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HUMANUM EST INFINITA POTESTAS…

-The power of human imagination is limitless….

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Table of Contents ABSTRACT: .......................................................................................................................................................... 3

1.LITERATURE REVIEW ............................................................................................................................... 4

1.1.INTRODUCTION ........................................................................................................................................ 4

1.2.MISSION SCIENCE OBJECTIVES: ................................................................................................. 6

1.3.LOCATION OF THE MISSION: .......................................................................................................... 6

1.4.ENVIRONMENT: ........................................................................................................................................ 8

1.5.NASA ARES MISSION: ........................................................................................................................ 13

2. CALCULATIONS ........................................................................................................................................ 16

2.1. Considerations: ........................................................................................................................................ 16

2.2. Initial Thrust requirement: .................................................................................................................. 17

2.3.Rotor Blade Calculations: .................................................................................................................... 21

2.4. ACTUATOR DISC THEORY-AUTOGIRO ................................................................................. 23

2.5. BLADE ELEMENT THEORY ........................................................................................................... 26

2.6.CYCLIC PITCH ANGLE ........................................................................................................................ 29

2.7.THRUST PER BLADE USING BLADE ELEMENT THEORY ............................................. 33

2.8. MEAN ROLLING MOMENT ......................................................................................................... 35

2.9. AUTOROTATION AND MINIMUM VELOCITY CALCULATIONS ................................. 38

2.10 INDUCED POWER FACTOR , FIGURE OF MERIT, LIFT & DRAG

COEFFICIENT .................................................................................................................................................. 42

2.11 GLIDE PHASE CALCULATIONS ................................................................................................. 43

3. DESIGN TRADE-OFF .............................................................................................................................. 46

4.PROPULSION SYSTEM .......................................................................................................................... 49

5.FINAL DESIGN GENERATION ............................................................................................................ 52

6. STOWING CONFIGURATION ............................................................................................................. 69

7. ONBOARD SCIENCE INSTRUMENTS ............................................................................................ 72

7.1. DUST DETECTOR .................................................................................................................................. 72

7.2.SPECTROMETER ..................................................................................................................................... 73

7.3. RADIATION ASSESSMENT DETECTOR: .................................................................................. 74

7.4. 3-D TERRAIN MODELLING ............................................................................................................. 75

7.5. CAMERA ........................................................................................................................................................ 76

CONCLUSION ........................................................................................................................................................... 77

Word Count: 9487

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ABSTRACT:

The Mars Autogyric Neutroscopic Explorer mission is a concept of a rotorcraft

UAV named X-I for a Mars Mission. The primary mission objective is to look at

the possibility of finding water on Mars using a Neutron Spectrometer. A UAV

design was chosen as an alternative because of its ability to perform survey

over a larger area in comparison with a rover. During the project, a design

optimisation study was done to create a concept which incorporates both

autogiro as well as glider configurations of flight to increase mission lifetime and

control stability. Calculations for the design parameters were done based on a

combined theory of Powered lift and forward flight of an aircraft. This was

followed by a feasibility study of such an option while undertaking a

performance analysis of such a mission. The significant challenges associated

with such a project is getting an aeroplane to Mars and flying through a thin C

in Maritian atmosphere; defining a vehicular geometry, aerodynamics and

mission constraints around the fairing as well as atmospheric parameters. The

design of X-I tries to answer such questions through this project.

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1.LITERATURE REVIEW

1.1.INTRODUCTION

Mars during recent years has been a great field of study for researchers. Since

the first close-up pictures of Mars were taken in 1965, spacecraft voyages to

the Red Planet have revealed a world strangely familiar to our predetermined

perception about its origin. Every time we feel close to understanding Mars;

new discoveries send us straight back to the drawing board to revise existing

theories challenging our views about the planet.[1]

Wernher von Braun’s book ‘The Mars Project’ was the first instance when an

option of landing gliders was discussed as a transfer option from Maritain Orbit

to the surface. The option of having an unmanned UAV as an observation

desk offers the best alternative for a science platform while reducing the risk of

sending a manned mission or a slower option of rover. The drawback for using

rover as an observational platform is that it minimises the surveying area. With

a planet of the size of Mars, rover missions allow us the possibility of having

limited scope of observing hence better alternatives for remote sensing were

required.

Manoeuvrability plays a vital role in defining mission objectives as observation

requirements may change over varied circumstances(atmospheric flux, new

objectives) hence a weather balloon looks like an unviable option. Factors like

weight of instruments and the craft mass itself, aerodynamic design, low

Reynolds number, high flight Mach number, low atmospheric density are the

major design constraints.

Propulsion for such an aircraft is too considered as a main design driver as

atmospheric density in Mars provides us less mass for momentum transfer.

Hence a lower thrust is generated for a given propulsion system when its

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performance is calculated to that on Earth. Lack of Oxygen is too an issue for

an air breathing engine as it leads to added mass on the craft. The propulsive

thrust generated by electric driven propellers causes added weight of batteries.

Hence the concept of using rocket engines for propulsion is the most viable

option.

Apart from these, the challenge to get an airplane to Mars would be a

significant design driver. The geometric arrangement best suited for stable

atmospheric flight is much different from that best suited for launch and

atmospheric re-entry. Efficient packaging of the aircraft is critical to provide

sufficient wing area (lift capability) within the geometric constraints of the

launch and entry vehicles. Deployment of the stowed design into flight

configuration is a stiff challenge. A mid-air deployment strategy provides the

challenge of transition from ‘falling’ to flying. This mid-air conversion from the

stowed configuration to flight, in which the airplane must take the final shape,

orient itself, and execute a pull-out manoeuvre, is a critical design point.[2]

After this conversion, the flight needs to glide down until it achieves the desired

velocity. Thereafter, the deployment of the rotor wings of the autogiro takes

place which requires for added stress analysis on the wings as they act both

as part of aerofoil in the initial flight and later as a lift generating mechanism.

Reduced air-flow over the new aerofoil would cause turbulent flow as well as

the transition from glider to rotor-craft can cause serious conditions of stall

which would be difficult to overcome in low density hostile Maritain

Atmosphere.

Studies by NASA and ESA have been undertaken to investigate feasibility of

having an UAV mission to Mars.[3] NASA under its MARS SCOUT PROGRAM

developed ARES which is considered here for our base design and X-I’s

design for the rotorcraft is its optimisation of it as its future concept. ESA on

the other hand has undertaken studies for an inflatable rotor concept under its

SOW program.

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The Mars plane thus offers an extensive scaled, high measurement surveyor

from a varied geographical and topological regions providing with a new and

better understanding of the Martian atmosphere as well as surface and its

geological interiors.

The primary objectives of MANE mission are explained below. Note the mission objectives of MANE mission aren't the primary objectives of the design X-I but the aircraft under MANE mission would achieve such science goals with on-board instrumentation.

1.2.MISSION SCIENCE OBJECTIVES:

Extensive study of the Martian atmosphere identifying its key characteristics.

Understand the atmospheric changes, pattern of Martian storms and

atmospheric composition with dust detector and spectrometer.

Perform a radiation analysis of Martian atmosphere

Perform remote sensing of the observation area through a high spatial

resolution on-board camera.

To analyse and create a profile of water (in the form of hydrated minerals,

adsorbed water, or possibly ice at the deepest level) and mineral

abundances near the surface through on-board spectrometer.

Identify key landing sites on Mars for future manned mission by performing a

detailed analysis of the landscape features using 3-D mapping.

1.3.LOCATION OF THE MISSION:

After studying the Maritain Geology and going through the reports of

previous mission to Mars conducted by NASA and Soviet Union[4] ;

Hellas Planitia was chosen for the mission as described through Figure

2.1. It’s a circular crater located at the southern hemisphere of Mars

(42.7°S , 70°E) in Hellas Basin with a crater depth of 7,152 m which

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extends over 2,300km as shown in Figure 2.1 and 2.2. The reason for

choosing Hellas Planitia for the mission was it offers 1,155 Pa of

atmospheric pressure at the bottom which was 89% higher than the

planet’s average. Such high pressure increases the possibility of finding

water in its liquid phase (under Martian temperature). Also the gullies

around it namely: Dao Vallis and Reull Vallis are low into the Martian

crust, making water exist there in its liquid form.

Figure 2.1: Location of Hellas Planitia on Mars[4]

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Figure 2.2: Hellas Planitia extends across about 50° in longitude and more

than 20° in latitude. From data from the Mars Orbiter LaserAltimeter (MOLA) [4]

1.4.ENVIRONMENT:

The Martian atmosphere is the least friendly atmosphere in terms of

design engineering in the solar system. The study here is based around

the Mars Pathfinder Atmospheric Structure Investigation/Meteorology

(ASI/MET) Experiment. The atmospheric density, pressure and

temperature profiles which were observed by the Pathfinder EDL were

compared to the Viking-1 results which are shown in the Figure 2.3.

