%

18%

10%

17%

envelope

gondola

payload

lifting gas

tail structure

propulsion

Page | 29

Diesel: It is the most widely available fuel, which also has the added advantage of being

cheap. It has high energy content. The airships of the previous era have flown with diesel as

the fuel.

Jet Fuel: Its availability is restricted to airports and the cost per litre is slightly lower when

compared to petrol. This fuel is highly inflammable and has a lower energy content than

diesel.

In order to choose the type of fuel, a comparison was done with weightages given to the

various desirable parameters of the fuel. The weightages were given on a scale of 1-5 and the

scoring was done on a similar scale with 1 being the lowest and 5 being the highest. The total

score was calculated and the fuel decided based on this.

Table 4.3: Comparison of fuel options.

Fuel Aviation

Gasoline

Petrol Diesel Jet

Fuel

Weightage

Parameters

Availability of fuel at

base station

1 4 5 2 4

Cost of fuel 3 1 5 2 5

Engine Options 4 1 2 3 2

Safety 3 2 4 3 4

Energy Content 3 3 4 3 3

Total Score 48 40 77 45

Diesel is chosen as the fuel due to its low cost and wide availability.

4.4.2 Engine

The aero-diesels, after the advent of gas turbines were phased out. There has been revived

interest in them due to the possible scarcity of avgas in the future. As they are being

developed, the certified engines are a handful in number. In order to select an engine, we

need an estimate for the power required from the engine.

From the drag estimation of the hull (see chapter 5), for a fineness ratio of 4.5, we have

Assuming an average speed of 20 m/s, we get

Page | 30

But, this is only the drag of the hull. From historical data [12]

, the drag of the hull is known to

be 50% of the drag of the entire airship.

Hence, drag of the airship,

Total power required to just overcome this drag,

Let efficiency of the engine system be 85%.

Then, total power required becomes,

As the power required is low, the option is available to use either one engine or two engines.

To provide redundancy and because the weight penalty was not very significant, the two

engine configuration was chosen. In addition to this, the two engine configuration provides

less complex low speed yaw control.

Estimated power required per engine,

Engine options available:

DeltaHawk’s DH160V4, Wilksch Airmotive’s WAM-120 and Thielert’s Centurion 2.0 are

the available aero-diesel engines in this class.

Table4.4: Comparison of available engines.

Name DH160V4 WAM-120 Centurion 2.0

Max. Power 160 hp 120 hp 135 hp

Best SFC 21.22 kg/hr 13 kg/hr 15.34 kg/hr

Weight 148.3 kg 127.2 kg 134 kg

Economy Power 84 hp 67 hp 97 hp

Certified? non-certified In process certified

Price in USD 29500 23700 59575

Page | 31

Considering the cost and the power required, the Wilksch Airmotive’s WAM-120 is selected.

The WAM-120 is a 3 cylinder, 2-stroke, compression ignition, liquid cooled engine of

inverted configuration with pressure fed lubrication and integral sump.

Performance

Table4.5: Performance parameters of the WAM-120

Power Torque Engine rpm Specific Fuel

Consumption

Max 5 min 90kW (120bhp) 312Nm 2750 300gm/kW hr

Max. continuous 75kw (100bhp) 275Nm 2575 275gm/kW hr

Economy Cruise 50kW (67bhp) 249Nm 2300 260gm/kW hr

For the propulsor, a ducted fan is proposed to be used as it is more efficient than a normal

propeller[12]

.

4.4.3 Fuel weight

The engine, while running at its economical speed, consumes fuel at the rate of 13 kg/hr.

Using this estimate, and the duration of the total mission as 9 hours, this gives a fuel weight

of 117 kg.

Figure4.3: The WAM-120 engine on display (Source: Geocities)

Page | 32

Table4.6: Margins taken for estimating fuel weight

Margin For Assumption

30% Hovering, Landing and take-off These operations are fuel-intensive

15% Emergency Regulations recommend this margin

2% Unusable fuel Fuel trapped in pipes and tanks

Applying all margins, the total fuel to be carried is 172 kg.

Using density of diesel as 0.832 kg/l,

Volume of diesel required,

4.4.4 Speed Optimization

The cruise velocity of the airship required to be optimized for the response time and specific

fuel consumption. The power required was calculated for different velocities from the drag

data and plotted. The power available from the engine was plotted to obtain the maximum

velocity and the economy velocity. The response time constraint was added to the graph.

Figure4.4: Speed optimization exercise

Page | 33

The maximum achievable velocity of the airship is 93.6 kmph. The most economical velocity

to fly at is 77 kmph. Thus, the range of velocities the airship can fly are from 50 kmph to 93.6

kmph.

A conservative velocity of around 60 kmph is chosen as the cruise velocity. This is lower

than the economical speed, but due to thrust vectoring for fuel compensation, the load on the

engine is expected to be equivalent to the economical power.

4.4.5 Hydrogen fuel cell

Table4.7: Comparison of Hydrogen fuel cells[22]

Fuel Cell Nexa Mark 902 Xcellis HY 205 US STS Shuttle PC-29

Manufacturer Ballard Ballard Ballard UTC IFC/UTC

Fuel Hydrogen Hydrogen Hydrogen Hydrogen Methanol

Type PEM PEM PEM KOH PEM

Max. Output Power (kW) 1.2 8.5 205 12 50

Mass (kg) 13 96 2170 116 141

Table 4.7 shows the comparison of Hydrogen fuel cells currently available in the market. For

the ALFRA-1, the power required is around 50 kW. Multiple PC-29 fuel cells can be used to

power the motors.

The main advantage of a fuel cell is that the loss of fuel weight during flight is negligible and

can easily be compensated by venting out a little lifting gas. The only drawback is storage of

Hydrogen on board in cylinders in large quantities, as Hydrogen storage technologies that are

portable are not developed enough. [23]

One option is to take hydrogen from the inner envelope and use it as a fuel for the on board

fuel cell which can be explored further. This has the potential to solve the inflight buoyancy

control problem.

