MANE-Mission to Mars

HUMANUM EST INFINITA POTESTAS… -The power of human imagination is limitless…. Page 1 of 78 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 Page 2 of 78 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. Page 3 of 78 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 Page 4 of 78 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. Page 5 of 78 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 Page 6 of 78 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] Page 7 of 78 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. Page 8 of 78 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 Page 9 of 78 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 Page 10 of 78 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. Page 11 of 78 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. Page 12 of 78 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] Page 13 of 78 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. Page 14 of 78 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. Page 15 of 78 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) = √ m/s Page 16 of 78 =√ = 245 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 Lift Resultant force on rotor Flow is upward through the rotor Net drag from rotor & airframe Thrust Weight b) Autogiro Figure 3.1 Forces acting around the centre of gravity while on steady flight Page 17 of 78 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 Page 18 of 78 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 Page 19 of 78 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 sinθ + L cosθ= W + Dsinθ + ………….(1) = An experimental value for =0 was calculated using Blade Element Theory [9], for a flight angle of attack, θ= 5° and Blade angle, where and forward propulsion thrust of 85 N, equation(1) =>  = ( ( (  = 42.715 N And Page 20 of 78 ( ( ∑ =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 design section. Now, Page 21 of 78 . This is further explained in the 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 Page 22 of 78 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. Page 23 of 78 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. Page 24 of 78 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 ϕ- Velocity normal to the rotor disc, Hence, Overall Inflow Velocity, V’= √ = 21.69 m/s = 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 Page 25 of 78 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 Page 26 of 78 For an incoming velocity forward velocity is , the blade incoming velocity in the disc plane due to 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] Page 27 of 78 Again, velocity perpendicular to the blade, (   = ( √ Total inflow velocity, √(   [From figure 3.5] = ( ( ) [ since for Now, inflow angle, ϕ =  Φ= ( ( ( For small angles, ϕ = ( ( = ( Page 28 of 78 ………(4) 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 Page 29 of 78 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 = Page 30 of 78 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) Page 31 of 78 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 that of Mars; the ( ( (LTPT) under similar parameters ( like were atmosphere to drag coefficient 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 which corresponds to an angle of attack ( 3.10.