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Figure 2.3: Atmospheric density and Temperature profile of Martian

Atmosphere with respect to altitude [5]

[5]The atmospheric densities measured by the Pathfinder varied from ~5 x

kg/ at the threshold of detection to 8 x kg/ at 9 km. An increase in

density was measured by the Pathfinder MET from 90 to 80 km height, which

was in respect to the temperature minimum. Below 30 km, the Pathfinder

measured lower values of density and pressure which was in consistent to the

lower overall mass and surface pressure of the Martian atmosphere at the time

of Pathfinder landing. A lower atmospheric density requires a higher minimal

velocity to assure steady flight, which in turn infers that we would have higher

power requirements to counteract it. This leads to higher fuel mass and hence

design optimisation through studying the trade-off between various propulsion

engines would be undertaken.

With respect to Martian thermosphere, the temperature increased rapidly with

altitude due to heating by solar ultraviolet radiation which was measured above

125 km. The temperature minimum was measured at 92K at 80 Km. Here the

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temperature profile is lower than the CO2 condensation temperature.

Possibilities of CO2 forming high-altitude clouds in this region are highly likely

hence an increasing atmospheric density. Below 16.5 km, the temperature

decreases from 200 to 181 K at 10 km. This inversion is below the condensation

temperature of water vapour and hence may be included in our observation

strategy while landing.

The pressure profile in Mars depends upon the Martian day and night where

variations were observed in between 0.2 mbar to 0.3mbar as shown in Figure

2.4. During daily pressure cycles, two maxima and two minima were observed.

This was accompanied by presence of a large semidiurnal tidal oscillation which

indicated atmospheric dustiness between altitude ranging from 10 to 20 km.

Comparing this figure to that of earth, the Martian atmospheric density is roughly

equivalent to the density at an altitude of 100,00 feet on Earth. This could be

used to test the designs flight performance and the results thus obtained could

be used to optimise the lift. Figure 2.4 suggests that it is preferable to conduct

the mission at 6hr-12 hr local time duration of Maritain Day to have minimum

pressure fluctuation.

Hence from the study of the Maritain environment, it was found that the average

density of Martian atmosphere is 0.02 kg/ while the optimum height of cruise

flight was 1.5 km and the average Mach speed at that height was 242 km/hr

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Figure 2.4: Time averaged surface pressure measurement by the

MET instrument over 30 days of Pathfinder mission [5]

A pressure, wind and temperature variation was observed during Martian dust

storm. It caused a rhythmic fluctuation in temperature and pressure minimum

and maximum as shown in Figure 2.5. The vortex speed of such storm is quite

high compared to the regular Martian wind and hence mission should be

avoided under such unfavourable circumstances.

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FIG 2.5:Pressure, wind and temperature change associated with

dust devil(storm). [5]

Solar flux density on Mars is 43% in comparison to Earth. This means the solar

cell area required to generate the same power in comparison to Earth would

need to be larger by a factor of 2.3. Such requirements put a constraint on the

availability of solar power as a method of propulsion and hence it was discarded

an option.

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1.5.NASA ARES MISSION:

Under NASA’s Mars Scout Opportunity program; NASA’s Langley

Research Centre teamed with the Jet Propulsion Laboratory, Lockheed

Martin, Aurora Flight Sciences, Charles Stark Draper Laboratory, Malin

Space Science Systems and other academic researchers came with the

idea of Mars Aerial Regional-scale Environment Survey (ARES). The

main objective of the aircraft was to autonomously fly a pre-planned

aerial survey approximately 1.5 km above the surface of Mars with a

variation of <10% over 10 km height and range ≥ 500km in the southern

highlands. It carried scientific instruments like MAG Sensor, Mass

Spectrometer and Mars Observer Camera (MOC) as shown in Figure

2.6. It was designed to fit within 2.65m diameter Viking-derivative aero

shell shape(max. internal diameter of 2.48m) while surviving the G-force,

radiation and thermal environment involved with different aspects of the

mission.[6]

FIG 2.6: ARES Instrumentation[6]

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While designing, the aerodynamic analysis of the aircraft was undertaken to

determine design drivers which was performed using the software VORVIEW.

The input geometry for VORVIEW was generated using Rapid Aircraft Modeller

(RAM) which was developed by NASA Ames Research Centre. Prediction of

maximum lift coefficient (CL max) and drag requires methods which encompass

viscous flow effects. Hence handbook estimates couldn’t explain the unusual

Mach and Reynolds Number found during simulation. The aerofoil section was

estimated using the MSES code developed by Mark Drela of MIT. The 3-D

aerodynamic predictions were undertaken using a non-linear Weissinger

method developed at UC Davis. The final results thus are shown below in

Figure 2.7.[7] These parameters were considered as the benchmark figures on

which the X-I design was optimised and its performances were compared to

these graphs.

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Figure 2.7: Ares performance characteristics as calculated using VORVIEW[7]

In Ares for stowing, spring-loaded folds were considered as they are the

simplest and offer low risk packaging. The main objective and a key design

parameter was to have minimum number of folds to maintain aircraft’s structural

integrity. Both Rogallo type parawing and inflatable wing were considered as a

wing option to accommodate the plane inside the aero shell. The main

drawback for the inflatable wing was its performance under Martian atmosphere

after traversing a year in the cold space environment. Hence the idea was

discarded and a traditional method was chosen.

For Propulsion in ARES, unconventional options were considered because of

the low atmospheric density in Martian atmosphere. Monopropellant fuel was

considered as they carried both fuel and oxidizer on-board In the end, rocket

motors (60N bi-propellant thruster) were considered because of its low risk. The

propeller thrust system risk involves blade sizing to ensure efficiency in low

density Martian atmosphere.

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2. CALCULATIONS

2.1. Considerations:

To generate the principle behind X-I’s flight, an experimental design was

generated using a craft mass of 154 kg. The fuel mass budget thus

allocated was 54 kg with a dry mass of 100 kg. The constant values used

in the calculations were:

Mass: 154 kg

Weight on Mars (W): 562.1 N

Acceleration due to gravity on Mars (g): 3.65 m/ [8]

Average density of Maritain atmosphere at 1.5 km (ρ): 0.02 kg/ [8]

Cruise Flight Height (h): 1.5 km

Drag Co-efficient for the glider configuration: 0.04

Lift Co-efficient for the glider configuration: 0.6

Using NASA’s Mars Atmospheric model at a height of 1.5 km[8],

Temperature (T) = -10.34-0.001217h = -26.73 C = 246.27 K

Pressure at the height (h) = 0.699 x = 611.51376 Pa

Gas constant on Mars = 188.92 J/Kg/K[9]

Gamma ( on Mars = 1.3(as the atmosphere has 95% C

Speed of sound at the height(h) = √ = √ = 245

m/s

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2.2. Initial Thrust requirement:

For the initial calculation for thrust requirements, from the rotor blades

and engine thrust, the Breguet Range Equation diagram for a steady

flight as shown in figure 3.1was modified into figure 3.2.

a) Aeroplane

b) Autogiro

Figure 3.1 Forces acting around the centre of gravity while on steady

flight

Thrust

LiftResultant force

on rotor

Net drag

from rotor

& airframe

Flow is

upward

through

the rotor

Weight

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Figure 3.2: Modified Breguet Range Diagram for X-I

Resolving the force vectors, the vectors were shifted around the centre of

gravity to find a steady state equation. The symbols in the figure stand for the

following:

= Rotor thrust

= Rocket thrust

W= Weight of the craft

D = Glider drag force

= Rotor drag force

L = Lift force generated by the glider

Θ = Flight path angle

Φ = Rotor angle with respect to the plane of hub

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For a design with total surface area of 10.8 (explained in details in Design

section) an initial calculations was performed to find :

√

/s

Thereafter, the minimum velocity required through the rotor for autorotation was

calculated using the powered lift theorem [8] where was found.

Now resolving this vector, the minimum forward velocity required for the steady

flight could be calculated as shown in figure 3.3.

Figure 3.3

19.939 m/s = x sin(15)

x = 77.03 m/s

Minimum Forward Velocity required for autogyro performance = 74.405 m/s

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Hence the minimum velocity margin is (91.918-74.405) = 17.513 m/s. This

means, the craft is able to generate a lift even it falls below its stall speed up to

74.405 m/s and maintain a forward flight without diving.

Now, total lift force required to be created =

= 0.5 x 0.02 x x 10.8 x 0.6

= 548.22 N

Using a lift-drag ratio of 14:1[9], Drag generated by the aerofoil = 0.071 x 726 N

= 51.546 N

Now applying conservation of momentum in a steady flight as shown in Figure

3.2 i.e. Sum of forces in all directions is zero; we have

= 0 ∑ = 0

∑ = 0

+ sinθ + L cosθ= W + Dsinθ +

=

………….(1)

An experimental value for was calculated using Blade Element Theory [9],

where for a flight angle of attack, θ= 5° and Blade angle,

and forward propulsion thrust of 85 N, equation(1) =>

= ( ( ( (

(

= 42.715 N

And

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∑ = 0

……………. (2)

For small flight angles, and rotor blade angle,

Therefore,

For an autogiro, the Drag component of the blade must be equivalent to the

sin component of the rotor thrust in cruise flight. Hence Equation(1) =>

Hence, the forward propulsive thrust requirement is 63 N while the total rotor

thrust required to optimise our design is 42.715 N.