The main drawback of fuel cells is the prohibitive cost currently. In the future, if the world

has a Hydrogen economy, the cost of fuel cells is bound to come down given the widespread

use. So, it is concluded that Hydrogen remains a viable option for propelling the ALFRA-1 in

the future.

Page | 34

Figure 4.5: 3D model of the proposed ALFRA-1 (rendered using Solidworks)

Page | 35

5. Flow Studies

The airship hull is known to create the maximum drag in an airship. [12]

This is owing to its

large surface area, which increases the skin friction drag and also the large frontal area it

projects, which increases the pressure drag. The National Physical Laboratory’s shape is

chosen as the airship envelope shape due to the possibility of optimization of fineness ratio

for this shape. The airship fineness ratio is a variable parameter for the volume and hence,

CFD simulation is performed using ANSYS FluentTM

to optimize for the least drag.

The Fluent code is validated by using [19]. The results are found in Appendix 4

For the estimated volume, six N.P.L shapes of fineness ratios ranging from 2.5 to 5.0 are

generated. An axisymmetric structured grid is built around the upper half of the N.P.L shape.

Figure 5.1: Axisymmetric structured grid around the upper half of N.P.L shape

Page | 36

Domain independence is verified by increasing the computational domain from 3l to 5l.

The boundary conditions provided are as follows for 600 m altitude cruise conditions:

Flight speed: 20 m/s

Pressure Far field: 0.998 bar

Static Temperature: 298 K

The Reynolds number for the flow is,

⁄

2.89 x 107.

Due to the turbulent nature of the flow, k- turbulence model is used coupled with enhanced

wall function. As we are interested in the drag, the pressure is iterated using the body force

weighted scheme to provide accurate results in the least computation time. The flow field is

thus solved to find the drag for each fineness ratio.

Figure 5.2: Pressure distribution on the upper surface of the envelope

Page | 37

From the drag obtained, the volumetric drag coefficient,

⁄ , is calculated.

In addition to this, the drag coefficient is calculated from the empirical formula [12]

{ (

)

(

)

(

)

}

The simulation and the empirical results are plotted in Figure 5.3

Figure 5.3: Volumetric Drag Coefficient (Empirical and Simulated) for different fineness ratios

The empirically calculated drag is lesser than the drag obtained from simulation results. This

offset can be attributed to the low Reynolds number used to formulate the empirical relation.

From Figure 5.3, it is clear that the drag reaches a minimum at a fineness ratio of 4.5. As the

minimum is a flat one, any fineness ratio from 4.5 to 5 can be selected.

Page | 38

6. Stress Analysis

It is crucial to identify the location of the Centre of Buoyancy (C.B.) and Centre of Gravity

(C.G.) for the airship envelope. The envelope carrying the lifting gas is not only subjected to

buoyant lift but is also loaded under its own weight. Thus, the axial distribution of weight for

the envelope or the linear gravity density and the buoyant lift distribution over the envelope

(axial) were studied.

The linear gravity density, qe is given by [18]

where, ρe= surface density of the envelope material;

r(x)= the radius of envelope varying with the longitudinal x-axis of the airship.

Secondly, the basic load applied on the envelope is buoyancy dependent on the displaced

volume of envelope and the unit buoyancy, the difference between the ambient atmosphere

density and the lift gas density. The linear unit net buoyancy can be calculated from the

theoretical envelope profile [18]

Where, ρn is the unit net buoyancy at specific altitude.

The loads due to buoyancy and gravity were plotted and are shown in figure 7.1.

From the graph, we obtain,

Location of center of gravity = 0.0029m

Location of center of buoyancy = 4.426m

It is seen that the C.B. lies ahead of C.G. which was likely to result in a nose up pitching

moment for the airship. Different structural components like the gondola, stabilizers,

propulsion systems etc. require to be configured and positioned in order to balance this.

As these structural components are added, the weight of these would result in a static bending

moment about the C.G. When there is a certain angle of attack for the airship, due to thrust

imbalance or improper load distribution, etc. the aerodynamic loading of the envelope would

also result in a dynamic bending moment about the C.G. which is dependent on the angle of

attack [18]

.

Page | 39

Figure6.1: Loads due to buoyancy and gravity applied on the envelope

The airship design criteria (ADC) of the Federal Aviation Administration (FAA) provides an

overall formula for the dynamic analysis check of the airship’s flight envelope including the

dynamic effect of the structural component and the aerodynamic effect. FAA maximum

bending moment formula is proposed as:

[ (

) ]

where, d= the maximum diameter of hull (m ) ;

u= gust velocity (m/s);

v= airship equivalent speed (m/s);

Ve= total envelope volume (m3).

The above equation is applicable for 4<l/d < 6. For l/d < 4, assume that l/d = 4. If l/d = 4, the

maximum bending moment formula can be further simplified as

First, the bending moment on the envelope for different values of L/D was computed and

plotted.

It was observed that as L/D ratio increases the bending moment also increases. This data was

used along with the aerodynamic data (obtained from flow studies) and an optimization was

Page | 40

carried out to find out the optimum L/D for the airship. The optimum L/D was found to be

4.5.

Figure6.2: Variation of bending moment on envelope with L/D ratio for a given lifting

capacity

The load capacity of the envelope hull of an airship is limited by the tendency to wrinkle

under high bending moment. To calculate the pressure required for maintenance of envelope

rigidity in flight, the maximum applied bending moment, the sum of static and dynamic

bending moments acting the hull must be established.

The lift gas static head effect (pressure gradient from bottom to top) must also be taken into

account. (This comes into picture prominently when the airship goes into a high angle of

attack orientation, during mooring or while flight).

The required minimum pressure for rigidity to withstand the related bending moment can be

calculated as,

Where, rc = max(r(x));

The longitudinal and hoop forces of airship with radius, r(x) under internal pressure p can be

written as a function of the radii of curvature in hoop direction ρH and in longitudinal

direction ρL.

