Innovative Low Cost Mars Flyby Spacecraft for Safe ...Mars, the red planet, has always been a matter of interest to astronomers and space enthusiasts. Since Neil Armstrong made his

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Innovative Low Cost Mars Flyby Spacecraft for Safe Interplanetary Human Mission

Aswath Suresh1, Gautam Ranjan2, Sri Harsha3, Kolluri Surya4, Adarsh Ranjan5, Nitin Ajithkumar6,

Sajin Sabu7, Vinay Teja8, Abel Varghese David6, Swapnasheel Sonkamble9, Shreyance Singhvi1 and Ganesha Udupa6

1New York University, Brooklyn, New York, USA

2Binghamton University, 4400 Vestal Parkway East, NY, USA 3San Francisco State University, Holloway Ave, San Francisco, CA 4Institute of Aeronautical Engineering, Hyderabad, Telangana, India

5University of Manchester, Oxford Rd, Manchester M13 9PL, United Kingdom 6Amrita School of Engineering, Kollam, Kerala, India

7College of Engineering, Trivandrum, Kerala, India 8University of South Alabama, Mobile, AL, United States

9M.I.T College of Engineering Pune, India (aswathashh10, gautamranjan91, sriharsha.yerrabelli, ammasganesh,

nitinakumar)@gmail.com

ABSTRACT. The human civilization has reached a point in its progression, where inhabiting planets other than Earth is fast becoming a necessity. Mars is considered the first step in this colonization scheme, due to the large number of findings that have been made that indicate favorable conditions for human survival. This paper describes in detail an innovative, cost efficient and safe mars flyby human mission. The rocket PSLVXL of ISRO is proposed as the launch vehicle due to its cost effectiveness and high success rate. The rocket with its alternate four stage propulsion system will carry a spacecraft having two humans on board and put it into Earth`s elliptical orbit. The spacecraft`s orbit around the earth is raised by firing its liquid engine. After the 6 main engine fire, the spacecraft finally gains escape velocity to travel out of the sphere of influence of the Earth. Then the spacecraft is beamed onto the Inspirational Mars Free Return Trajectory which basically uses Hofmann Transfer. This ensures the usage of least amount of fuel. The spacecraft is equipped with star sensors and a gyroscope for accurately keeping track of the trajectory. Four thrusters and a main liquid engine is used for trajectory correction whenever it is required. Innovative deep sleep mechanism has been incorporated to limit the food and water requirement for the humans on board. The humans will be able to go into deep sleep for approximately 27 days. The spacecraft has been designed in such a way as to sustain two humans for around 600 days. The spacecraft will flyby mars in approximately 253 days and continue its free return trajectory in approximately 273 days reaching back to the Earth’s sphere of influence. Composite material heat shield which can withstand nearly 2500 degree Celsius and a parachute landing system will be used to safely land humans on the Earth’s ocean. The implementation of a mission based on this idea will most definitely be a new chapter in the history of human civilization.

KEYWORDS: PSLV-XL, Star Sensors, Spacecraft, Trajectory, Deep Sleep Mechanism, Humans