2.3.Rotor Blade Calculations:

Since the maximum rotor area is a function of the fairing size, initial

calculations for the sizing of rotor blade were done considering a pure

helicopter configuration were the required thrust to be generated was

56.351N [+15% margin on ] .The maximum fairing diameter of 4-5

m(figures based on Atlas 5, Ariane 5, Delta rocket fairing sizes) [10] allows

us a maximum rotor area of 16.04 . This is further explained in the

design section. Now,

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Induced velocity, =√

=√

=9.36 m/s

Applying conservation of momentum on the controlled volume:

Thrust = Mass flow rate x 2 x Induced Velocity [from actuator disc theory]

Thrust =

=

= 3.01 kg/m/s

Now,

=> = 3.01kg/m/s[Where rdr is the elemental radius of the

actuator disc]

rdr =

= 1.131 m

Now integrating both sides with respect to r

∫

∫

R = 2.26 m

Applying blade element calculation for a two-blade rotor(N=2) and a solidarity

value(s) of 0.1 for a figure of merit of 75% [10]

S =

[where c is the chord length]

c = 0.3555 m

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2.4. ACTUATOR DISC THEORY-AUTOGIRO

Figure 3.4: Velocity Vectors and distribution around the rotor

blade (actuator disc) [9]

In figure 3.4, the elementary theory of actuator disc is being applied

to an autogiro. Instead of creating a thrust using rotor power, an

autogiro utilises the upward inflow from a horizontal velocity vector

to generate a torque around the disc, which in turns creates the

necessary power required to generate rotor lift. This is explained

through figure 3.5.

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Figure 3.5: Co-ordinate system of the incoming velocity component around

an autogiro blade, where V is the upward velocity vector [9]

The upward velocity vector around the rotor is resolved and vectors are shifted

around the blade plane as shown in Figure 3.5. The design has been optimised

around the maximum available theoritical forward velocity of 120 m/s (a function

of forward thrust and propulsion system) and a disc tilt of 15°. Though , a forward

velocity of 74.405 m/s is required to sustain the autogyro’s lift, the maximum

available velocity in the upward direction is 31.05 m/s [V sin (15)]. Hence the

performances of the design were calculated within 19.9 m/s – 31.05m/s range.

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Figure 3.6: Overall Inflow Velocity around the blade plane[9]

Velocity component parallel to the rotor disc: = V cosϕ = 120 * cos 15

115.911 m/s

Velocity normal to the rotor disc, ϕ - = 21.69 m/s

Hence, Overall Inflow Velocity, V’= √ = 120.364 m/s

For an upward velocity, V perpendicular to the disc plane; V’ tends to be equal to

the total rotor tip velocity. Therefore, ( = 120.36 m/s = Mach 0.5 = ΩR

Where Ω is the angular velocity of the rotor blade = 0.94 rad/s

Normalising the velocity component with respect to tip velocity:

µ =

= 0.997 =

0.077

=

= 0.96 =

= 0.2579

=

0.18

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2.5. BLADE ELEMENT THEORY

Figure 3.7 : Elemental forces acting around a blade

In figure 3.7, the elemental forces acting around an autogiro blade are shown

……….. (3)

Where, = angle of attack

= pitch angle

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For an incoming velocity , the blade incoming velocity in the disc plane due to

forward velocity is as explained in Figure 3.8

Figure 3.8 Incoming velocity around a Rotor

Here, for the advancing side is = sinψ +ΩR [to determine total incoming

velocity]

Where ψ is the azimuth angle

For an elemental distance along the rotor blade x, where x =

Ωr = ΩRx =

Therefore,

= ( ) [From normalisation of vector components]

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Again, velocity perpendicular to the blade, [From figure 3.5]

(

= (

Total inflow velocity, √

√( (

= ( ) [ since for

Now, inflow angle, ϕ =

Φ = ( (

(

For small angles, ϕ = (

( =

( ………(4)

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2.6.CYCLIC PITCH ANGLE

The pitch angle is a function of the cyclic pitch which changes with the

azimuth angle ψ. Cyclic pitch is defined as the change of the angle of

attack of the rotor blades, which controls the movement of the aircraft

forward, backwards or sideways. Expressing this change around the first

order of Fourier series gives a formula for θ as:

(

Where the cyclic pitch is applied longitudinally and is the cyclic pitch

applied laterally. This is explained in details in figure 3.9.

Figure 3.9 : Variation of pitch angle as a function of azimuth angle

where is the rotor coning angle

Simplifying Equation 5 by considering, since we have flaps on the

wing surface and tails to provide us with longitudinal pitch

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From above,

= +

( ……………(6)

Applying Blade Element Theory to Figure 3.6, the elemental lift and drag forces

on the blade are as shown in figure 3.10:

Figure 3.10: Elemental forces on the blade[9]

From lift equation, dL =

=

[

from lift- curve

slope]

And drag equation, dD =

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For ϕ << 1 and a lift to drag ratio of 14:1 [stated above], and applying it on forces

around the blade through resolving of vector components:

dT = dL cosϕ + dD sinϕ

14*1 +1 *0.1

dT +dD

dĤ = dD cosϕ – dL sinϕ

1*1-14*0.1

dĤ 0

Hence no torque is generated by the lift and drag on the blade which explains the

absence of tail rotor in an autogiro to sustain a counter-torque to the forces of the

main rotor blade. The minimum torque generated could be overcome through

pitching the blades at different angles around the rotor hub.

Now, dT = dL + dD

=

+

= ( ) ; dr =Rdx and equation (6) substituted in the above

equation

dT= =

( (

+

( +

(

*

……………(7)

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From experimental verifications of a blade with similar characteristics to that

of the above specified at NASA’s Ames Research Centre and NASA’s Langley

Low-Turbulence Pressure Tunnel (LTPT) under similar atmosphere to

that of Mars; the parameters like drag coefficient

( ( ( were estimated for initial

calculations.[11]

The blades under this experiment were tested at a tip Mach number of 0.65

and a Reynolds number of 50,000 for an Eppler 387 airfoil. A software called

OVERFLOW-D was used for the experimentation by the Research Centre. It

was found that in order to produce a thrust co-efficient greater than 0.1 and a

figure of merit of 0.4; there was a requirement of collective pitch equivalent to

8.4 degrees. While the design of X-I was improvised around a value of 0.6

which corresponds to an angle of attack ( of 5 degrees as shown in figure

3.10.[12] Also, the value of for theoretical evaluation was considered to be

0.04(streamline boy). Hence the value of a( was found to be 0.12 . This

value of collective pitch was used an experimental value to determine the

actual collective pitch through method of substitutive evaluation,

Figure 3.11: Vs α experimental graph for Eppler 387 at a Reynold’s[12]

Number of 60000

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2.7.THRUST PER BLADE USING BLADE ELEMENT

THEORY

In order to find the Thrust along the blade element, x is integrated from 0 to

1 :

This elemental thrust is then integrated again as a function of ψ to give the

total thrust generated per blade:

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Solving the above equations analytically using Wolfram Alpha and

simplifying the terms of the general solution using the values of constants

derived above:

Integrating the above equation from 0 to 2π along the azimuth angle would

provide an average thrust around the two blades:

Using Wolfram Alpha and Mathematica to integral of equation 10 and above;

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From initial calculations, let the average thrust required to be generated from the

Rotor = Average thrust required to maintain autogyro rotation = 56.351 N =

Substituting this into equation 11:

56.351 = 20.7944-0.96

=> =-37.03°

Hence, the non-dimensional thrust coefficient,

= 0.02424

2.8. MEAN ROLLING MOMENT

The mean rolling moment experienced by the blade is explained in Figure

3.12

Figure 3.12: Mean Rolling Moment around a blade[9]

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Here an elemental portion of the blade experiences a mean torque

due to the incoming velocity defined by:

Page 37 of 78

Page 38 of 78

2.9. AUTOROTATION AND MINIMUM VELOCITY

CALCULATIONS

Figure 3.13: Elementary forces around a blade while autorotation[9]

When undergoing autorotation, a blade performs on the thrust generated due to

the downward movement of the craft because of its weight. This thrust

provides the necessary lift to maintain a steady rate of descent without

undergoing stall, which is the same principle used in an autogiro. This is

explained through figure 3.13

Page 39 of 78

From elemental Lift equation, dL =

( (

dT =

( (

)

dT =

( (

T =

∫ (

T =

(

Now, dĤ = dD-dLφ =

(

(

(

)

dQ = r* dĤ =

(

)

Q= N ∫

(

] *

Here, Thrust (T) =

(

]

Therefore, Q =

(

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Now,

(

&

(

Therefore,

(

)

=

(

)

Since

(

)

From actuator Disc theorem, T = (

Substituting this in the equation for

(

)

(

)

Therefore,

(

)

(

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Substituting the values from above results;

Now, =

∫ ∫ (

)

, the equation

was solved using Wolfram Alpha

= 53.1065°(unrealistic)

Again, in figure 3.11:

= (

) (

)

For smaller angles,

=

For sinψ =1 and x =1,

= 5.46°

And, for sinψ =0 and x =1,

=10.641°

Again, - = 47.64°

And

= +

= 10.616°

Page 42 of 78

2.10 INDUCED POWER FACTOR , FIGURE OF MERIT, LIFT &

DRAG COEFFICIENT

For n= 2 for a parabolic distribution of induced velocity from the rotor

centre to the tip;