O

Page | 41

√

|[ ]

|

(

)

f2 and f3 are the loads on the envelope due to the longitudinal and hoop’s stress. These were

computed and their axial distribution over the envelope was plotted.

Figure6.3: distribution of loads due to longitudinal and hoop’s stress on the envelope

Now the maximum value of longitudinal tension=1.49KN/m.

The maximum value of hoop’s tension=2.96KN/m

So the selected envelope material must have a allowable shear strength greater than the

maximum hoop’s tension.

Page | 42

7. Building indoor airship envelopes

Before building the scale model of the ALFRA-1, some hands on experience in building

airship envelopes was required to understand the challenges. To achieve this objective, small

indoor airship models were made.

7.1 Shape of the envelope

Envelope design for other projects was based principally on mathematical shape modelling,

drag estimations and aesthetics. The same NPL shape as mentioned in the preliminary

configuration chapter was decided to be used in this models and the volume of the envelope

was arbitrarily fixed at 5 litres considering ease of manufacture.

Figure 7.1: The NPL Shape

7.2 Materials used

Materials used for airship envelopes are based on helium retention capability of the material,

lower stiffness to aide easier manufacturing etc. The National University of Singapore[21]

and

the Rowan University used polyurethane sheets. The University of Adelaide[20]

and The

Zurich University used Mylar, i.e., BoPET (Biaxially-oriented Polyethylene Terephthalate).

7.3 Fabrication

Mylar (50 microns) was the only workable material we were successful in obtaining and was

used to make small prototypes of the actual model. The aim here was to test different glues

and try to understand how the gores can be joined to get gastight envelopes.

Page | 43

The first envelope made was a four petal sphere, which was almost impossible to join owing

to the low number of petals and stiffness of the material at hand. An attempt to use the

cellotape-heat sealing process also proved to be unsuccessful. Some experts were consulted

online and they suggested that we make a gore tool and increase the number of petals. The

gore tool is basically a support cut to the profile shape on which the gores are kept and glued

together.

Figure 7.2: Gore joining tool diagrams (Source: Alan Sherwood via rcgroups.com/forums)

The gore tool, as shown above was manufactured using a wooden plank. The top surface was

smoothened using a strip of balsa wood. A layer of foam double sided tape was pasted on the

top of the balsa sheet and a removable cello tape layer was pasted on top of that.

Figure 7.3: Manufacturing the gore tool from wood

The second envelope was made with 12 gores using the gore tool. Fevi-Bond, a popular

rubber based adhesive was used and a closed ellipsoid was successfully made. However, the

Page | 44

glue began to crystallize and became stiff after a few days and there were visible holes at the

seams.

Figure 7.4: Joining the mylar gores on the gore tool

A third envelope was then made using thin double sided tape. It had six gores. Cello tape was

used to seal the seams so that small leakages at creases could be prevented. A valve from an

inflatable toy was used and the balloon was inflated using compressed air. However leakages

were found at the sections where the valve was joined with the envelope. Thus, it was

concluded that the double sided tape and cello tape did give an airtight envelope and it was

the way the valve was joined that was faulty.

UHU power glue, an adhesive tested and recommended for joining Mylar on various forums

on the internet was procured. The next envelope that was made used a combination of double

sided tape and UHU power glue and gave reasonable gas tightness.

Although these models did not fly and were not capable of lifting more than 5 grams, they

provided the much needed knowledge, which aided in addressing the problems that cropped

up during the fabrication of the scale model of ALFRA-1.

Page | 45

8. Preliminary design of the ALFRA-1 scale model

8.1 Weight Estimation

Airship components that had been purchased were weighed and the rest were estimated. The

table below shows the weights of the components and the quantity that will be used in the

airship. Weights of components from serial number 10 to 13 (in bold below) are estimated

while the rest are weighed on a digital scale.

Table8.1: Estimated weights

8.2 Envelope

The design of the envelope was kept similar to the one of ALFRA-1. It has an inner hydrogen

chamber surrounded by a multi-chamber outer envelope containing helium. From the weight

estimation carried out earlier it is clear that the amount of disposable lift required out of the

envelope is 1100gm.

Sl. No. Component Quantity Total Weight (g)

1 Brushless Motors

2 122

2 ESC

2 40

3 LiPo Battery

1 45

4 Receiver

1 09

5 Servo

1 35

6 Board

1 50

7 Battery

1 35

8 Ultrasonic Sensor

2 40

9 Camera + Battery

1 125

10 Tow Line

1 100

11 Stabilizers

4 100

12 Velcro

N/A 100

13 Gondola

1 300

Total 1.1 Kg

Page | 46

A MATLAB code (refer appendix 3) was developed to iterate and compute the dimensions

and coordinates of the envelope with enough volume to generate the required lift. For this,

the material density and thickness of the material were used as specified by National Balloon

Facility.

The statistical analysis carried out in the literature survey (refer appendix 1) had shown that

the length to diameter ratio of airships show a reducing trend with time. In earlier days,

before WWII, large airships with length to width ration of 6 or more were built. But recent

airships are small with a length to width ratio of around three. This is because materials are

now available which can withstand the stresses at such low length to diameter ratios. The

University of Adelaide went for a length to width ratio of 3.7 and the National University of

Singapore went with a value of 2.266.

As the airship was fabricated as a scaled prototype of the bigger ALFRA-1, the length to

diameter ratio for both airships was to be kept same. However the optimization of length to

diameter ratio for minimum drag and structural loading was not completed by the time the

preliminary design for the prototype began and thus a fineness ratio of 3, which is on a

slightly higher side to what was obtained from the statistical analysis, was chosen.

The outer envelope dimensions were obtained in this way but for the inner envelope the

volume of hydrogen to be carried was required to be known.