1 Introduction Mars, the red planet, has always been a matter of interest to astronomers and space enthusiasts. Since Neil Armstrong made his first steps on Moon, man has dreamed of colonizing new worlds to live in. Mars, being the sister planet to earth due to its size, shape, composition and distance to sun, can be considered our best bet for inter-planetary habitation. From cultural references of Martians and secret societies, to serious scientific discussions on terraforming the red planet, all indicate our enthusiasm to visit the red planet and hopefully making it the next planet that harbors human life. The conditions in Mars are believed to be hospitable since the planet is similar to Earth in many ways. Mars and Earth have almost equal period of revolution around the axis. Mars takes 24 hours and 37 minutes to complete a revolution around its axis. While Earth takes approximately 365 days to orbit around the sun, Mars takes 687 days for an orbit around the sun. The gravity of Mars is roughly one-third of Earth`s gravity and it has a thin atmosphere with a pressure of 1% that of Earth. The atmosphere, water, ice and geology interact with each other to produce a dynamic Martian environment as in Earth. Mars has surface features reminiscent of both the impact craters of the Moon and volcanoes, deserts and polar ice of Earth. It inspires vision of an approachable world. For ages humans have been speculating about life on Mars. But, the question that is to be still answered is whether Mars has a biosphere or ever had an environment in which life could have evolved and sustained. [1] It quite evident from environmental facts and figures that humans have overused and exploited the Earth and it fast reaching a crossover period in history where moving to another planet is no longer a fancy of the imagination, but a crucial necessity of survival. Much damage has already been done to the planet’s environment even while knowing the long term effects. One should embrace the fact that any catastrophe like global warming, nuclear war, asteroid collision can easily wipe out the human race from the planet and in such a catastrophe we may find ourselves having no place to stay as it sure to destroy the home planet. The best way to avoid it such a disastrous possibility would be to move on and establish colonies in other planets. The Kardashev scale (Kardashev, 1964) gives us a scaling relation between different types of developed intelligent civilizations in the universe and their relative energy usage. Robert Zubrin adapted the Kardashev scale to refer to the advancement of civilization by measuring their spread in space. According to him, we are a Type-I civilization who have transcended our regional boundaries to occupy the entire planet and we have all the ideas and technological capabilities to become a Type II civilization which can extensive colonies in our stellar system. All that we see today both as possibilities are as our reality indicate that we are ready to make a transition to Type II civilization by colonizing the red planet. Human settlement of Mars will be the next giant leap for humankind. Exploring the solar system as a united humanity will bring us all closer together. Mars is the stepping stone for our voyage into the universe. Human settlement on Mars will aid our understanding of the origins of the solar system, the origins of life and our place in the universe. We can use this knowledge to advance our understanding of future space explorations and colonizing strategies on other planets, moons and other such heavenly bodies. As with the Apollo Moon landings, a human mission to Mars will inspire generations to believe that all things are possible and that anything can be achieved. [2]


2 Mission Plan Mission planning is done in conjunction with the defined Gemini Mars Mission Objectives. The Mars Mission can be envisaged as a rendezvous problem. The rendezvous mission consists of following steps

i) The spacecraft with two humans on board is proposed to be lift off by India’s very own PSLV-XL using alternate solid and liquid propulsion in 4 stages. (See Section - 3)

ii) The spacecraft (Mars flyby capsule) is separated from PSLV-XL rocket and put into elliptical earth orbit.

iii) As the capsule goes away towards apogee its velocity decreases, but as it comes back its velocity increases because of the earth gravitational pull.

iv) Engine to be fired when velocity is high i.e. when the spacecraft is close to the earth. This raises the orbit and increases the velocity with least consumption of fuel.

v) After six main engine burns, the spacecraft gains escape velocity, to escape the sphere of influence of the Earth and beamed tangentially from Earth’s orbit to inspiration mars trajectory when Mars is in right position. This process of setting the spacecraft in a trajectory, which will cause it to arrive at Mars is called Trans Mars Injection [3].

vi) The spacecraft is equipped with star sensor and gyroscope to accurately follow the trajectory. The spacecraft is powered by solar cells and antenna is used for communication with Earth. Radio waves are continuously sent out to the orbiter for the communication with Earth.

vii) The spacecraft is built in such a way to sustain humans for around 600 days. Innovative deep sleep mechanism has been incorporated with health monitoring to limit the amount of food and water required to be carried. (See Section- 8)

viii) The spacecraft will further continue to stay in motion unless acted upon by external forces. Therefore, fuel will only be used to correct the trajectory. This process is called trajectory correction maneuver [3].

ix) It is expected to flyby mars on 21st August 2018 (approximately after 253 days) and mark the new era of space exploration. To make this happen, the angle between Earth, Mars and Sun should be approximately 44 degrees. Such arrangement occurs at an interval of 780 days.

x) Now the spacecraft continues in the free return of inspiration mars trajectory reaching back to earth sphere of influence by 21st May 2019.

xi) Composite heat shield which can withstand nearly 2500 degree Celsius and a parachute landing system will be used by the both the humans to safely land on the Earth’s ocean.

