Mane: Mission to mars

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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
. ONBOARD SCIENCE INSTRUMENTS............................................................................................ 72
.1. DUST DETECTOR.................................................................................................................................. 72
.2.SPECTROMETER..................................................................................................................................... 73
.3. RADIATION ASSESSMENT DETECTOR:.................................................................................. 74
.4. 3-D TERRAIN MODELLING............................................................................................................. 75
. . 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
Lift
Resultant 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:

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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

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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°

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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

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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;
-

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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

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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

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WINGED/GLIDERCONFIGURATION
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
WingArea(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
DragForce(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
LiftForce(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
THRUSTCALCULATIONS
AutorotationVelocity(Vd) 25 25 25 25 25 25 25 25 25 25 25 25
RotorTiltangle(inradians) 0.2618 0.2617994 0.2617994 0.261799 0.261799 0.2618 0.261799 0.261799 0.261799 0.261799 0.261799 0.261799
MinimumForwardVelocity(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
RotorThrustrequired(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
RotorDrag(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
RotorThrust(functionRotorDrag)(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
ForwardPropulsionThrust(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
RotorThrust(withconstant liftandforwardvelocity)2.64447 24.60927 56.76447 108.2445 157.0845 203.284 246.8445 287.7645 326.0445 361.6845 394.6845 425.0445
RotorDrag(ConstLift)(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
ForwardPropulsionThrust(ConstAreaLift)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

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ROTORCALCULATION
InducedVelocitygenerated(Vi)(m/s)2.0306.1939.40612.98915.64717.80019.61521.17822.54323.74324.80225.739
VelocityParalleltothedisc(Vx)(m/s)101.42299.49096.59391.76386.93382.10477.27472.44467.61562.78557.95653.126
Vz(m/s)27.17626.65825.88224.58823.29422.00020.70619.41118.11716.82315.52914.235
VelocityNormaltothedisc(Vz-Vi)(m/s)25.14620.46516.47611.5997.6474.2001.091-1.767-4.425-6.920-9.273-11.504
OverallInflowVelocity(Vt)(m/s)107.969105.013101.34895.70590.32485.10480.00775.02170.14065.36760.71256.190
Rotormachnumber(Mach)0.4460.4340.4190.3950.3730.3520.3310.3100.2900.2700.2510.232
µ(V/Vt)0.9730.9810.9870.9930.9960.9991.0001.0000.9980.9940.9880.979
µx(Vx/Vt)0.9390.9470.9530.9590.9620.9650.9660.9660.9640.9600.9550.945
λi(Vx/Vt)0.0190.0590.0930.1360.1730.2090.2450.2820.3210.3630.4090.458
µzd(Vz-Vi/Vt)0.2330.1950.1630.1210.0850.0490.014-0.024-0.063-0.106-0.153-0.205
µz(Vz/Vt)0.2520.2540.2550.2570.2580.2590.2590.2590.2580.2570.2560.253
RotorTiltangle(φ)6.8485.7154.7583.5402.4701.4390.397-0.686-1.840-3.091-4.468-6.007
CoefficientofThrust(Ct)functionofconstantAvgthrust0.0300.0320.0340.0380.0430.0490.0550.0620.0710.0820.0950.111
CoefficientofThrust(Ct)functionofsolidirity,theta_not,a0.0320.0320.0310.0310.0310.0310.0310.0310.0300.0300.0300.030
AverageThrust(Tavg)asafunctionofCt59.38755.86051.77845.88940.65335.90231.56427.60123.98820.70517.74115.084
CyclicPitch(B1)(degrees)-40.200-36.527-32.274-26.140-20.686-15.737-11.218-7.091-3.3260.0933.1815.949
BladeLiftCoefficient(Cl_average)2.1152.1032.0932.0802.0692.0582.0472.0362.0252.0121.9981.984
ProfileDragCoefficent(Cdo)0.7240.7400.7640.8060.8510.9000.9551.0141.0801.1531.2341.324
Vd(asafunctionoffunctionofs,,a)33.31332.08430.70828.69326.83025.05623.34621.68920.08118.51917.00615.544
Vd(asafunctionofconstantthrust)30.46432.57135.60141.35548.41257.20168.25482.280100.241123.463153.774193.701
IdealInducedpower(forn=2)(W_156.435448.875631.914773.370825.339829.164803.306758.444701.619637.857570.922503.728
TotalRotorPower(W)319.681529.160651.709734.727752.679736.558700.139651.338595.335535.791475.420416.296
FigureofMerit0.3770.6540.7470.8110.8450.8680.8840.8970.9080.9180.9260.933
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