Several different criteria were studied to see if there can be some constraint over the volume

of hydrogen. Increasing the size of inner envelope (which means increasing the percentage

volume of hydrogen) meant more lift and lesser cost. There was however a weight penalty as

the overall weight increased owing to an increased size of the inner envelope. But it was seen

that the increase in lift far outweighed the increase in overall weight. Moreover, the only

purpose of using helium or having a helium envelope over the hydrogen is to prevent the

interaction of hydrogen and air in case of leakage in the hydrogen envelope. This would be

catered even by a thin layer of helium used as a jacket over the hydrogen envelope.

As no condition could be imposed to constraint the size of the hydrogen envelope, the inner

envelope was chosen such that it aides the manufacturing as well as makes it more visually

distinct and demonstrates the two envelope design better.

Page | 47

Table8.2: Envelope dimensions

Length(m) Diameter (m)

Outer envelope 3.86 1.28

Inner envelope 3 1

8.3 Lifting Gas

The envelope was designed to hold hydrogen in the inner chamber and helium in the outer

chamber. The lifting capacity of both gases are tabulated below

Table8.3: Comparison of lifting gas properties

Gas Density

(Kg/m3)

Lifting capacity

(Kg of Lift per m3 of gas)

Hydrogen 0.089 1.191

Helium 0.178 1.102

The above values show that hydrogen has a higher lifting capacity than helium by 7.56%.

Although hydrogen being inexpensive, it can be dangerous as it forms an inflammable

mixture with oxygen when allowed to mix with air. Helium on the other hand has a slightly

lower lifting capacity and is expensive. However it’s an inert gas and comes with the

advantage of being safe, particularly for indoor handling conditions.

Considering safety policies in the campus, it was decided to use helium in both the outer and

inner envelopes for the flight tests.

8.4 Propulsion system

All the small airships studied during the literature survey used electric motors to drive the

propellers/ ducted fans. Lithium Polymer batteries are lightweight, small and efficient and

hence are ideal for use in an airships.

Thus, a few parameters of flight were fixed so that power estimation could be made for

choosing the motors for the propulsion system. The maximum take-off weight of the airship

is approximately 3 kg. The cruise velocity was selected as 1m/s and the time required for

Page | 48

attaining the cruise velocity was taken to be 10 sec all of which are in accordance with the

historical data was obtained from the literature survey. The drag at cruise is calculated by

using a CDV for the entire airship of 0.06[11]

.

Table8.4: Scale model parameters

Parameter Value

Weight 3kg

Cruise velocity 1m/s

Time to reach the cruise velocity 10sec

Drag at cruise 0.07 N

Max thrust required 60gms

The maximum thrust required is calculated for the condition of hover and take-off. As two

engines are being used the thrust required from each engine should be equal to 30gms which

is half the total thrust estimated using historical data. A thrust rig was designed in order to

test the thrust produced by the motor for different propeller speeds (in terms of percentage of

maximum thrust applied). The construction of the set-up is shown in the schematic diagram

below.

Figure8.1: Schematic and picture of the thrust measurement rig built in the lab

The rig has two arms of equal length joined perpendicular to one another (AB and BC in the

figure). The motor and the propeller were mounted to AB as shown and the end C is rested on

a digital weighing balance. The motor was connected to the battery through an ESC

(Electronic Speed Controller) and the transmitter-receiver set was attached. The thrust was

gradually increased and the weighing machine readings were taken at different thrust settings

Page | 49

(position of the control stick). “T” is the thrust generated by the engine and “N” is the

reaction by the weighing machine.

Applying moment balance about B:

The data was tabulated and the analysis showed at what percentage of maximum thrust the

motors needed to be run

Table8.5: Results obtained from the thrust measurement rig

% of full thrust applied Thrust measured (grams)

0 0

16.67 35

33.34 107

50 176

66.67 265

83.34 350

100 485

Figure8.2: Operating point of the motors for cruise

Page | 50

Thrust required from a single motor is 30gms. As the motor was capable of producing thrust

as high as 485grams which is larger compared to our zone of operation, it was decided to

have the maximum thrust set at 60grams for each motor (which is twice the designed thrust as

calculated above). A graph was plotted for the tabulated data using a soothing-spline to fit the

data. This showed that the maximum thrust setting for each motor is 25%.

8.5 Gondola

The components contained in the Gondola are:

Payload (camera)

Engine (motors + propellers)

Electronic Speed Control

Battery

Receiver

Servo for thrust vectoring

The frame of the Gondola was made of Balsa wood to have a minimum weight. The two

motors were mounted to their respective fire-walls that was connected to a plate (made of

balsa and ply) hinged to the under-side of the gondola body. This assembly was connected to

a servo that would rotate it about the hinge, thus allowing thrust vectoring for climb and

descent.

The gondola was designed such that there is enough clearance between the rotating propellers

and the envelope. This governed the height of the gondola. The size of the gondola was kept

small enough to reduce the net weight but sufficiently large enough to house all the

components. A trapezium shape instead of a regular rectangle was adopted because having a

larger area at top meant that when suspended the load of the gondola is distributed over a

larger area of the envelope.

Page | 51

Figure8.3: Gondola drawings –side and front views

8.6 Empennage design

Stabilizers have a very small frontal area compared to their longitudinal area. The larger

longitudinal area produces more resistance to motion in the yaw, roll and pitch planes.

Stabilizers are needed on the tail of the envelope to add stability to the airship. The airship’s

small weight, relative to its surface area, means that it was susceptible to the effects of air

currents. A four-stabilizer, plus configuration was chosen for the airship replicating the tail

design for the flood relief airship. However, the stabilizers were not to be provided with any

control surfaces because at the low speeds of airship operation, the effectiveness of the

control surface would be very less. The fins are fabricated as a balsa wood frame covered

with foil paper.

Table8.6: Comparison of empennages of indoor airships[20]

Airship Envelope

Longitudinal Profile

Area (m2)

Stabiliser

Longitudinal Area

(m2)

Area Ratio

Airship Solutions 4m3

3.32 0.18 18.84

Airship Solutions 11m3

6.88 0.34 20.48

NUS Airship 0.55m3

1.01 0.03 33.50

Average 3.74 0.18 24

Page | 52

Figure8.5: Stabilizer dimensions

The literature survey revealed that, the University of Adelaide[20]

had conducted a survey on

small airships (mostly indoor) and had studied the ratio of longitudinal profile area to the

stabilizer longitudinal area.