3 Launch Vehicle A launch vehicle provides the velocity needed by a spacecraft to escape Earth's gravity and set it on its course to Mars [4]. Figure 1 shows the Launch vehicle parts. Launch vehicle, in spaceflight, a rocket-powered vehicle used to transport a spacecraft beyond Earth’s


atmosphere, either into orbit around Earth or to some other destination in outer space. Practical launch vehicles have been used to send manned spacecraft, unmanned space spacecraft, and satellites into space since the 1950s. [5]. Designed to launch in 2018, we can take advantage of a favourable launch opportunity when Earth and Mars are in advantageous positions in their orbits for a Mars landing [6]. That means that it would take less power to get to Mars relative to other times when Earth and Mars are in different positions in their orbits around the sun. When mission planners are considering different launch vehicles, what they take into consideration is how much mass each launch vehicle can lift into space. The Mars 2020 rover mission would use heritage technologies based on the Mars Science Laboratory mission, which was designed to allow large-mass payloads to reach the surface of Mars safely [6]. Launch vehicle selection criteria would encompass the final architecture for this mission, including total mass. For a sense of overall potential mass, the Mars Science Laboratory mission had a total launch mass, including the intermediate class rocket that lifted away from Earth, a weight of about 531,000 kilograms (1.17 million pounds). Every 26 months, Earth, Mars and the Sun align for the most efficient, least energy-consuming path between Earth and Mars. [7]

Figure 1 Launch Vehicle Parts (ISRO)

A. PSLV-XL Vehicle Description

The PSLV has four stages using solid and liquid propulsion systems alternately. The first stage, one of the largest solid rocket motors in the world, carries 138 tonnes of hydroxyl- terminated polybutadiene urethane-bound (HTPB) propellant and develops a maximum thrust of about 4,800 kN. The 2.8-m diameter motor case is made of maraging steel and has an empty mass of 30,200 kg. Pitch and yaw control during first stage flight is provided by the Secondary Injection Thrust Vector Control System (SITVC) [8], which injects an aqueous solution of strontium perchlorate into the nozzle to produce asymmetric thrust. The solution is stored in two cylindrical aluminum tanks strapped on to the solid rocket motor and pressurized using nitrogen. Roll control is provided by two small liquid engines on opposite sides of the stage, the Roll Control Thrusters (RCT) (from figure 2). Stage by stage description of the PSLV [PSLV-XL] is as follows

● First Stage


○ On the PSLV and PSLV-XL, the first stage thrust is augmented by six strap-on solid boosters.

○ Four boosters are ground-lit and the remaining two ignite 25 seconds after launch. ○ In the standard PSLV, each booster carries nine tonnes of propellant and produces 510 kN

thrust. ○ The PSLV-XL uses larger boosters which carry 12 tonnes of propellant and produce 719

kN thrust. ○ Two strap-on boosters are equipped with SITVC for additional attitude control as shown in

figure 1.2. The PSLV-CA uses no strap-on boosters.

Figure 2 PSLV configuration. (ISRO)

● Second Stage ○ The second stage employs the Vikas engine and carries 41.5 tonnes (40 tonnes till C-5

mission) of liquid propellant – unsymmetrical dimethyl hydrazine (UDMH) as fuel and nitrogen tetroxide (N2O4) as oxidizer.

○ It generates a maximum thrust of 800 kN (724 till C-5 mission). ○ The engine is hydraulically gimbaled (±4°) to provide pitch and yaw control, while roll

control is provided by two hot gas reaction control motors.

● Third Stage ○ The third stage uses 7 tonnes of HTPB-based solid propellant and produces a maximum

thrust of 240 kN. ○ It has a Kevlar-polyamide fiber case and a submerged nozzle equipped with a flex-

bearingseal gimbaled nozzle (±2°) thrust-vector engine for pitch & yaw control. ○ Roll control is provided by the fourth stage reaction control system (RCS).