√ √

For a thrust, T = 56.351 N

The induced power factor of for n =2 is 1.30

Hence Total rotor power; P = √

P=803.28 W

Figure of Merit =

= √

Figure of Merit = 0.6561

Profile Drag co-efficient is given by:

For a Mach flow of 0.50,

M =

√

√

= 0.202

Page 43 of 78

2.11 GLIDE PHASE CALCULATIONS

Figure 3.14 Aircraft in a symmetric flight

Drawing a free body diagram of the aircraft in glide phase as shown in

figure 3.14, here is the symmetric flight path angle, θ is the pitch angle

between the ground plane and longitudinal axis while is the angle of

attack. Now, applying Breguet Range Equation to the above diagram and

considering thrust vector in the direction of flight in a steady flight:

………12

The power required to perform this steady flight is defined as:

, where V is the forward velocity

In gliding phase, T=0, so substituting and simplifying equation 12;

-

Page 44 of 78

In order to have a constant velocity over the aerofoil (maintaining a speed

equivalent to ), the aircraft descents at a rate of descent (

with a

decent angle of – implying

So equation 12 implies: D = Wsin

(

And, Lift =

= W cos (

tan (-

……..13

For maximum range,

=

(

=>

=

……..14

Now,

Applying this to Equation 14:

Where e is Oswald’s Factor

Page 45 of 78

Rate of Descent (RD):

For a constant dry mass weight (W), surface area(S) and density; the rate of

descent is minimum when

is minimum.

Hence, the time it would hit the ground is given by =

t =

√

And maximum range travelled is:

S =

For, h =1.5 Km and tan (-

, RD = 5.585 m/s

t = 369.258 seconds

s= 21126.76056 m =21.126 km

These calculations allow us to extend the mission time beyond the cruise flight

limit (i.e. when fuel mass -> 0) and also calculate the initial distance travelled

by the X-I in its supersonic glider configuration. For this, it is considered that

the flight is gliding throughout its descent from a height of 10 km.

t’ = 1983.71sec s = 166666.666 m = 166.66 km

Page 46 of 78

WINGED/GLIDER CONFIGURATION

Vstall(m/s) 105.00 103.00 100.00 95.00 90.00 85.00 80.00 75.00 70.00 65.00 60.00 55.00

Wing Area(m2) 8.50 8.84 9.38 10.39 11.57 12.98 14.65 16.67 19.13 22.19 26.04 30.99

Drag Force(N) 40.75 39.21 36.96 33.36 29.94 26.70 23.65 20.79 18.11 15.62 13.31 11.18

Lift Force(N) 582.12 560.16 528.00 476.52 427.68 381.48 337.92 297.00 258.72 223.08 190.08 159.72

THRUST CALCULATIONS

Autorotation Velocity(Vd) 25 25 25 25 25 25 25 25 25 25 25 25

Rotor Tilt angle(in radians) 0.2618 0.2617994 0.2617994 0.261799 0.261799 0.2618 0.261799 0.261799 0.261799 0.261799 0.261799 0.261799

Minimum Forward Velocity(m/s) 24.1481 24.148146 24.148146 24.14815 24.14815 24.1481 24.14815 24.14815 24.14815 24.14815 24.14815 24.14815

Rotor Thrust required(N) -27.103 57.762412 182.00056 380.904 569.6073 748.11 916.4133 1074.516 1222.419 1360.121 1487.623 1604.925

Rotor Drag(N) -7.0148 14.950012 47.105212 98.58521 147.4252 193.625 237.1852 278.1052 316.3852 352.0252 385.0252 415.3852

Rotor Thrust(function Rotor Drag)(N) -82.821 24.009624 180.40298 430.7864 668.3297 893.033 1104.896 1303.918 1490.101 1663.443 1823.946 1971.608

Forward Propulsion Thrust(N) -41.066 66.271727 223.40762 474.9799 713.651 939.421 1152.29 1352.257 1539.324 1713.489 1874.754 2023.117

Rotor Thrust(with constant lift and forward velocity)2.64447 24.60927 56.76447 108.2445 157.0845 203.284 246.8445 287.7645 326.0445 361.6845 394.6845 425.0445

Rotor Drag(Const Lift)(N) 0.68444 6.3693479 14.691726 28.01573 40.65645 52.6139 63.88805 74.47893 84.38652 93.61083 102.1519 110.0096

Forward Propulsion Thrust(Const Area Lift)43.4799 64.630102 95.592767 145.1635 192.1922 236.679 278.6233 318.0257 354.886 389.2042 420.9804 450.2144

3. DESIGN TRADE-OFF

From the formulas derived in section 2, calculations were done to optimise the

design parameters as well as performance of the aircraft. Initially, the

fundamental design parameters were identified to be that of in terms of

being related to the wing area, Rotor blade radius(R) in terms of being within the

fairing diameter constraints, Upflow Velocity ( in terms of determent of

autogyro performance) and Rotor Tilt angle. These values were taken arbitrarily

around the theoretical predictions and as shown through the tables below to have

an optimum design which itself produces the final requirements for these

fundamental parameters. The values shown within the yellow columns are the

parameters derived from theoretical predictions to prove the authenticity of the

calculations.

Initially, a trade-off was done around the Glider configuration to determine Thrust

and Lift requirements as shown in Table 3.1.

Table 3.1: Glider Design Trade-off

From this table, values for Lift force requirement for the craft were generated to

generate the Thrust requirements as shown in Table 3.2 below

Table 3.2: Thrust and Drag calculations

Page 47 of 78

ROTO

R CAL

CULA

TION

Induc

ed Ve

locity

gene

rated

(Vi)(

m/s)

2.030

6.193

9.406

12.98

915

.647

17.80

019

.615

21.17

822

.543

23.74

324

.802

25.73

9

Veloc

ity Pa

rallel

to th

e disc

(Vx)(

m/s)

101.4

2299

.490

96.59

391

.763

86.93

382

.104

77.27

472

.444

67.61

562

.785

57.95

653

.126

Vz(m

/s)27

.176

26.65

825

.882

24.58

823

.294

22.00

020

.706

19.41

118

.117

16.82

315

.529

14.23

5

Veloc

ity No

rmal

to the

disc(

Vz-V

i)(m/

s)25

.146

20.46

516

.476

11.59

97.6

474.2

001.0

91-1.

767

-4.42

5-6.

920

-9.27

3-11

.504

Overa

ll Infl

ow Ve

locity

(Vt)(

m/s)

107.9

6910

5.013

101.3

4895

.705

90.32

485

.104

80.00

775

.021

70.14

065

.367

60.71

256

.190

Rotor

mach

numb

er(Ma

ch)

0.446

0.434

0.419

0.395

0.373

0.352

0.331

0.310

0.290

0.270

0.251

0.232

µ(V/V

t)0.9

730.9

810.9

870.9

930.9

960.9

991.0

001.0

000.9

980.9

940.9

880.9

79

µx(V

x/Vt)

0.939

0.947

0.953

0.959

0.962

0.965

0.966

0.966

0.964

0.960

0.955

0.945

λi(Vx

/Vt)

0.019

0.059

0.093

0.136

0.173

0.209

0.245

0.282

0.321

0.363

0.409

0.458

µzd(V

z-Vi/V

t)0.2

330.1

950.1

630.1

210.0

850.0

490.0

14-0.

024

-0.06

3-0.

106

-0.15

3-0.

205

µz(V

z/Vt)

0.252

0.254

0.255

0.257

0.258

0.259

0.259

0.259

0.258

0.257

0.256

0.253

Rotor

Tilt a

ngle(

φ)6.8

485.7

154.7

583.5

402.4

701.4

390.3

97-0.

686

-1.84

0-3.

091

-4.46

8-6.

007

Coeff

icien

t of T

hrust(

Ct) fu

nctio

n of c

onsta

nt Av

g thru

st0.0

300.0

320.0

340.0

380.0

430.0

490.0

550.0

620.0

710.0

820.0

950.1

11

Coeff

icien

t of T

hrust(

Ct) fu

nctio

n of s

olidir

ity,th

eta_n

ot, a

0.032

0.032

0.031

0.031

0.031

0.031

0.031

0.031

0.030

0.030

0.030

0.030

Avera

ge Th

rust(T

avg)

as a f

uncti

on of

Ct59

.387

55.86

051

.778

45.88

940

.653

35.90

231

.564

27.60

123

.988

20.70

517

.741

15.08

4

Cycli

c Pitc

h(B1)(

degre

es)

-40.20

0-36

.527

-32.27

4-26

.140

-20.68

6-15

.737

-11.21

8-7.