The above data was used to compute the stabilizer area.

The envelope longitudinal profile area=4.25m2.

Average area ratio from the above mentioned study=24

So, stabilizer longitudinal area=0.175m2.

The dimensions of the stabilizers were selected in

proportion to that for the flood relief airship.

The front taper angle was chosen as 370, a value

derived from historical data.

8.7 Free Body Diagram of Airship

The centre of gravity and centre of buoyancy for the envelope can be obtained using the

distribution of gravity and buoyancy load on the envelope.

Centre of gravity is 10.56 cm behind the centre point (point of maximum diameter) of the

envelope and the centre of buoyancy is 9.9 cm behind the centre point. For the envelope

alone, the net buoyant force acts on the centre of buoyancy and the weight acts on the centre

of gravity, thus resulting in a pitch up moment. The fins and the gondola are installed on to

the envelope with proper positioning so that trim condition is attained. The line of thrust is

below the centre of gravity of the system due to which a net pitch up moment tends to be

created when the engines produced thrust.

A force analysis was carried out in order to find the proper positioning of the different

components of the airship.

Page | 53

Figure8.6: Forces acting on the airship

The above figure shows the forces acting on the airship.

G=centre of gravity for the envelope where the weight of the envelope acts.

B=centre of buoyancy where the net lift force acts

F= the position of the fins and the weight of the fins act here

T= position of the Gondola. The thrust from the engines and the gondola weight act here

The position of the Gondola was fixed in the y-direction due to its design dimensions. Our

aim here was to find the position of the gondola axially so that trim could be achieved.

Table 8.7: Position and forces on different components

component Position Forces

Envelope C.G.= -10.56cm

C.B.= -9.9cm

Weight at C.G.= 2064gm

Buoyancy at C.B = 3145gm

Fins C.G= -126cm Weight = 75gm

Gondola C.G.:

Position along y-axis=65 cm from axis

of envelope

position along x-axis= to be obtained

Weight = 800gm

Thrust (adjustable)

Page | 54

So, applying moment balance about the z-axis: the position of the Gondola in axial direction

=3.675cm ahead of the centre point (for zero angle of attack at no thrust condition).

However, due to the thrust applied, a net pitch up moment will develop. A study was carried

to find out what should be the position of the gondola such that for a given thrust value the

airship will be in trim condition. The graph shown below was plotted between the position of

the Gondola and thrust required. So, for a maximum thrust required of 60gm as estimated

earlier, the position of the gondola was fixed at 8cm ahead of the centre. This ensured that at

cruise the flight path angle was zero and thus minimum aerodynamic loading.

Figure 8.7: thrust required vs axial position of gondola ahead of centre

Page | 55

9. Fabrication of the scale model

9.1 Envelope

The coordinates of the gores for the outer and inner envelopes were obtained using the

MATLAB code and so were the coordinates of the chamber walls. A solid works model was

then built to verify the coordinates and visualize the envelope. Once the coordinates were

frozen, the fabrication process started.

Figure 9.1: Gore profiles of the inner and outer envelope

The fabrication of the envelope was undertaken at the National Balloon Facility, TIFR,

Hyderabad. Tables extending more than 180m in length, heat sealing machines for gores,

hand-held heat sealing machines for sealing more intricate seams, tapes and powders for use

with polyethylene and gore cutting machine for larger gores were some of the facilities

available there.

9.2 Heat sealing machines

Different kinds of heat sealing machines were used at NBF. The main heat sealing machine

(depicted in figure 9.1). was used to seal long seams with lesser curvature. It had two heating

elements and two cooling elements in succession and rollers were present at both ends to

facilitate the positioning and movement of the seam between the elements. The feed rate

could be adjusted in order to vary the time of heating so that different thickness of material

could be handled.

Page | 56

Figure 9.2: Main Heat sealing machine

Figure 9.2 shows the hand-held heat sealing machine which was similar to a hand held

clothes iron. It had a narrow heating element that had to be pressed on to the seam that was

supposed to be sealed. The other hand held machine was similar in structure to a soldering

iron with a roller at its tip. It was used for sealing seams with high curvature particularly

those encountered during sealing of the nose, tail and filling tubes.

Figure 9.3: Hand held heat sealing machines

The fabrication began by marking the coordinates of the gores and the walls on polyethylene

sheet of desired thickness. As the outer envelope required good weathering, load carrying and

gas retention properties, it was fabricated using 75 micron thick sheets. For the inner

envelope where gas retention was the only major concern, 38 microns thick polyethylene was

used.

A sealing width of 2cm was added to the designed coordinates and the marking was done.

This excess was provided for sealing. After marking, the gores and the walls were cut out.

Page | 57

However the work came to a halt when the gores and the chamber walls were to be sealed.

Sealing is straightforward when the contours of the sheets which are being sealed are same.

The procedure involves overlapping the sheets and sealing over the seam width provided. But

in this case as the sealing was supposed to be carried out over a 3D curve (the curved

envelope surface sealed to the straight chamber walls which are perpendicular to the

envelope), the sealing was not easy and pinches were needed to be introduced at the seams

and then sealed. Introduction of pinches to the outer envelope however meant that the shape

of the envelope was not maintained and this can lead to increase in drag and a loss in terms of

aesthetics. A solution to the problem was to first use a sheet larger than the designed wall, cut

to the profile of the outer envelope at one end. So this end could be sealed to the outer

envelope first without pinches appearing on the outer envelope. Implementing this idea, the

inner envelope was placed over extra outer envelope gores and the shape cut out making sure

the wall width was maintained. After the outer envelope gore was sealed to this, the inner

gores were sealed at the other end, introducing pinches near the nose section where the

curvature was high.