● Fourth Stage ○ The fourth stage is powered by twin engines burning monomethyl hydrazine (MMH) and

mixed oxides of nitrogen (MON). ○ Each engine generates 7.4 kN thrust and is gimbaled (±3°) to provide pitch, yaw & roll

control during powered flight. Coast phase attitude control is provided by RCS. ○ The stage carries 2,500 kg of propellant in the PSLV and PSLV-XL and 2,100 kg in the PSLV-CA.

PSLV is developed with a group of wide-range control units as shown in Table 1. Table 1. PSLV Control Units

Stage 1 Stage 2 Stage 3 Stage 4

Pitch SITVC Engine Gimbal Flex Nozzle Engine Gimbal

Yaw SITVC Engine Gimbal Flex Nozzle Engine Gimbal

Roll RCT and SITVC in 2 PSOMs

HRCM Hot Gas Reaction Control

Motor

PS4 RCS PS4 RCS

ISRO has envisaged a number of variants of PSLV to cater to different mission requirements. There are currently three operational versions of the PSLV — the standard (PSLV), the corealone (PSLV-CA) without the six strap-on booster motors, and the (PSLV-XL) version, which carries more solid fuel in its strap-on motors than the standard version. These configurations provide wide variations in payload capabilities ranging from 3800 kg in LEO to 1800 kg in Sun synchronous orbit.

B. All Variants of PSLV PSLV (Operational)

The standard version of the PSLV has four stages using solid and liquid propulsion systems alternately and six strap-on boosters. It currently has the capability to launch 1,678 kg to 622 km into Sun synchronous orbit.

PSLV-CA (Operational)

The PSLV-CA, CA meaning "Core Alone", model premiered on 23 April 2007. The CA model does not include the six strap-on boosters used by the PSLV standard variant. Two small roll control modules and two first-stage motor control injection tanks were still attached to the side of the first stage. The fourth stage of the CA variant has 400 kg less propellant when compared to its standard version. It currently has capability to launch 1,100 kg to 622 km Sun synchronous orbit.

PSLV-XL (Operational)

PSLV-XL is the updated version of Polar Satellite Launch Vehicle in its standard configuration boosted by more powerful, stretched strap-on boosters. Weighing 320 tonnes at


lift-off, the vehicle uses larger strap-on motors (PSOM-XL) to achieve higher payload capability. PSOMXL uses larger 1-metre diameter, 13.5m length motors, and carries 12 tonnes of solid propellants instead of 9 tonnes used in the earlier configuration of PSLV. On 29 December 2005, ISRO successfully tested the improved version of strap-on booster for the PSLV. The first version of PSLV-XL was the launch of Chandrayaan-1 by PSLV-C11.The payload capability for this variant is 1800 kg compared to 1600 kg for the other variants. Other launches include the RISAT Radar Imaging Satellite and GSAT-12.

C. Launch Vehicle Specification Specifications of the complete Launch Vehicle [10] is as shown in Table 2. Table 2. Launch Vehicle Specification

STAGE-1 PSOM- XL

STAGE-2 STAGE-3 STAGE-4

Propellant Solid Solid Liquid Solid Liquid

(HTPB Based)

(HTPB Based)

(UH25 N2O4)

+ (HTPB Based)

(MMH MON-3)

+

Propellant Mass 138(Tonne) 12.2 42

7.6 2.5

Peak Thrust (kN) 4800 718 799 247 7.3 X 2

Burn Time (sec) 103 50 148 112 525

Diameter (m) 2.8 1 2.8 2.0 2.8

Length (m) 20 12 12.8 3.6 2.7

D. Weight distribution (in kilograms) of the payload

Launch Pod: - 2500 kg Dry Mass: - 500 kg Propellant Mass: - 500 kg Total Payload Mass: - 3500 kg

E. Cost Estimation Launch Vehicle - PSLV XL - �90 crores ($15M) - �108 crore ($16M)


F. Launch Dates The following table 3 shows the possible launch schedule. The schedule includes the detailed timeline from earth launch, reaching mars and landing back to earth [11].