091

-3.32

60.0

933.1

815.9

49

Blade

Lift C

oeffi

cient(

Cl_av

erage

)2.1

152.1

032.0

932.0

802.0

692.0

582.0

472.0

362.0

252.0

121.9

981.9

84

Profile

Drag

Coeff

icent(

Cdo)

0.724

0.740

0.764

0.806

0.851

0.900

0.955

1.014

1.080

1.153

1.234

1.324

Vd(as

a fun

ction

of fu

nctio

n of s

,, a)

33.31

332

.084

30.70

828

.693

26.83

025

.056

23.34

621

.689

20.08

118

.519

17.00

615

.544

Vd(as

a fun

ction

of co

nstan

t thru

st)30

.464

32.57

135

.601

41.35

548

.412

57.20

168

.254

82.28

010

0.241

123.4

6315

3.774

193.7

01

Ideal

Induc

ed po

wer(f

or n=

2)(W_

156.4

3544

8.875

631.9

1477

3.370

825.3

3982

9.164

803.3

0675

8.444

701.6

1963

7.857

570.9

2250

3.728

Total

Rotor

Powe

r(W)

319.6

8152

9.160

651.7

0973

4.727

752.6

7973

6.558

700.1

3965

1.338

595.3

3553

5.791

475.4

2041

6.296

Figure

of M

erit

0.377

0.654

0.747

0.811

0.845

0.868

0.884

0.897

0.908

0.918

0.926

0.933

Hence, the values obtained for Rotor Thrust requirement from the above

table were used as the design parameters in Table 3.3.

Table 3.3: Rotor Performance Calculations

Page 48 of 78

In the above three tables, the values showed in red are the ones around which

the final design of the baseline concept of X-I has been scaled around. The

values which were chosen finally for our concept are shown below in Table 3.4.

Drag Coefficient(glider) 0.74

Drag Coefficient(autogiro) 0.202

Coefficient of Lift(plane) 0.6

Coefficient of Lift(autogiro blade) 1.61

Wing Area(m2) 8.8

Autorotation Velocity(Vd)(m/s) 30

Rotor Tilt angle(phi) 5

Rotor Drag(Dr)(N) 6.589

Rotor Radius( R) 2.26

Area of Rotor(m2) 16.04

Forward Velocity(m/s) 120

Lift Force(N) 560

Velocity Parallel to disc(Vx)(m/s) 115.911

Velocity Normal to disc(Vz-Vi)(m/s) 21.69

Overall Inflow Velocity/Vtip(Vt)(m/s) 120.36

Theta_not(θo) 10.615

Lateral Cyclic Pith(B1) -37.03

Theta max(θmax) 10.615°

alpha(α) 5°

Average Thrust(Tavg)(N) 56.351

Coefficient of Thrust(Ct) 0.0242

a(Cl/alpha) 0.01

Solidirity(s) 0.09

Coefficing of rolling moment(Mr) 0.0482

Mean Rolling Moment 523.55

Disc Loading(N/m2) 35.06

Chord Length(m) 0.35

Vstall(m/s) 103.215353

Number of blades 2

Lift to Drag Ratio 14:1

Page 49 of 78

4.PROPULSION SYSTEM

The requirement of propulsion system for the chosen configuration from Table 3.2

was in the range of 43 N – 100N. Propeller driven propulsion was discarded in

the initial phase due to the complexity in stowing arrangement. Electric

Propulsion proved insignificant because of their low thrust performance (20mN-

90mN). Hence it was decided that in order to meet the forward propulsive thrust

requirement, options of chemical thrusters would be the most suitable option for

X-I design. Thrusters were considered from EADS Astrium as well as Northrop

Grumman[13]. These options along with the fuel alternatives like cold gas,

monopropellant, bi-propellant were considered which is shown in details in table

4.2.

Storage of Bi-propellant fuel proved to be design constraint to the X-I fuselage as

it reduces the amount of total fuel which could be carried in the mission while

imposing external constraints on tank volumes. Hence the final choice of

propellant was chosen to be Mono-propellant Hydrazine thruster. Northrop

Grumman through its MRE-15 thruster met the requirements of X-I mission [14],

hence was considered as the final choice of thruster type. The characteristics as

well as performance data of MRE-15 thruster is as shown below in the table

Table 4.1: MRE-15 Technical Data[14]

Page 50 of 78

Component type

Descripti

on/

Designati

on Mass I_sp

Inlet

pressure Thrust

Valve

(+heater)

power

accumulated

burn life

Length

(max)

Diameter

(max)

Mixing

ratio

(1) Mono-

prop

(2) Bi-

prop

(3) Cold-

gas

Propella

nt Oxidiser

V(exhaus

t)

per piece per piece per piece

OX /

PROP

[kg] [s] [bar] [N] [W] [hours] [mm] [mm] (m/s)

BI

Astrium 400 N Bi-Propellant Engine

Model

S400-12 3.6 318. 10. 420. 35. 8.3 503. 248. 1.65 2. MMH

(N2O4,

MON-1,

MON-3) 3119.58

BI

Astrium

400 N Bi-Propellant Engine

Model

S400-15 4.3 321. 10. 425. 35. 12.8 669. 316. 1.65 2. MMH

(N2O4,

MON-1,

MON-3) 3149.01BI

Astrium

10 N Bi-Propellant Thruster

(single Seat Valve) 0.35 291. 10. 70. 158.5 103. 1.65 2. MMH

(N2O4,

MON-1,

MON-3) 2854.71

MONO

AMPAC In-Space Propulsion

MONARC-

445 1.6 235. 445. 410. n/a 1. N2H4 n/a 2305.35

MONO

AMPAC In-Space Propulsion

MONARC-

90 1. 235. 90. 300. n/a 1. N2H4 n/a 2305.35

MONO

Astrium

400 N Hydrazine Thruster 2.7 214. 400. 0.485 325. n/a 1. N2H4 n/a 2099.34

MONO

Astrium

20 N Hydrazine Thruster 0.395 224. 20. 10.500 195. n/a 1. N2H4 n/a 2197.44

MONO

Astrium

1 N Mono-Propellant Thruster 0.29 220. 1. 15.9 50.000 172. 30. n/a 1. N2H4 n/a 2158.2

COLD GAS

Moog

0.12 N GN2 Thruster

Solenoid

Actuated

58E142A 0.016 57. 6.9 0.12 16,666.667 n/a 3. GN2 n/a 559.17

COLD GAS

Moog

3.5 N GN2 Thruster

Solenoid

Actuated

58-118 0.022 71.5 14.82 3.5 16,666.667 n/a 3. GN2 n/a 701.415

BI

Astrium

22 N Bipropellant Thruster 0.65 290. 22. 212. 55. 2. MMH MON 2844.9

BI

Northrop Grumman

Dual Mode Liquid Apogee Engine TR-308 322. 6.719 0. 0. 1. 2. N2H4 N2O4 3158.82

BI

Northrop Grumman

High Performance Dual Mode Liquid

Apogee Engine

TR-312-

100YN 330. 35. 6.944 0. 0. 1.06 2. N2H4 N2O4 3237.3

MONO-85N

Northrop Grumman

Monopropellant Thruster MRE-15 1.1 228. 66. 72. 318. 119. n/a 1. H2H4 n/a 2236.68

Table 4.2: Options for Thrusters and their specifications

Page 51 of 78

From table 4.1, design thrust available from MRE-15 = 66 N

Exhaust velocity ( = = 2236.68 m/s

Now,

Force = mass flow rate x exhaust velocity

= 0.0281 kg/s

Total Fuel Budget in final design( = 54 kg

Total flight time =

Now, unifying Rocket Equation and Breguet Range equation for horizontal flight

using rocket thrust, we have:

Range =

Range = 14*228*ln(

= 165389.93 m = 165.389 km

In table 4.2 performances of the thruster were compared with all the available

options from Table 3.1 are shown below:

Table 4.2: Thruster performances for different stall speeds

Vstall(m/s) 120 105 103 100 95 90 85 80 75 70 65 60 55

Wing Area(m2) 8.8 8.5034014 8.8368366 9.375 10.38781 11.5741 12.97578 14.64844 16.66667 19.13265 22.1893491 26.04166667 30.99174

Forward(thrust) 70.93 50.42349 70.934019 100.9602 149.0318 194.638 237.7793 278.4552 316.6659 352.4114 385.691747 416.506844 444.8567

Fuel Mass flow rate(kg/s) 0.031712 0.0225439 0.031714 0.045138 0.066631 0.08702 0.106309 0.124495 0.141579 0.15756 0.1724394 0.186216555 0.198892

Burn Time(s)(for Mfuel=54 kg) 1702.816 2395.3264 1702.7193 1196.32 810.4359 620.54 507.9531 433.7528 381.4137 342.7264 313.153499 289.9849588 271.5048

Range(m) 165389.9 144716.19 141959.7 137824.9 130933.7 124042 117151.2 110260 103368.7 96477.46 89586.2158 82694.96839 75803.72

Range in (Km) 165.3899 144.71619 141.9597 137.8249 130.9337 124.042 117.1512 110.26 103.3687 96.47746 89.5862158 82.69496839 75.80372

Page 52 of 78

5.FINAL DESIGN GENERATION

For the initial phase of concept generation, alternative designs of rectangular, v

shaped as well as delta wings were considered. The aircraft in initial glide stage

required to undergo speed up to Mach 3 while the autogiro requirement speed

was of Mach 0.50. Delta wings have been known to be effective under such flight

regime compared to rectangular wings. Concepts were generated to

accommodate both delta wing with a convertible wing section were the slats of

the wing could be transformed into the rotor blades. This was done in order to

accommodate the design within the aero shell constraints instead of having a

typical autogyro/rotor craft configuration with separate body and rotors. Further

study into NASA’s Mars Ares Mission[15] suggested that delta wing can sustain

better performance under stowing compared to rectangular wings and bluff

bodies in terms of fuselage structural integrity. The concept of X-I was initially

generated as figure 5.1.