Figure 9.4: The outer envelope and a close look at the seam

In the preliminary design phase, design was to have three chambers in the helium envelope.

But, due to the difficulties faced in manufacturing and the shortage of time, it was decided to

go ahead with a two-chamber outer envelope configuration during the fabrication stage. This

ensured simplicity of sealing at nose and tail and sealing the walls at the junctions near the

ends.

For filling gas into the chambers, filling tubes had to be attached. Filling tubes were first

fabricated using thin strips of polyethylene cut to a specific profile and sealed in the form of a

tube. These were then hand sealed to the chambers. The filling tube for the inner hydrogen

Page | 58

chamber was passes through the filling tubes of one of the outer chambers. This was to

facilitate the inner chamber to be filled first allowing the inner filling tube to be closed (tying

with rope) and then the outer chambers filled one after the other.

After the sealing of the gores was done, the nose and the tail portions were sealed. This was

done by using a circular sheet for the nose sealed to the open end using the hand-held heat

sealing device. The tail sealing, due its lower curvature was straight-forward.

After the sealing was complete the air tightness and the retention of the original design shape

was to be tested. This was carried out using compressed air. The envelope was found to be

airtight.

Next, patches were fixed on to the envelope using pressure tapes. The pressure tapes are

designed to carry large loads and are adhesive to polyethylene. Specified lengths were cut out

and used to distribute the loads over a large area. The purpose of these patches was for tying

ropes that would be used for holding the fins in position, suspending the gondola and for

tying the handling ropes.

After the entire process was complete, the first test was done using hydrogen as the lifting

gas. Though originally designed to carry a weight of 1100gm with helium as the lifting gas,

the envelope was neutrally buoyant only when a weight of 900gm was suspended. This

reduction in lift was later attributed to the impurity of the hydrogen used.

9.3 Stabilizers

The stabilizer dimensions were already known from the preliminary design phase. In order to

reduce the weight it was decided that a frame would be made out of balsa wood and then foil

paper would be pasted and pressed (using iron box) to make the fins.

The frame was made out of 25mm wide pieces of the specific lengths. They were joined

using cyanoacrylate based glue. The bottom piece was wider so that there was sufficient

room to put the tapes while fixing the fins on to the surface of the envelope.

Page | 59

Figure 9.5: Fin frame dimensions in mm and photograph

Foil paper was used to form the skin of the fins by pasting it over the frame of the fins using

Fevi-bondTM

. The fins were arrested in their position by using thin nylon ropes. On the

envelope small tubes were attached using pressure tapes that formed a load patch. A load

patch is a certain arrangement of pasting tapes that can carry large loads and is generally used

for suspending the gondola or tethering ropes. A small attachment was provided on the fin so

that the ropes can be passed and the fin can be tethered.

According to the preliminary design, the plan was to have a plus configuration for the

stabilizers so that there is simplicity in control. However, when the load patches were fixed

on to the envelope, there was an error in the positioning of the nose-end load patches.

However, as the fins were designed simply to give stability in pitch and roll and were not

designed to have any control surface. This helped us to change the design and opt for a cross

configuration of stabilizers so that no modifications needed to be made that would have

further delayed the fabrication.

Figure 9.6: Front view schematic of the scale model

Page | 60

The net weight of all the fins along with the ropes used to tether them was measured as

90gms which is close to our initial estimate of 100gms.

9.4 Gondola

The gondola was fabricated using balsa wood. The frame of the casing was made from 6mm

balsa sheet. The design was made such that it reduced the overall weight and still imparted

enough strength to take the loads. The tilting mechanism was made using a laminated sheet of

balsa and plywood, so that it could carry the load of the motors and propellers fixed to it and

would be rigid enough for the components to be screwed on to it. The firewall instead of

being suspended from this arm was incorporated in the same. A stainless steel door hinge was

used to connect this piece with the main casing of the gondola. The motors were mounted to

the extreme ends of the arm and the servo was fixed so as to attain free movement about the

hinge axis and have the maximum possible throw.

The whole structure was then covered with foil paper using fevi-bond on all sides except the

back, so as to have access to the components in case some modifications were desired. The

finished gondola is shown in Figure 9.7.

Figure9.7: Fabricated gondola photographs

The weight of the Gondola without its components was measured as 105gms, which is

195gms less than our preliminary estimate. This reduction in weight was deliberately

introduced by use of balsa wood and using a light weight casing (designed as a balsa wood

frame covered with foil paper) so that the increase in weight of the envelope could be

compensated for to some extent.

Page | 61

9.5 Revised weight estimation

After the fabrication of the gondola and the fins was carried out, the weights of all

components were taken again to have a final weight estimate. The measured values are given

in the table 9.1.

Table 9.1: Comparison of estimated and final weights of components used in flight

Sl. No. Component Quantity Estimated

weight(g)

Final Weight (g)

1 Brushless Motors

2 122 70

2 ESC

2 40 46

3 LiPo Battery

2 45 183

4 Receiver

1 09 09

5 Servo

1 35 35

6 Camera + Battery

1 125 105

7

Propellers + collets 2 N/A 33

8 Stabilizers

4 100 100

9 Gondola

1 300 257

10 Velcro

N/A 100 N/A

Tow line

1 100 N/A

Total 1100 838

Page | 62

10. Flight Testing of Scale Model

10.1 Flight Test

Two flight tests were conducted using the scale model. The first was to test the stability as

well as the handling characteristics of the airship. Since, an airship was never flown by

anyone in the team, a practice run was required to obtain experience in flying an airship. In

the first flight test, the yaw control was found to be ineffective. This was attributed to the

small moment arm offered by the motor in the yaw direction. It was concluded that the

moment arm had to be increased before the next flight test. The airship was also subjected to

wind loads and was found to be stable in roll, pitch and yaw.

The moment arm was increased from 185mm to 305 mm before the second flight test.