Table 3. Launch Dates

Launch Earth Departure Earth’s orbit

Reach Mars Land On Earth

0 days + 25 to 30 + 228 days + 273 days

Thursday, 7 January 2016

Saturday, 6 February 2016

Wednesday, 21 September 2016

Wednesday, 21 June 2017

Sunday, 25 February 2018

Tuesday, 27 March 2018

Saturday, 10 November 2018

Saturday, 10 August 2019

Wednesday, 15 April 2020

Friday, 15 May 2020

Tuesday, 29 December 2020

Tuesday, 28 September 2021

Saturday, 4 June 2022 Monday, 4 July 2022 Friday, 17 February 2023

Friday, 17 November 2023

Tuesday, 23 July 2024 Thursday, 22 August 2024 Monday, 7 April 2025

Monday, 5 January 2026

4 Planetary positions of the planets during the missions Following shows the position of planets during the course of our mission. Two launch cases has been used [12] i) Mission [25.02.2018 to 10.08.2019] ii) Mission [15.04.2020 to 28.09.2021]

A. Mission [25.02.2018 to 10.08.2019]


Figure 3.1 to 3.4 shows the planetary position at different timeline during the mission [25.02.2018 to 10.08.2019].

Figure 3.1 Planetary position as per date 25:02:2018 AD





B. Mission [15.04.2020 to 28.09.2021]

Figure 4.1 to 4.4 shows the planetary position at different timeline during the mission [15.04.2020 to 28.09.2021].







5 Trajectory Low cost rocket PSLV-XL is used to inject the spacecraft into a parking orbit of the Earth (See Section - 3). As the spacecraft goes away towards apogee, its velocity decreases, but as it’s comes back its velocity increases because of the earth gravitational pull. Engine is fired when velocity is high i.e. when the spacecraft is close to the earth. This raises the orbit and increases the velocity with least amount of fuel. With six main engine burns, the spacecraft is gradually maneuvered into a departure hyperbolic trajectory with which it escapes from the Earth`s Sphere of Influence with Earth’s orbital velocity + ΔV boost as shown in the figure 5. [3] The sphere of influence of earth ends at 918347 km from the surface of the earth beyond which the perturbing force on the orbiter is due to the Sun only. Now the sun is much more massive than any other planets and its gravity dominates the solar system. Gravitational influence of a planet, as compared to the sun is only significant near the planet. Figure 6 shows the Earth-Mars-Earth trajectory (Inspiration Mars Trajectory). Mars being in the right position, the spacecraft, leaves the earth orbit at a tangential path to it as shown in the figure. This described the process of the spacecraft being affixed in a trajectory to mars is referred as Trans Mars Injection [3]. Also, by Newton`s first law of motion, the spacecraft will move along the trajectory without external forces and thereby requiring fuel only to correct its trajectory. This operation is known as trajectory correction maneuver [3]. For the spacecraft to reach Mars successfully, an approximate of 44-degree angle should exist relative to earth, Mars and Sun which happens at an interval of 780 days. Now, the spacecraft will finally flyby mars on 21st August 2018 (which is approximately after 253 days) as shown in figure 8 and who’s trajectory from earth as shown in Figure 7 (Left) marking a new era for space exploration. It takes again a 273 days to return back to earth on 21 May 2019 whose trajectory is as shown in figure 7 (Right).