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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

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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]

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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 400N 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
400N 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) 3149.01
BI
Astrium
10N 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
400N Hydrazine Thruster 2.7 214. 400. 0.485 325. n/a 1. N2H4 n/a 2099.34
MONO
Astrium
20N Hydrazine Thruster 0.395 224. 20. 10.500 195. n/a 1. N2H4 n/a 2197.44
MONO
Astrium
1N Mono-Propellant Thruster 0.29 220. 1. 15.9 50.000 172. 30. n/a 1. N2H4 n/a 2158.2
COLD GAS
Moog
0.12N GN2Thruster
Solenoid
Actuated
58E142A 0.016 57. 6.9 0.12 16,666.667 n/a 3. GN2 n/a 559.17
COLD GAS
Moog
3.5N GN2Thruster
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
22N 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

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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
WingArea(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
FuelMassflowrate(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
BurnTime(s)(forMfuel=54kg) 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
Rangein(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

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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

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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.

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Figure 5.3: Identification of dimensional parameters from concept
generation-Glider (distances in metres) [15]

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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.

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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)

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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,

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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.

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

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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.

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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)

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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.

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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

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Figure 5.16: Dimensional Parameters of final autogiro design

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Figure 5.17: Panoramic view of final autogiro design

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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

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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.

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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

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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

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

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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.

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. 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 of 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 .1: Speciﬁcations for the Piezo Dust Detector (PDD)

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Figure .1: CAD model of Piezo Dust Detector (80mm x 40mm x 40 mm)
.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 .2: CAD image of the C1XS Instrument showing coalligned front
detectors, deployable radiation shield and 140◦ﬁeld-of-view.
Parameter Value
Mass 5.56 kg
Power (Operating) 25.5 W
Spatial Resolution 25 km
Table .2: Speciﬁcations for the The Chandrayaan-1 X-ray Spectrometer
(C1XS)
.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.

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The RAD combines both charged and neutral particle detection
capability over a wide dynamic range in a compact, low mass, low power
input.
.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).

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. . 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

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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
0.010
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

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Now using the two laws of linear motion:
S =Ut +
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
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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
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„http://www.russianspaceweb.com/spacecraft_planetary_mars.html‟, Accessed on
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[6] „Evolution of a Mars airplane concept for the ARES Mars Scout Mission; Mark
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[7] „Evolution of a Mars airplane concept for the ARES Mars Scout Mission; Mark
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[8] http://nssdc.gsfc.nasa.gov/planetary/factsheet/marsfact.html; NASA; Accessed
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[9] „Blade Element Theory‟ ; Powered Lift lecture notes; Dr. Scott J I Walker;
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[10] „Aerodynamic design of helicopter‟; Basic Helicopter Aerodynamics; pg 286;
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[11] „Computational Analysis of a Protoype Martian Rotorcraft Experiment‟; Kelly
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[12] „Computational Analysis of a Protoype Martian Rotorcraft Experiment‟; Kelly
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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
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[15] „Aircraft Centre of Gravity Calculator‟; http:// adamone.rchomepage.com
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[16] „Deployable Rotor‟;Richard H Hollrock; US Patent Classification:
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[17] „ ARES MISSION OVERVIEW- Capabilities and Requirements of the Robotic
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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,
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[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
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/Status%20and% 20performance%20of%20the%20CALIOP%20lidar.pdf
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[23] Olsen Technology, Inc., “Satellite Plus L-Band Optical Transport,” available
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[24] Mars Orbiter Camera, Malin Space Science System,
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[25] Vertically Planetary Aerial Vehicles: Three planetary bodies and four
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NASA;

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