[12] Also, the value of value of 0.6 of 5 degrees as shown in figure 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 Page 32 of 78 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: Page 33 of 78 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; Page 34 of 78 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] Page 35 of 78 Here an elemental portion of the blade experiences a mean torque due to the incoming velocity defined by: Page 36 of 78 Page 37 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 38 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 = Page 39 of 78 Now, & ( ( ( Therefore,  = ( ) ) Since ( ) ( From actuator Disc theorem, T = Substituting this in the equation for ( ) ( ) ( ) Therefore,    ( Page 40 of 78 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 41 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 42 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 43 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 (  = W cos ( And, Lift =  tan (- ……..13 = For maximum range, => ( = ……..14 Now, Applying this to Equation 14: Where e is Oswald’s Factor Page 44 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 45 of 78 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. WINGED/GLIDER CONFIGURATION Vstall(m/s) Wing Area(m2) Drag Force(N) Lift Force(N) 105.00 8.50 40.75 582.12 103.00 8.84 39.21 560.16 100.00 9.38 36.96 528.00 95.00 10.39 33.36 476.52 90.00 11.57 29.94 427.68 85.00 12.98 26.70 381.48 80.00 14.65 23.65 337.92 75.00 16.67 20.79 297.00 70.00 19.13 18.11 258.72 65.00 22.19 15.62 223.08 60.00 26.04 13.31 190.08 55.00 30.99 11.18 159.72 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 THRUST CALCULATIONS Autorotation Velocity(Vd) 25 25 Rotor Tilt angle(in radians) 0.2618 0.2617994 Minimum Forward Velocity(m/s) 24.1481 24.148146 Rotor Thrust required(N) -27.103 57.762412 Rotor Drag(N) -7.0148 14.950012 Rotor Thrust(function Rotor Drag)(N) -82.821 24.009624 Forward Propulsion Thrust(N) -41.066 66.271727 Rotor Thrust(with constant lift and forward 2.64447 velocity) 24.60927 Rotor Drag(Const Lift)(N) 0.68444 6.3693479 Forward Propulsion Thrust(Const Area Lift) 43.4799 64.630102 25 0.2617994 24.148146 182.00056 47.105212 180.40298 223.40762 56.76447 14.691726 95.592767 25 0.261799 24.14815 380.904 98.58521 430.7864 474.9799 108.2445 28.01573 145.1635 25 0.261799 24.14815 569.6073 147.4252 668.3297 713.651 157.0845 40.65645 192.1922 25 0.2618 24.1481 748.11 193.625 893.033 939.421 203.284 52.6139 236.679 25 0.261799 24.14815 916.4133 237.1852 1104.896 1152.29 246.8445 63.88805 278.6233 Table 3.2: Thrust and Drag calculations Page 46 of 78 25 0.261799 24.14815 1074.516 278.1052 1303.918 1352.257 287.7645 74.47893 318.0257 25 0.261799 24.14815 1222.419 316.3852 1490.101 1539.324 326.0445 84.38652 354.886 25 0.261799 24.14815 1360.121 352.0252 1663.443 1713.489 361.6845 93.61083 389.2042 25 0.261799 24.14815 1487.623 385.0252 1823.946 1874.754 394.6845 102.1519 420.9804 25 0.261799 24.14815 1604.925 415.3852 1971.608 2023.117 425.0445 110.0096 450.2144 ROTOR CALCULATION Induced Velocity generated(Vi)(m/s) Velocity Parallel to the disc(Vx)(m/s) Vz(m/s) Velocity Normal to the disc(Vz-Vi)(m/s) Overall Inflow Velocity(Vt)(m/s) Rotor mach number(Mach) µ(V/Vt) µx(Vx/Vt) λi(Vx/Vt) µzd(Vz-Vi/Vt) µz(Vz/Vt) Rotor Tilt angle(φ) Coefficient of Thrust(Ct) function of constant Avg thrust Coefficient of Thrust(Ct) function of solidirity,theta_not, a Average Thrust(Tavg) as a function of Ct Cyclic Pitch(B1)(degrees) Blade Lift Coefficient(Cl_average) Profile Drag Coefficent(Cdo) Vd(as a function of function of s,, a) Vd(as a function of constant thrust) Ideal Induced power(for n=2)(W_ Total Rotor Power(W) Figure of Merit 2.030 101.422 27.176 25.146 107.969 0.446 