Figure 5: Showing the spacecraft interchanging the orbit with orbit velocity + ΔV boost to achieve escape velocity. (ISRO)


Figure 6: Showing the Inspiration Mars Trajectory merged to ISRO orbit velocity gain track (Inspiration Mars/ISRO) For a communication link between earth and the orbiter, radio waves are continuously sent out from the spacecraft. Due to the fact that the spacecraft is moving away from the earth, each successive radio wave from earth has to travel further on to reach the spacecraft before being reflected and redetected back to earth. This causes a change in wavelength, giving rise to Doppler Effect which in turn is used to compute the velocity of the orbiter. To reach mars, the spacecraft follows a heliocentric path (orbit around the sun). The shortest distance between Earth and Mars is 54.6 million km. launching to this path and decelerating to catchup with mars speed, required an extremely large amount of fuel. The route which require least amount of fuel is an elliptical orbit of about 680 million km which forms a tangent to the mars and the earth orbit around the sun. This kind of transfer is called Hofmann Transfer.


Figure 7: LEFT: Shows the trajectory took to reach mars and in RIGHT: the return trajectory took to reach back to earth. (Inspiration Mars)

Figure 8. At its closest approach, the Inspiration Mars' mission will pass within 100 miles of the Red Planet's surface. (Inspiration Mars) 6 Spacecraft Design

Detailed Discussion of Spacecraft Design The Spacecraft has been designed using CATIA keeping standards in mind. The overall

dimension of the Spacecraft is 8m x 4m x 3.5m. Figure 9 shows the Spacecraft Design.


Figure 9 Spacecraft Design

Figure 10 Spacecraft Inside View 1

Figure 11 Spacecraft Inside view 2


Figure 12 Spacecraft Section View 1 & 2

The above figure 10-12 shows the inside and sectional view of the space craft. In above figures: -

Figure 13 Spacecraft section View 3 & 4

Basic Cost Analysis Cost analysis has been done on the basis of standard Indian market price of materials and

gadgets. 4 – layer Composites The four layer Composites have been used for designing the spacecraft casing i.e Outer

Casing system. These layers consist of following materials: -

1. Ablating tiles (Outer most layer)


2. Carbon – Carbon composites 3. RXF1 4. Al – Alloy (Inner most layer)

Ablating tiles (Outer most layer) Shielding must be fitted to a spacecraft, such as a manned capsule or the Space Shuttle,

if it is to survive the intense heat generated during reentry. The high heating experienced by a spacecraft when entering the atmosphere is caused by a high-pressure bow shock in front of the vehicle (not, as is sometimes supposed, friction with the air). This strong shock wave is caused by the craft flying at hypersonic speeds, or high supersonic speeds. Hypersonic refers to speeds greater than Mach 5. The shock wave is where the atmosphere is rapidly compressed by a factor of 50 to 100, depending on the speed of the vehicle. Because of this rapid compression the gas gets heated up to temperatures of 6,000 K or more. This hot gas then impinges on the front of the spacecraft, transferring heat to the surface. As the ablative material absorbs heat it changes chemical or physical state and sheds mass, thereby carrying the heat away from the rest of the structure.

Carbon – Carbon composites

Carbon-carbon composites range from simple unidirectional fibre reinforced structures

to complex woven 3-dimensional structures. The variety of carbon fibres and multidirectional weaving techniques now available allow tailoring of C/C composites to meet complex design requirements. By selection of fibre-type, lay-up (or fibre-weave), matrix and composite heat treatment, the properties can be suited to different applications. The strength of carbon–carbon with unidirectional reinforcement fibres is up to 700 MPa. Carbon–carbon materials retain their properties above 2000 °C.

RXF1 This material has been recently developed by NASA scientists which is one of the best

materials available for Heavy Radiation Shield. However, its main limitation is that it starts melting at high temperature. Otherwise it is very light weight and strong and has a tensile strength of about 3 times to that of Aluminum. Therefore, it can be used as heavy radiation shield along with some other heat shield materials.

Al – Alloy (Inner most layer)

This layer has been just used for structural support and other fitment convenience.