Figure 5.1: Concept Generation of X-I

Page 53 of 78

From inputs from the initial calculations as shown in section 2, the wing half span

had to match the rotor blade radius in order to maintain structural symmetry to

minimise wingtip vortices. Hence, the half span for the design formed the chief

dimensional parameter(Y = 2.26m). The sweep angle was decided to be 30° in

order to sustain a delta wing formation. A wing tip chord (B) was chosen to be

equivalent to the blade chord length of 0.35 m. The design requirement provided

with a constraint to have a minimum wing surface area of 8 to 15 and

hence, an online tool called Aircraft Center of Gravity Calculator[15] was used to

generate rough dimensional figures to accommodate this surface area

requirement as shown in figure 5.2 and figure 5.3. These values formed the

foundation for the final design of X-I.

Page 54 of 78

Figure 5.3: Identification of dimensional parameters from concept

generation-Glider (distances in metres) [15]

Page 55 of 78

Figure 5.4: Identification of dimensional parameters from concept

generation-Autogyro (distances in metres) [15]

The values thus obtained from the Centre of Gravity calculator were then

transformed into a rough sketch as shown in Figure 5.5 using SolidWorks 2013.

As Fig 5.5 shows, some parameters were twirled to maintain structural symmetry.

Also to calculate the optimum CG for the design, an average of CG, NP and MAC

was taken in the generation of the final design. This sketch was used as the

foundation for generating the individual components of the autogiro.

Page 56 of 78

Figure 5.5: 2-D sketch of X-I(units in metres)

From figures 5.6 – figure 5.12 below, the exact parameters along with its

dimensional measurements are explained in details:

Figure 5.6: Dimensions of the main body (units in metres)

Page 57 of 78

In Fig 5.6, a traditional delta wing design was optimised to meet the design

requirement. In order to accumulate maximum control volume of air through the

autogyro blades; the aerofoil section was cut around the side of the leading

edges to accommodate an arced aerofoil. Also, an arced section was used in the

trailing edge of the aerofoil to minimise wingtip vortices as well as to facilitate

absorption of maximum control volume from the column of air below the trailing

edge.

Figure 5.7: Dimensions of the rotor hub (units in metres), height = 0.10 metres

In figure 5.7, the rotor hub was chosen to be of 0.10 meters height to

accommodate a 15°-5° tilt of the rotor blade plane. In the glider phase of the X-I's

mission, the rotor hub sits in front of the propulsion system while the hinges are

stowed around its sides at 15° to its horizontal plane(Fig 5.8).

The mechanism used to control the lateral and longitudinal pitch of the blade is

used in glider phase to control the pitching angle of the rotor blades (when it is

used as extended flaps to the aerofoil section). The deployment of the rotor hub

takes once the aircraft reaches a speed of 0.65 Mach, where the section rises,

Page 58 of 78

rotates by 90° and slides over the cylindrical central fuselage as shown in figure

5.14. The fillet section around the fuselage in the SolidWorks assembly

represents the tracks along which the hub slides over the fuselage. It moves

forward and sits on an extruded hole section (located behind the centre of gravity)

to maintain the nose tip high in air with respect to the central plane of the aircraft

(as shown in Figure 5.18). This is done to optimise as explained in figure 3.13.

Such kind of alignment ensures a steady rate of ascent which is explained in

details in the observation strategy section.

Figure 5.8: Dimensions of the rotor hinges (units in metres), radius = 0.02 metres

The rotor hinges used in X-I design are based on the deployable rotor invented

by Richard H. Hollrock from Kaman Aerospace Corporation [16]. Figures 5.9

explain the deployment of the rotor hinges with the blades attached to its free

end 22.

Page 59 of 78

Fig 5.9: Deployment of rotor hinges. Point 30 and Point 32 are the top edges of

the rotor hub

Page 60 of 78

The design has two tail spars of 2 metre length each placed at an angle of 5°

from the normal plane of the trailing edge of the aerofoil (figure 5.10). They have

been divided into two asymmetric sections for both spars in-order to incorporate

stowing requirements of the design as shown in Figure 6.3.The spars are divided

in an uneven symmetry such that the hinges around which they are revolved after

deployment are placed on an elongated extension from the other half to maintain

structural integrity. The extensions are in a parallel configuration from opposite

sides for both the spars in-order to balance out the moment forces around the

hinges. This allows locking the movement of the hinges due to reaction to the lift

forces around the aerofoil and tail section. The spars allow sustaining the tail at a

height of 44 cm from the main aircraft plane. A height of 44 cm was chosen to

avoid being in the line of discharge of exhaust straight out from the engine

nozzle.

Figure 5.10: Dimensions of the tail spars deployed(units in metres), diameter =

0.05 metres

To accommodate a tail with high manoeuvrability capability, an elongated tail

body was considered instead of a traditional T-tail or V-tail(Figure 5.11).A primary

requirement for the X-I design is being able to undertake sudden changes in flight

path (based around a new science requirement) and hence to perform such

manoeuvres; elongated tail body was used. This gives X-I a high efficiency in

vertical stability of the aircraft.

Page 61 of 78

Figure 5.11: Dimensions of the main tail body (units in metres), thickness = 0.10

metres

The tail blades are aligned at angle of 30° with the main tail section as shown in

figure 5.12. This allows the tail blades to provide vertical stability to the aircraft;

while being inclined at angle they perform as a compact horizontal stabilizer

where they create a pitching moment relative to the main section causing a

torque. The two ruddervators located at its trailing edge provides symmetric pitch

control and asymmetric yaw control to X-I. These are controlled by an

autonomous control system which is driven by a brushless DC motor.

Figure 5.12: Dimensions of the tail blades (units in metres)

Page 62 of 78

The rotor blades were designed around the dimensional parameters achieved

from the initial calculations. The key characteristics of the blade element is as

shown in Figure 5.13 The blades perform both as flaps to the wing surface as

well as their chief role of being rotor blades. The whole design of X-I was based

around the chief requirement of having transformable wings hence making it he

main design parameter around which the rest of the design has been optimised

upon. The blades have the ability to undergo a turn of 180° around its axis which

is controlled by the rotor hinges.

The performance characteristics of the blade as shown in figure 5.13 has been

discussed in details in the calculation section above while the figure explains its

dimennsional parameters.

Figure 5.13: Dimensions of the rotor blades(units in metres), length = 2.26metres

After defining individual component of the aircraft and its properties; it was all

finally assembled into one single unit using SOLIDWORKS 2013. The design was

optimised based around the symmetry of the final configuration which is shown

from Figures 5.14 to 5.17.

Page 63 of 78

Figure 5.14: Dimensional Parameters of final glider design

Figure 5.15: Panoramic view of final glider design, location of rotor hub with

respect to the main fuselage section

Page 64 of 78

Figure 5.16: Dimensional Parameters of final autogiro design

Page 65 of 78

Figure 5.17: Panoramic view of final autogiro design

Page 66 of 78

Figure 5.18: Frontal View of the final design showing the maximum height and

total length of the final design

Figure 5.19: Location of the centre of gravity (in yellow) and its relation with the

rotor hub

Page 67 of 78

Figure5.20: Final design Wing Area estimation and characterisation of

ideal Centre of Gravity (CG), Aerodynamic Centre (AC) and Neutral Point

(NP)[15]

A final calculation based study was done to investigate the fluctuation of

performance parameter as a function of flight velocity. The range was considered

from Vstall to 104 m/s(maximum cruise velocity). The final range as a function of

cruise velocity and thrust requirement is shown in Table 5.1.