0.973 0.939 0.019 0.233 0.252 6.848 0.030 0.032 59.387 -40.200 2.115 0.724 33.313 30.464 156.435 319.681 0.377 6.193 99.490 26.658 20.465 105.013 0.434 0.981 0.947 0.059 0.195 0.254 5.715 0.032 0.032 55.860 -36.527 2.103 0.740 32.084 32.571 448.875 529.160 0.654 9.406 96.593 25.882 16.476 101.348 0.419 0.987 0.953 0.093 0.163 0.255 4.758 0.034 0.031 51.778 -32.274 2.093 0.764 30.708 35.601 631.914 651.709 0.747 12.989 91.763 24.588 11.599 95.705 0.395 0.993 0.959 0.136 0.121 0.257 3.540 0.038 0.031 45.889 -26.140 2.080 0.806 28.693 41.355 773.370 734.727 0.811 15.647 86.933 23.294 7.647 90.324 0.373 0.996 0.962 0.173 0.085 0.258 2.470 0.043 0.031 40.653 -20.686 2.069 0.851 26.830 48.412 825.339 752.679 0.845 17.800 82.104 22.000 4.200 85.104 0.352 0.999 0.965 0.209 0.049 0.259 1.439 0.049 0.031 35.902 -15.737 2.058 0.900 25.056 57.201 829.164 736.558 0.868 19.615 77.274 20.706 1.091 80.007 0.331 1.000 0.966 0.245 0.014 0.259 0.397 0.055 0.031 31.564 -11.218 2.047 0.955 23.346 68.254 803.306 700.139 0.884 21.178 72.444 19.411 -1.767 75.021 0.310 1.000 0.966 0.282 -0.024 0.259 -0.686 0.062 0.031 27.601 -7.091 2.036 1.014 21.689 82.280 758.444 651.338 0.897 22.543 67.615 18.117 -4.425 70.140 0.290 0.998 0.964 0.321 -0.063 0.258 -1.840 0.071 0.030 23.988 -3.326 2.025 1.080 20.081 100.241 701.619 595.335 0.908 23.743 62.785 16.823 -6.920 65.367 0.270 0.994 0.960 0.363 -0.106 0.257 -3.091 0.082 0.030 20.705 0.093 2.012 1.153 18.519 123.463 637.857 535.791 0.918 24.802 57.956 15.529 -9.273 60.712 0.251 0.988 0.955 0.409 -0.153 0.256 -4.468 0.095 0.030 17.741 3.181 1.998 1.234 17.006 153.774 570.922 475.420 0.926 25.739 53.126 14.235 -11.504 56.190 0.232 0.979 0.945 0.458 -0.205 0.253 -6.007 0.111 0.030 15.084 5.949 1.984 1.324 15.544 193.701 503.728 416.296 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 47 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 48 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 (20mN90mN). 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 49 of 78 Component type Descripti on/ Designati on Mass per piece [kg] [s] BI Astrium 400 N Bi-Propellant Engine BI Astrium 400 N Bi-Propellant Engine BI Astrium 10 N Bi-Propellant Thruster (single Seat Valve) Model S400-12 Model S400-15 Valve (+heater) accumulated Length power burn life (max) Inlet I_sp pressure Thrust per piece per piece [N] [W] [hours] [bar] 3.6 318. 4.3 321. 10. 10. 420. 425. 0.35 291. 10. 35. 35. Diameter Mixing (max) ratio OX / PROP [mm] [mm] 8.3 12.8 70. (1) Monoprop (2) Biprop (3) Cold- Propella V(exhaus gas nt Oxidiser t) 503. 669. 158.5 248. (m/s) 1.65 316. 1.65 103. 1.65 2. MMH (N2O4, MON-1, MON-3) 3119.58 2. MMH (N2O4, MON-1, MON-3) 3149.01 2. MMH (N2O4, MON-1, MON-3) 2854.71 MONO AMPAC In-Space Propulsion MONARC445 1.6 235. 445. 410. n/a 1. N2H4 n/a 2305.35 MONO AMPAC In-Space Propulsion MONARC90 1. 235. 90. 300. n/a 1. N2H4 n/a 2305.35 2.7 214. 400. 0.485 325. n/a 1. N2H4 n/a 2099.34 0.395 224. 20. 10.500 195. n/a 1. N2H4 n/a 2197.44 0.29 220. 1. 50.000 172. 30. n/a 1. N2H4 n/a 2158.2 MONO Astrium 400 N Hydrazine Thruster MONO Astrium 20 N Hydrazine Thruster MONO Astrium 1 N Mono-Propellant Thruster COLD GAS Moog 0.12 N GN2 Thruster COLD GAS Moog 3.5 N GN2 Thruster Solenoid Actuated 58E142A Solenoid Actuated 58-118 0.016 15.9 57. 6.9 0.12 16,666.667 n/a 3. GN2 n/a 559.17 0.022 71.5 14.82 3.5 16,666.667 n/a 3. GN2 n/a 701.415 2. MMH MON 2844.9 BI Astrium 22 N Bipropellant Thruster BI Northrop Grumman Dual Mode Liquid Apogee Engine TR-308 322. BI Northrop Grumman High Performance Dual Mode Liquid Apogee Engine TR-312100YN 330. MONO-85N Northrop Grumman Monopropellant Thruster MRE-15 1.1 228. 