The table 4 below shows the detailed parts and structure of the spacecraft with its material

used, cost estimation etc. Table 4 Spacecraft parts details with cost estimation

Sr.No PartName PartImage BaseMaterial

Costin Crore(Ind Rupee/USD)

1 SpacecraftCasing

4-LayerComposite 30/4.48

2 Gyroscope

Al 0.05/0.01

3 Micrometer-Camera

NA 1/0.15

4 Accelerometer

Al 0.05/0.01


5 Antenna

Ag 2/0.3

6 Control

Systemanddataanalysiscenter

NA 20/3

7 Drive

MechanismAccuator

Fe 0.05/0.01

8 DriveMechanismdisk

Fe 0.05/0.01


9 Infraredcamera

NA 2/0.3

10 LigtningCamera

NA 3/0.45

11 Hibernation

Cabine(fortwoastronauts)

NA 40/6

12 Casing

5-LayerComposite 5/0.75

13 SolarPanel

Si 4/0.6


Other expenses except the cost shown above in table has been considered around 25 Cr. It includes the fitment, assembly and other costs.

So, Total Cost = 107.2 + 25 =132.2 Crore Indian Rupee= 19.73 Million USD Detailed Discussion A. Spacecraft Casing: -This is the most important part of the spacecraft as it protects

from heavy radiation and high temperature faced by the spacecraft during its journey from earth to mars.

B. Gyroscope: -It helps in the proper navigation of the spacecraft with all its balanced

movement. C. Micrometer- Camera: - Micrometer makes high-speed measurements of fast moving

targets on production and assembly lines via high-intensity, low-noise GaN LED light source. D. Accelerometer: -It measures the acceleration of the spacecraft and the sensors acts on

the basis of its inputs. E. High Gain Antenna and Communication: - Communication management at

distance from 214 million (Range at Mars capture) to 375 million Km (Range of Mars after 6 month) are done using S-Band deep space frequencies for both TTC and Data transfer with two 230 W TWTAs and two coherent transponders. S-Band Δ- DOR transmitters for improving the orbit determination accuracy. [3]

F. Control System and data analysis center: - This cabin has all the setups which are responsible for controlling everything going inside

and outside of the spacecraft. It receives data from outside and inside and processes it to plan some decision.

G. Drive Mechanism Actuator: - This drive mechanism is responsible for transferring power from source to the load. It

requires consideration of motion, load, power requirements and the placement of the actuators with respect to the joints. The primary considerations in transmission are stiffness and efficiency, which have taken care of.

H. Infrared camera: - It is an imaging camera, designed to detect light at near- and

midinfrared wavelengths - in other words, light with wavelengths between 3.6 and 8.0 microns (1 micron is one-millionth of a meter).

I. Solar Array: - Two 850W Solar Panel Powers the spaceship with 1700W.

J. Hibernation Cabin: - Using the proposed deep sleep mechanism as shown in figure 14.

People can go into deep sleep for about 27 days which can limit the food and water requirement


(See Section 8)

Figure 14: Deep Sleep Hibernation Cabin Setup

7 Astronaut Health and Safety Measure “A better mission is with better people. Healthy astronauts lead to a successful mission.”

- Anonymous team member

Space is unpredictable. But can regulate known circumstances that causes trouble to astronauts during the mission. This will help avoid mishaps. Few predictable problems for astronauts during space voyage are given as under-

Astronaut during space travel in deep sleep will experience

A. Muscle and bone Atrophy B. Body Toxic Management C. Cardiac Control D. Blood Control E. Brain Control F. Psychological Health.

A. Muscle and bone Atrophy

The pressure exerted on the neck, collar, ankles, the pelvis and vertebrae bones are due to gravity acting upon human body helps to the body maintain the bone density even without proper and regular exercising [16]. But zero gravity in the absence of space gravity causes many problems including muscle and bone density loss. In the weightless environment during the journey to Mars, bone loss poses as a serious threat. Lack of support by gravity to astronaut’s body results in a loss of calcium and hence bones heal slowly and are weaker [17]. Bone density deterioration is also caused due to other factors like low light effect in spacecraft, no direct exposure to sunlight. Although bone density varies from person to person, on an average 1 to 2 % per month. For a trip of one year, this loss might be from 6 to 24% which would later cause problems when on Mars or even worse cause fractures easily