The second flight test of a duration of 87 minutes demonstrated better controllability and

thus, a total untethered flight was made possible.

Figure 10.1: Scale model in flight

Page | 63

10.2 Flight parameters

10.2.1 Turn Rate:

The turn rate was calculated by measuring the time taken to turn 180° starting from a static

position. The turn rate was achieved by giving full thrust to outer engine and switching off

the inner engine. Three readings were taken and the average is presented in the table 10.1.

Table 10.1: Turn rates for different directions

Direction Time Taken (seconds) Turn Rate (deg/sec)

Clockwise 15 12

Counter-clockwise 18 10

The turn rate was observed to be different for the two different directions because the

maximum thrust obtained from the two motors were different due to separate batteries being

used to power the two separate motors.

10.2.2 Speed:

The speed data obtained from the GPS device was collected and plotted to find the maximum

speed obtained during the test flight. The maximum speed, as shown in figure 10.2, was

found to be 16.7km/h (4.64 m/s).

Due to the lack of available open space to carry out the flight test, the tests could only be

carried out at a maximum throttle of 50%. The throttle had to be cut regularly and full throttle

could not be given for a sufficient period of time. So, the maximum speed that the airship can

attain is expected to be a higher value.

Page | 64

Figure10.2: GPS speed data

10.2.3 Rate of climb :

The GPS altitude data for the segment of test was collected and plotted as shown in the

altitude vs. time graph (figure 10.3). The values of maximum and minimum altitude attained

by the airship during the flight test were obtained from the data.

Table 10.2: Maximum and minimum altitude attained

Maximum altitude 122m (above MSL)

Minimum altitude 91m (above MSL)

Page | 65

Figure10.3: GPS altitude data

The climb rate was calculated by bringing the airship to the lowest altitude possible and then

providing an upward thrust. The altitude difference and time difference is obtained from GPS

data (Figure 10.3). The slope gives the ROC as 0.32 m/s. This ROC was obtained at 50% of

maximum thrust.

Page | 66

11. Recommendations

Preliminary design of the ALFRA-1 is complete. A scale model of the ALFRA-1, the first

airship to be built and the flight tested in IIST, has demonstrated the proof of concept.

Future scope:

The scope of future work to be done in this project is large. Some of the recommendations are

listed below:

1) The detailed design of ALFRA’s subsystems needs be carried out.

2) The scale model can be modified for use as a payload carrier for many experiments at

IIST, which require a flying platform.

3) Autonomy can be implemented in the scale model to check the effectiveness of

control algorithms.

4) A larger scale model can be fabricated and flight tested.

5) A complete aerodynamic analysis, including wind tunnel testing, if possible can be

carried out.

6) A detailed 6 DOF analysis needs to be carried out to verify the adequacy of the

control effectiveness.

Page | 67

Appendix 1

The length to width ratio of airships shows a reducing trend with time. In earlier days, before

WWII, large airships with length to width ration of 6 or more were built. But recent airships

are small with a length to width ratio of around three. The University of Adelaide went for a

length to width ratio of 3.7 and the National University of Singapore went with a value of

2.266.

Figure A1.1: Length to weight ratio of airships with We< 1 tonne [20]

The variation of the Thrust to weight ratio of airship is over a wide range, i.e., within 0.1 and

0.7. A study done by the University of Adelaide regarding the Thrust to weight ratio of small

airships shows that the values are close to 0.3.

The main reason behind such huge variations is that different values of thrust to weight ratio

cater to different purposes the airships are designed in the first place. An airship designed to

cruise at a high velocity will have a large Thrust to weight ratio than designed to Hover.

Over the years, however the empty weight to take-off weight ratio of the airships has stayed

at a relatively constant value of 0.6.

Page | 68

Appendix 2

Figure A2.1: Orange areas show the flood prone areas in India. The legend gives the

elevation of those areas. (Composite map created using data from Wikipedia.)

Page | 69

Appendix 3

Code to generate envelope shape, volume and weight of the envelope:

th=75*10^(-6);%thickness of envelope material and payload weight can be

changed here

dena=1290; %density of air

denhe=178.6; %density of helium

denh=89; %density of hydrogen

weight_payload=1100;

ld=3;%

ab=ld*2/2.414;

vol=weight_payload/1000; %initial volume of envelope based on empirical

formula

%iteration to get the final value of weight and volume

for j=1:1:100

b=((vol/(4.46389*ab))^(1/3));

a=ab*b;

i=1;

X=0;

Y=0;

for x1=-sqrt(2)*a:0.001:0

X(i)=x1;

Y(i)=(b-b*x1^2/(2*a^2));

wt_env(i)=-2*pi*Y(i)*0.92*70;

lift_env(i)=(dena-denhe)*pi*Y(i)^2;

i=i+1;

end

for x=0.001:0.001:a

X(i)=x;

Y(i)=sqrt(b^2-b^2*x^2/a^2);

wt_env(i)=-2*pi*Y(i)*0.92*70;

lift_env(i)=(dena-denhe)*pi*Y(i)^2;

i=i+1;

end

[X Y]; %envelope contour

length=2.414*a;

surface_area=pi*b*(b+(a^2/sqrt(a^2-b^2))*asin(sqrt(a^2-b^2)/a))+

1.481*a*b*(1+3.0144*sqrt(1+0.493*b^2/a^2)); %surface area of the envelope

weight_env=2*surface_area*th*0.92; %weight of envelope

Page | 70

weight_gross=weight_payload+weight_env+vol*178.6;

vol_new=weight_gross*10^(-3)*9.81/10.35;

if (abs(vol-vol_new)<0.001)

break

else

vol=vol_new;

end

end

weight_helium=vol_new*178.6;

cg=0;

cb=0;

for j=2:1:numel(X) %computing position of C.G. and C.B.

cg=cg+X(j)*(-1)*wt_env(j)*0.001/(weight_env);

cb=cb+X(j)*lift_env(j)*0.001/((dena-denhe)*vol);

end

Code to generate gore contour:

a=1.6; %1.24

b=0.64; %0.5

i=1;