Page 68 of 78

Table 5.1: Velocity Vs performance parameters for final design

WINGED/GLIDER CONFIGURATION

Vstall(m/s) 92.00 94.00 96.00 98.00 100.00 102.00 104.00

Wing Area(m2) 11.30 11.30 11.30 11.30 11.30 11.30 11.30

Drag Force(N) 31.28 32.66 34.06 35.50 36.96 38.45 39.98

Lift Force(N) 446.90 466.54 486.60 507.09 528.00 549.33 571.08

THRUST CALCULATIONS

Autorotation Velocity(Vd) 25.00 25.00 25.00 25.00 25.00 25.00 25.00

Rotor Tilt angle(in radians) 0.09 0.09 0.09 0.09 0.09 0.09 0.09

Minimum Forward Velocity(m/s) 24.90 24.90 24.90 24.90 24.90 24.90 24.90

Rotor Thrust required(N) 1292.78 1074.24 851.01 623.07 390.44 153.11 -88.93

Rotor Drag(N) 116.19 96.55 76.49 56.00 35.09 13.76 -7.99

Rotor Thrust(function Rotor Drag)(N) 1402.38 1164.21 920.91 672.49 418.94 160.28 -103.51

Forward Propulsion Thrust(N) 1438.43 1200.83 958.12 710.29 457.36 199.32 -63.83

Rotor Thrust(with constant lift and forward velocity) 144.01 124.37 104.30 83.82 62.91 41.58 19.82

Rotor Drag(Const Lift)(N) 12.94 11.18 9.37 7.53 5.65 3.74 1.78

Forward Propulsion Thrust(Const Area Lift) 175.87 157.53 138.79 119.65 100.12 80.20 59.88

ROTOR CALCULATION

Induced Velocity generated(Vi)(m/s) 14.98 13.92 12.75 11.43 9.90 8.05 5.56

Velocity Parallel to the disc(Vx)(m/s) 91.63 93.62 95.61 97.60 99.60 101.59 103.58

Vz(m/s) 8.27 8.45 8.63 8.81 8.99 9.17 9.35

Velocity Normal to the disc(Vz-Vi)(m/s) -6.71 -5.47 -4.12 -2.62 -0.91 1.12 3.79

Overall Inflow Velocity(Vt)(m/s) 92.24 94.16 96.09 98.04 100.00 102.01 104.07

Rotor mach number(Mach) 0.38 0.39 0.40 0.41 0.41 0.42 0.43

µ(V/Vt) 1.00 1.00 1.00 1.00 1.00 1.00 1.00

µx(Vx/Vt) 0.99 0.99 1.00 1.00 1.00 1.00 1.00

λi(Vx/Vt) 0.16 0.15 0.13 0.12 0.10 0.08 0.05

µzd(Vz-Vi/Vt) -0.07 -0.06 -0.04 -0.03 -0.01 0.01 0.04

µz(Vz/Vt) 0.09 0.09 0.09 0.09 0.09 0.09 0.09

Rotor Tilt angle(φ) -2.09 -1.67 -1.23 -0.77 -0.26 0.31 1.05

Coefficient of Thrust(Ct) function of constant Avg thrust 0.04 0.04 0.04 0.04 0.04 0.03 0.03

Coefficient of Thrust(Ct) function of solidirity,theta_not, a 0.03 0.03 0.03 0.03 0.03 0.03 0.03

Average Thrust(Tavg) as a function of Ct 42.47 44.34 46.28 48.28 50.37 52.56 54.90

Cyclic Pitch(B1)(degrees) -22.57 -24.53 -26.54 -28.63 -30.81 -33.09 -35.53

Blade Lift Coefficient(Cl_average) 2.07 2.08 2.08 2.09 2.09 2.10 2.10

Profile Drag Coefficent(Cdo) 0.83 0.82 0.80 0.79 0.77 0.76 0.75

Vd(as a function of function of s,, a) 27.47 28.15 28.83 29.53 30.25 31.01 31.84

Vd(as a function of constant thrust) 45.69 43.21 40.92 38.80 36.84 35.02 33.31

Ideal Induced power(for n=2)(W_ 825.49 800.99 765.58 716.03 647.13 548.94 395.96

Total Rotor Power(W) 760.39 749.40 730.39 700.91 656.97 590.99 483.48

Figure of Merit 0.84 0.82 0.81 0.79 0.76 0.72 0.63

Fuel Mass flow rate(kg/s) 0.078631 0.070428 0.06205 0.053495 0.044763 0.035855 0.026771

Burn Time(s)(for Mfuel=54 kg) 686.7545 766.7357 870.2691 1009.445 1206.346 1506.049 2017.095

Range(m) 126799 129555.5 132311.9 135068.4 137824.9 140581.4 143337.9

Range in (Km) 126.799 129.5555 132.3119 135.0684 137.8249 140.5814 143.3379

Page 69 of 78

6. STOWING CONFIGURATION

During the descent into the Martian atmosphere, the X-I is stowed within an aero

shell to protect it from re-entry heat. The descent strategy used in its mission is

similar to that of NASA’s Pathfinder and Mars rover mission. The aero shell

enters the Martian atmosphere at angle of -13° and supersonic parachutes are

deployed to slow the craft down. At an approximate height of 15 km above the

surface, the bottom cover of the aero shell is jettisoned and the wings of X-I are

deployed. At a height of 12 km above the surface, the top half of the aero shell

separates from the main aircraft through an explosion of mortar canister and then

a drogue chute is deployed to slow down the aircraft to a lower Mach speed(<1)

and continue its flight in a glider configuration from until it reaches a speed of 0.5

Mach. At this speed, the rotor hub transforms vertically and slides over the

fuselage into the designated hole in the fuselage while the blades are being

deployed along with it. This would cause structural imbalance in the craft and the

thrusters would be initiated to overcome this imbalance. Thereafter it continues in

a flight as explained in the above sections. In order to describe the descent

sequence of events, Fig 6.2 is used which is the deployment sequence of Ares

Mars Mission whereas Fig 6.1 describes the stowed configuration within the aero

shell.[17]

Fig 6.1: X-I stowed in the aero shell with the dimensional parameters of the shell

Page 70 of 78

Fig 6.2: Descent strategy of X-I based on the Ares entry strategy[7]

Page 71 of 78

Fig 6.3: Stowed design of X-I

Fig 6.3 explains the stowed configuration of X-I. The main aerofoil section was

divided into three parts around the main fuselage body. During the deployment, a

series of pyro burns are undertaken to push the hinges to interlock around the

edges while carbon tape springs around the corners ensure the structural

stiffness. The design was optimised to have minimum number of folds to maintain

structural integrity.

Page 72 of 78

7. ONBOARD SCIENCE INSTRUMENTS

8.1. DUST DETECTOR

The dust environment of the Mars and the Martian atmosphere are of great

interest especially its summer dust storms. The prospect of future Martian

manned mission lies on the study of the effects of such storm on

Martian weather as well as the actual composition o f the particles of

dust. This would allow creating a profile of the Martian atmosphere and

implementing the findings for future Martian missions and descent

strategies. The dust detector used in MARE is a Piezo Dust Detector (PDD)

. It performs consistent dust monitoring for better understanding of dust migration

patterns on the Mars through direct detection of particle impacts. The PPD is

a modular, miniaturized in-situ measurement device. The modular design

allows an addition of detector units to increase the sensor surface or measure

impacts on multiple spacecraft surfaces. The detector has a low mass, low power

consumption, low data rate and small size. This flexible design makes the PDD

easy to accommodate on the spacecraft.

The detector will provide physical parameters of impacting dust and debris

particles such as velocity, mass and impact energy. The size of detectable

particles will be in the range of 1 µm to 1 mm at a velocity of up to 10 km/s. [18]

Parameter Value

Mass 0.5 kg

Power (Operating) 3 W

Data Volume 36 MB/orbit

Table 7.1: Specifications for the Piezo Dust Detector (PDD)

Page 73 of 78

Figure 7.1: CAD model of Piezo Dust Detector (80mm x 40mm x 40 mm)

7.2.SPECTROMETER

The primary science goal of MARE is to analyse and create a profile of

minerals as well as water near the surface of Mars. This is achieved with the

use of The Chandrayaan-1 X-ray spectrometer (C1XS) w h i c h w a s

designed to measure absolute and relative abundances of major rock-

forming elements (principally Mg, Al,Si, Ca, Ti and Fe) in the lunar crust with

spatial resolution 25 km for India’s Lunar mission.[19] The C1XS

spectrometer was designed by the Rutherford Appleton Laboratory (RAL) for

the Indian Space Research Organisation (ISRO) Chandrayaan-1 lunar

mission and launched in 2008. The following instrument has been chosen

as it’s the latest space proven spectrometer.

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Figure 7.2: CAD image of the C1XS Instrument showing coalligned front

detectors, deployable radiation shield and 140field-of-view.

Parameter Value

Mass 5.56 kg

Power (Operating) 25.5 W

Spatial Resolution 25 km

Table 7.2: Specifications for the The Chandrayaan-1 X-ray Spectrometer

(C1XS)

7.3. RADIATION ASSESSMENT DETECTOR:

The Radiation Assessment Detector (RAD) is an energetic particle

detector designed to measure a broad spectrum of energetic particle

radiation. [20] It is a lightweight and energy efficient passive detector which

acquires radiation data from Galactic Cosmic Rays (GCRs) and ionised

particles from Coronal Mass Ejection (CMEs). The acquired information will

be used to assess the potential radiation hazard for future Mars manned

missions and Mars based colonies; and how the radiation dosage effects

the spacecraft subsystems during the entry phase.

Page 75 of 78

The RAD combines both charged and neutral particle detection

capability over a wide dynamic range in a compact, low mass, low power

input.

7.4. 3-D TERRAIN MODELLING

Modern cartography and geological studies rely on satellite data to

enhance our knowledge of morphology. When planning to build scientific bases

on extra-terrestrial ground, a deep study on landing and construction site

requires to be done. By using information about the altitude of each point in

the ground a more useful map can be generated.

A secondary goal of the MARE mission is to reconstruct terrain morphology.

Commonly a combination of high-resolution satellite images and Digital Elevation

Models (DEM) are used to produce this. [21] DEMs contain information about the

relative altitude of each pixel in a picture. From this a 3D surface resembling the

actual terrain can be produced. An imaging algorithm then carries out a warp and

match to combine the high-resolution pictures with the 3D model to create a

virtual terrain.