0.65 290. 22. 35. 66. 72. 212. 55. 6.719 0. 0. 1. 2. N2H4 N2O4 3158.82 6.944 0. 0. 1.06 2. N2H4 N2O4 3237.3 1. H2H4 n/a 318. 119. n/a Table 4.2: Options for Thrusters and their specifications Page 50 of 78 2236.68 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   Total Fuel Budget in final design( Total flight time = ̇ ̇ ̇ = 0.0281 kg/s = 54 kg ̇ 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: Vstall(m/s) Wing Area(m2) Forward(thrust) Fuel Mass flow rate(kg/s) Burn Time(s)(for Mfuel=54 kg) Range(m) Range in (Km) 120 8.8 70.93 0.031712 1702.816 165389.9 165.3899 105 8.5034014 50.42349 0.0225439 2395.3264 144716.19 144.71619 103 8.8368366 70.934019 0.031714 1702.7193 141959.7 141.9597 100 9.375 100.9602 0.045138 1196.32 137824.9 137.8249 95 90 10.38781 11.5741 149.0318 194.638 0.066631 0.08702 810.4359 620.54 130933.7 124042 130.9337 124.042 85 12.97578 237.7793 0.106309 507.9531 117151.2 117.1512 80 14.64844 278.4552 0.124495 433.7528 110260 110.26 75 16.66667 316.6659 0.141579 381.4137 103368.7 103.3687 70 19.13265 352.4114 0.15756 342.7264 96477.46 96.47746 Table 4.2: Thruster performances for different stall speeds Page 51 of 78 65 22.1893491 385.691747 0.1724394 313.153499 89586.2158 89.5862158 60 26.04166667 416.506844 0.186216555 289.9849588 82694.96839 82.69496839 55 30.99174 444.8567 0.198892 271.5048 75803.72 75.80372 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 52 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 53 of 78 Figure 5.3: Identification of dimensional parameters from concept generation-Glider (distances in metres) [15] Page 54 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 55 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 56 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 57 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 58 of 78 Fig 5.9: Deployment of rotor hinges. Point 30 and Point 32 are the top edges of the rotor hub Page 59 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 60 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 61 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 62 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 63 of 78 Figure 5.16: Dimensional Parameters of final autogiro design Page 64 of 78 Figure 5.17: Panoramic view of final autogiro design Page 65 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 66 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 67 of 78 Table 5.1: Velocity Vs performance parameters for final design WINGED/GLIDER CONFIGURATION Vstall(m/s) Wing Area(m2) Drag Force(N) Lift Force(N) THRUST CALCULATIONS Autorotation Velocity(Vd) Rotor Tilt angle(in radians) Minimum Forward Velocity(m/s) Rotor Thrust required(N) Rotor Drag(N) Rotor Thrust(function Rotor Drag)(N) Forward Propulsion Thrust(N) Rotor Thrust(with constant lift and forward velocity) Rotor Drag(Const Lift)(N) Forward Propulsion Thrust(Const Area Lift) ROTOR CALCULATION Induced Velocity generated(Vi)(m/s) Velocity Parallel to the disc(Vx)(m/s) Vz(m/s) Velocity Normal to the disc(Vz-Vi)(m/s) Overall Inflow Velocity(Vt)(m/s) Rotor mach number(Mach) µ(V/Vt) µx(Vx/Vt) λi(Vx/Vt) µzd(Vz-Vi/Vt) µz(Vz/Vt) Rotor Tilt angle(φ) Coefficient of Thrust(Ct) function of constant Avg thrust Coefficient of Thrust(Ct) function of solidirity,theta_not, a Average Thrust(Tavg) as a function of Ct Cyclic Pitch(B1)(degrees) Blade Lift Coefficient(Cl_average) Profile Drag Coefficent(Cdo) Vd(as a