X=0;

Y=0;

for x1=-sqrt(2)*a:0.05:0

X(i)=x1;

Y(i)=b*(1-x1^2/2/a^2);

i=i+1;

end

for x=0.01:0.05:a

X(i)=x;

Y(i)=sqrt(b^2-b^2*x^2/a^2);

i=i+1;

end

[X Y];

arc_len1=quad(@(t)coordbig_parabola(t,a,b),-sqrt(2)*a,0);

arc_len2=quad(@(t)coordsmall(t,a,b),0,a);

no=12; %no. of petals

clear i j

Page | 71

for j=1:length(X)

if X(j)<=0

m(j)=quad(@(t)coordbig_parabola(t,a,b),-sqrt(2)*a,X(j)); %petal

x-coordinate

n(j)=pi*Y(j)/no; %petal y-coordinate

else

m(j)=quad(@(t)coordsmall(t,a,b),0,X(j))+arc_len1;

n(j)=pi*Y(j)/no;

end

end

plot(m,n)

daspect([1 1 1]) %set data aspect ratio to 1:1:1

function f=coordbig_parabola(t,a1,b1)

f=sqrt(1+b1.^2.*t.^4/a1.^4);

function f=coordsmall(t,a1,b1)

f=sqrt(1+(b1.^2./a1.^2).*(t.^2./(a1.^2-t.^2)));

Code for shape, volume, weight estimation and stress analysis of the ALFRA-1:

weight_payload=2000000;

dena=1290; %density of air

denhe=178.6; %density of helium

denh=89; %density of hydrogen

ld=4.5; %fineness ratio

ab=ld*2/2.414;

vol=weight_payload/1000; %initial volume of envelope

%iteration to get the final value of weight and volume

for j=1:1:100

b=((vol/(4.46389*ab))^(1/3));

a=ab*b;

i=1;

X=0;

Y=0;

for x1=-sqrt(2)*a:1:0

X(i)=x1;

Page | 72

Y(i)=(b-b*x1^2/(2*a^2));

wt_env(i)=-2*pi*Y(i)*0.35;

lift_env(i)=(dena-denhe)*pi*Y(i)^2;

i=i+1;

end

for x=0.1:0.1:a

X(i)=x;

Y(i)=sqrt(b^2-b^2*x^2/a^2);

wt_env(i)=-2*pi*Y(i)*0.35;

lift_env(i)=(dena-denhe)*pi*Y(i)^2;

i=i+1;

end

[X Y];

length=2.414*a;

surface_area=pi*b*(b+(a^2/sqrt(a^2-b^2))*asin(sqrt(a^2-

b^2)/a))+ 1.481*a*b*(1+3.0144*sqrt(1+0.493*b^2/a^2));

weight_env=2*surface_area*0.35;

weight_suspension=21000*vol/1000;

weight_nose_reinforcements=21000*vol/1000;

weight_subsystems=2390000;

% weight of subsystems include th e

following

%weight_container=550000;

%weight_watertank=120000;

%weight_pumps_pipes=20000;

%weight_transmission=100000;

%weight_props=200000;

%weight_fuel=650000;

%weight_engines=450000;

%weight_structuralframe=300000;

longitudinal_profile_area=pi*a*b/2+4*sqrt(2)*a*b/3;

weight_fin=4*longitudinal_profile_area/24*6000;

Page | 73

%weight of fin estimated from empirical

formula

weight_gross=1.2*(weight_payload+weight_env+weight_fin+weight_

subsystems+weight_suspension+weight_nose_reinforcements+vol*17

8.6);

vol_new=weight_gross*10^(-3)*9.81/10.35;

if (abs(vol-vol_new)<0.001)

break

else

vol=vol_new;

end

end

weight_helium=vol_new*178.6;

Mfaa=(0.0292*dena*20*20*vol*length^0.25)/1000;

plot(X,wt_env);hold on;plot(X,lift_env);

del_p=2*Mfaa/pi/(max(Y))^3;

x_cg=0;

x_cb=0;

for t=1:1:numel(X)

x_cg=x_cg+X(t)*-

1*wt_env(t)*0.1/(weight_env+weight_helium);

x_cb=x_cb+X(t)*lift_env(t)*0.1/((dena-denhe)*vol);

if X(t)<=0

Y1(t)=-1*b*X(t)/a^2;

Y2(t)=-1*b/a^2;

phiH(t)=Y(t)*sqrt(1+(Y1(t))^2);

phiL(t)=Y(t)*abs(((1+(Y1(t))^2)^(3/2))/Y2(t));

fH(t)=del_p*phiH(t)*(1-phiH(t)/2/phiL(t));

fL(t)=del_p*phiH(t)/2;

else

Y1(t)=-1*b*X(t)/(a^2)/sqrt(1-(X(t))^2/a^2);

Y2(t)=(b^2/(a^2-(X(t))^2))*((X(t)/b^2)-sqrt(1-

(X(t))^2/a^2)/b);

phiH(t)=Y(t)*sqrt(1+(Y1(t))^2);

Page | 74

phiL(t)=Y(t)*abs(((1+(Y1(t))^2)^(3/2))/Y2(t));

fH(t)=del_p*phiH(t)*(1-phiH(t)/2/phiL(t));

fL(t)=del_p*phiH(t)/2;

end

end

Page | 75

Appendix 4

Using a 6:1 prolate spheroid [19]

, the CFD code for axisymmetric bodies has been validated.

The solver used was ANSYS FluentTM

. The flow conditions are as given below.

Speed: 12 m/s

Nominal Unit Reynolds number: 8 x 105 m

-1

Figure 6.1: Model and the measuring probe

Page | 76

Figure 6.2: Pressure coefficient and skin friction coefficient vs. non dimensional length

The inaccuracies in the skin friction coefficient are due to measurement inaccuracies as it is

difficult to get measurements near the tail of the body [19]

.

Page | 77

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