To generate 3D images of the Martian surface, two instruments are required.

For MARE, apart from the onboard high resolution camera; t h e r e i s a n o t h e r

i n s t r u m e n t c a l l e d the InterFerometric Synthetic Aperture Radar (IFSAR)

system. IFSAR uses a method called phase interferometry to obtain a pair of high

resolution radar images needed to generate DEMs.[22] IFSAR images are

obtained by using a stereo configuration i.e., two instruments pointing on the

same target on the ground are mounted at a fixed distance on the spacecraft.

The expected total system mass is 1.6 kg.[23] This is highly reliable and

being a radar it can select the desired wavelength to be used and observe other

features rather than obtaining just an elevation map (crater composition,

minerals, etc).

Page 76 of 78

7.5. CAMERA

The Mars Orbiter Camera (MOC) was a camera developed for NASA’s

Mars Observer mission. It has three elements primary components: A

narrow angled camera having the capability to take images of very high

spatial resolution (1.4 m per pixel) and two wide angled cameras which

has various functionalities like global imaging, geodesy etc.[24] It is 80 cm

in length and 40 cm in diameter. This camera has been chosen as it has

proven its technological capability by being used in previous Mars mission

(Mars Observer and Ares Mission). It has a mass of approximately 10 kg

and the structure and the material of the camera is explained via figure 8.3

Figure 8.3: Components of the MOC

Page 77 of 78

CONCLUSION

The study done in the report is only a beginning to accomplish a successful rotor-

wing based exploration mission to Mars. Technical issues like vibration test,

deployment test, drop test, aerofoil performance test in wind tunnel under

simulated Martian conditions, stowing configuration optimisation, wing aerofoil

optimisation, control systems definition etc. needs to be undertaken to make a

matured technical approach to the concepts in comparison to other available

options as studied under NASA’s Mars Scout Mission Program namely MARV

and GTMARS[25].

The final range of X-I is derived using the Newtonian laws of motion in a glider-

autogiro-glider configuration. The autogiro with a final cruise velocity of 102 m/s

has a maximum range of 140 km. The final glider phase allows the mission to be

extended further 21 km. In-order to calculate the initial distance covered in the

initial glider configuration, an initial velocity of Mach 1 was considered which is

decreased over to Mach 0.5 thereafter which the rotor blades are deployed. This

is shown in details in the figure below

10 km

1.5 km 1.5 km

0.01 0

2

4

6

8

10

12

0 108 1506.04 1875.298 time->

height

Glider from t=0 - t=1875

autogiro t=108

glider from t = 1506

Final Observational Plan for X-I

Page 78 of 78

Now using the two laws of linear motion:

S =Ut +ÏȢυ

V = u+at

S initial gliger = 31 km.

Total flight distance = 31+140.58+21 = 192.53 km

Comparing the concept of X-I to NASA’s Ares mission, which is its main

competitor, the above described aircraft has a capability of performing a <200 km

range(glider-autogiro-glider configuration) within a flight time of <1 hour while the

Ares can perform a 500+km range in a similar flight period. The X-I wet mass is

21 kg lower than that of Ares while its stowed structure could be fitted within a

2.5m diameter aeroshell; while X-I needs to be within a 5 m diameter aeroshell

(due to single structure of rotor blades) . Further optimisation of stowing

mechanism could bring this diameter constraint down to a region of 2.5 m –

3m(function of the blade length). Further the range could be optimised if the

thruster are operated around a low frequency duty cycle(pulses) instead of single

stage burn. Approximation based around Ares mission data[15] suggests the

mission duration could be increased substantially by a factor of 10%-15% .

However X-I design has greater ability to perform manoeuvring than ARES due to

external stability at slower speeds. As well in this study, the contingency in design

constraints were quite high which led to such conclusion with respect to ARES

mission .

Hence, the mission doesn’t meet the optimum range capabilities as compared to

that of NASA ARES mission, but it displays an opportunity to use a transformable

wing aircraft in Mars and achieve greater performances in stability and flight

control at the cost of greater observational range. Similar design could be

applied to high altitude drones, where the primary requirement of the aircraft is to

fly at lower speeds with capability of performing sudden manoeuvres. Hence

through this project, a new configuration of flight was discussed and it was finally

concluded that such a design is practically possible of performing a steady flight

while experimental tests as described above would make the design realistic.

REFERENCES

1] „UAV mission design for the exploration of Mars‟;ESA/EUROAVIA DESIGN

WORKSHOP 2006;EUROAVIA

2] „Evolution of a mars airplane concept for the ares Mars scout mission‟; Mark D

Guynn and Mark A Croom; NASA Langley Research Center; American Institute

of Aeronautics and Astronautics

3] „Rotorcraft as Mars Scouts‟, G A Briggs, Center for Mars Exploration (CMEX),

NASA Ames Research Centre, NASA

4] „http://www.nasa.gov/mission_pages/mars/missions/‟,

„http://www.russianspaceweb.com/spacecraft_planetary_mars.html‟, Accessed on

March/2013

5] „The Mars Pathfinder Atmospheric Structure

Investigation/Meteorology(ASI/MET) Experiment; J.T. Schofield, J R Barnes;

NASA Ames Research Centre; NASA

[6] „Evolution of a Mars airplane concept for the ARES Mars Scout Mission; Mark

D Guynn , Mark A Croom; NASA Langley Research Center; NASA

[7] „Evolution of a Mars airplane concept for the ARES Mars Scout Mission; Mark

D Guynn , Mark A Croom; NASA Langley Research Center; NASA; AIAA 2003-

6578

[8] http://nssdc.gsfc.nasa.gov/planetary/factsheet/marsfact.html; NASA; Accessed

on March/2013

[9] „Blade Element Theory‟ ; Powered Lift lecture notes; Dr. Scott J I Walker;

University of Southampton

[10] „Aerodynamic design of helicopter‟; Basic Helicopter Aerodynamics; pg 286;

John Seddon and Simon Newman; Wiley

[11] „Computational Analysis of a Protoype Martian Rotorcraft Experiment‟; Kelly

J. Corfeld; Department of Aerospace Engineering; The Pennyslvania State

University; NASA; AIAA-2002-2815

[12] „Computational Analysis of a Protoype Martian Rotorcraft Experiment‟; Kelly

J. Corfeld; Department of Aerospace Engineering; The Pennyslvania State

University; NASA; AIAA-2002-2815

[13] http://cs.astrium.eads.net/sp/spacecraft-propulsion/hydrazine-

thrusters/index.html ; EADS; Accessed in March 2013

[14] http://www.northropgrumman.com

/Capabilities/PropulsionProductsandServices

/Pages/MonopropellantThrusters.aspx; Northropgrumman; Accessed in March

2013

[15] „Aircraft Centre of Gravity Calculator‟; http:// adamone.rchomepage.com

/cg_calc.htm;

[16] „Deployable Rotor‟;Richard H Hollrock; US Patent Classification:

416/142; 244/138.00A; 244/138.00R

[17] „ ARES MISSION OVERVIEW- Capabilities and Requirements of the Robotic

Aerial Platform‟; Henry Wright, Robert Braun; NASA Langley Research Centre;

AIAA 2003-6577

[18] (PDD) : a modular miniaturized In-situ measurement instrument

for dust research.; A. Wolf, R. Laufer, G. Herdrich, T.W. Hyde, Piezo

Dust Detector; CASPER (Center for Astrophysics, Space Physics

and Engineering Research), 43rd Lunar and Planetary Science

Conference (2012)

[19] ”The C1XS X-ray Spectrometer on Chandrayaan-1”; M. Grandea,

B.J. Maddison, C.J. Howe, B.J. Kellett; in Journal of Planetary and

Space Science, Vol. 57, Issue 7, June 2009, page 717-724

[20] The Radiation Assessment Detector (RAD) Investigation, D.

M.Hassler et al.; Space Sci Rev (2012) 170:503:558 DOI

0.1007/s11214-012-9913-1, 14 July 2012, available online:

http://link.springer.com/article/10.1007% 2Fs11214-012-9913-1

[21] D. M. Winker, W. Hunt & C. Hostetler, “Status and

performance of the CALIOP lidar,”proceedings of SPIE,

2004,available online from http://ccplot.org /pub/resources /CALIPSO

/Status%20and% 20performance%20of%20the%20CALIOP%20lidar.pdf

[22] M. A. Richards, “A Beginner‟s Guide to Interferometric SAR Concepts and

Signal Processing,” published in IEEE A&E Systems Magazine, 9 Sept 2007,

vol. 22, no. 9, pp. 5-29

[23] Olsen Technology, Inc., “Satellite Plus L-Band Optical Transport,” available

online from http://www.olson-technology.com/l-band.html

[24] Mars Orbiter Camera, Malin Space Science System,

http://www.msss.com/all_projects/mgs-mars-orbiter-camera.php;

accessed on March 2013

[25] Vertically Planetary Aerial Vehicles: Three planetary bodies and four

conceptual design case; Larry A Young; NASA Rotorcraft Division;

NASA;

“Space , is the final

frontier"-Star Trek