function of function of s,, a) Vd(as a function of constant thrust) Ideal Induced power(for n=2)(W_ Total Rotor Power(W) Figure of Merit Fuel Mass flow rate(kg/s) Burn Time(s)(for Mfuel=54 kg) Range(m) Range in (Km) 92.00 11.30 31.28 446.90 94.00 11.30 32.66 466.54 96.00 11.30 34.06 486.60 98.00 11.30 35.50 507.09 100.00 11.30 36.96 528.00 102.00 11.30 38.45 549.33 104.00 11.30 39.98 571.08 25.00 0.09 24.90 1292.78 116.19 1402.38 1438.43 144.01 12.94 175.87 25.00 0.09 24.90 1074.24 96.55 1164.21 1200.83 124.37 11.18 157.53 25.00 0.09 24.90 851.01 76.49 920.91 958.12 104.30 9.37 138.79 25.00 0.09 24.90 623.07 56.00 672.49 710.29 83.82 7.53 119.65 25.00 0.09 24.90 390.44 35.09 418.94 457.36 62.91 5.65 100.12 25.00 0.09 24.90 153.11 13.76 160.28 199.32 41.58 3.74 80.20 25.00 0.09 24.90 -88.93 -7.99 -103.51 -63.83 19.82 1.78 59.88 13.92 12.75 11.43 9.90 93.62 95.61 97.60 99.60 8.45 8.63 8.81 8.99 -5.47 -4.12 -2.62 -0.91 94.16 96.09 98.04 100.00 0.39 0.40 0.41 0.41 1.00 1.00 1.00 1.00 0.99 1.00 1.00 1.00 0.15 0.13 0.12 0.10 -0.06 -0.04 -0.03 -0.01 0.09 0.09 0.09 0.09 -1.67 -1.23 -0.77 -0.26 0.04 0.04 0.04 0.04 0.03 0.03 0.03 0.03 44.34 46.28 48.28 50.37 -24.53 -26.54 -28.63 -30.81 2.08 2.08 2.09 2.09 0.82 0.80 0.79 0.77 28.15 28.83 29.53 30.25 43.21 40.92 38.80 36.84 800.99 765.58 716.03 647.13 749.40 730.39 700.91 656.97 0.82 0.81 0.79 0.76 0.070428 0.06205 0.053495 0.044763 766.7357 870.2691 1009.445 1206.346 129555.5 132311.9 135068.4 137824.9 129.5555 132.3119 135.0684 137.8249 8.05 101.59 9.17 1.12 102.01 0.42 1.00 1.00 0.08 0.01 0.09 0.31 0.03 0.03 52.56 -33.09 2.10 0.76 31.01 35.02 548.94 590.99 0.72 0.035855 1506.049 140581.4 140.5814 5.56 103.58 9.35 3.79 104.07 0.43 1.00 1.00 0.05 0.04 0.09 1.05 0.03 0.03 54.90 -35.53 2.10 0.75 31.84 33.31 395.96 483.48 0.63 0.026771 2017.095 143337.9 143.3379 14.98 91.63 8.27 -6.71 92.24 0.38 1.00 0.99 0.16 -0.07 0.09 -2.09 0.04 0.03 42.47 -22.57 2.07 0.83 27.47 45.69 825.49 760.39 0.84 0.078631 686.7545 126799 126.799 Page 68 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 69 of 78 Fig 6.2: Descent strategy of X-I based on the Ares entry strategy[7] Page 70 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 71 of 78 . 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 Mass Power (Operating) Data Volume Value 0.5 kg 3W 36 MB/orbit Table .1: Specifications for the Piezo Dust Detector (PDD) Page 72 of 78 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 rockforming 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. Page 73 of 78 Figure .2: CAD image of the C1XS Instrument showing coalligned front detectors, deployable radiation shield and 140◦field-of-view. Parameter Mass Power (Operating) Spatial Resolution Value 5.56 kg 25.5 W 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. Page 74 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. .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 75 of 78 .. 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 76 of 78 CONCLUSION The study done in the report is only a beginning to accomplish a successful rotorwing 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 gliderautogiro-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 Final Observational Plan for X-I height 12 10 10 km 8 Glider from t=0 - t=1875 6 4 autogiro t=108 2 glider from t = 1506 0 0 1.5 km 108 Page 77 of 78 1.5 km 1506.04 0.01 1875.298 time-> Now using the two laws of linear motion: S =Ut +‘Ǥͷ 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. 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