during adverse conditions. According to several journals this reduction would lead to 40 - 60% decrease in load bearing capacity of an astronaut. In general, for a space voyage, bone density is maintained by regular exercising and generating an artificial gravity in space by spinning the spacecraft and creating a centrifugal force within the system to let astronauts place their feet on base for better workouts. Alternatively, exercising treatment such as low-frequency vibrations by mechanical devices, creating vibrations with low-frequency oscillations at around 30-95 Hz for few minutes a day would maintain the muscle and bone density at the same level of strength. Vibrations at high frequencies are harmful but low-frequency vibrations are risk-free and easy to workout when astronauts are in deep sleep. [17] Bone-enhancing drug, OPG (Osteoprotegerin), developed by the NASA-sponsored commercial agency is a protein that prevents bone loss and increases the metabolic rate of bones and muscles.

B. Body Toxic Management The levels of bacteria increase moderately with an increase in the duration of the mission while the levels of fungi decrease with an increase in the duration of the mission. The levels of bacteria range from a few hundred to 1,000 colony-forming units per cubic meter (CFU/m3) of air during longer missions. Fifteen species of bacteria were recovered from samples collected during space missions. Fungi tended to be present at a few hundred CFU/m3 early in the missions, but their quantities dropped to undetectable levels toward the ends of the missions. Nevertheless, low levels of Aspergillus and Penicillium species are found during more than 60 percent of the missions. [18]

C. Cardiac Control Loss of muscle mass also happens to the heart which could lead to cardiac problems. Due to long-time exposure to space, astronaut’s heart become more spherical in shape. According to latest studies, an astronaut’s heart becomes 9.4% more spherical in shape.

Even though the heart regains its original shape after returning to earth, effects of being in a space for a long time are known. Several other parameters were measured like an increase or decrease in arterial pressure, heart beat etc. and it was found that these changes could possibly cause cardiac problems in astronauts. They concluded that distension of the heart and associated central vessels during 0-g might induce the hypotensive effects through peripheral vasodilatation.

Some of the cardiovascular symptoms and abnormalities that astronauts may have during a long-duration space mission include high blood pressure with or without symptoms; atrial and ventricular premature beats; atrial arrhythmias, such as atrial fibrillation, atrial flutter, and supraventricular tachycardia; sustained and nonsustained ventricular fibrillation; chest pain, ischemic and nonischemic; shortness of breath, cardiac and noncardiac; orthostatic hypotension; syncope; vasovagal and other cardio neurogenic responses; and edema, cardiac and noncardiac[20].


Figure 15: Predicted change in a heart shape at end-diastole on Earth (green) and in microgravity (red). Credit: Dr. Chris May [19]

D. Blood Control

Earth’s gravity influences everything inside our body. When we stand straight blood flows into the body parts which are below the heart, when we raise any of the body part above the head, the veins get emptied when limbs regain their position veins are filled with blood and this is a different case in case of space blood flows from the brain down to the heart in the absence of gravity. When an astronaut is in space the heart pumps equal blood throughout the body. We have sensors in our body to measure fluids, so our upper body sensors start reporting our increase. The kidney’s start removing fluids so our brain thinks we are thirsty, so we go into hydration mode. [20] It’s easier for the blood to pump in space because of the lack of gravity, so the heart starts to shrink. Since there is no gravity the blood doesn’t get pulled down to the lower part of the body. Blood goes to the chest and head a lot resulting in puffy faces and a redder complexion. The problem is that there is not enough blood that goes to the brain so when they return to earth they feel dizzy because of the lack of blood flow. With this understanding of blood flow in space, it is possible that researchers can help treat headaches and other neurological symptoms reported by crew members while living aboard the space station. [21]

E. Brain Control Neuro mapping with the help of MRI for astronauts let us understand how space flight impacts astronauts’ central nervous system as shown in figure 16. It causes imbalance due to spatial orientation, figure out improper posture and thinks and confuses between the up and down orientation of the body. The balance system composed of gravity sensors in the inner


ear and the connections between the brain. the eyes, cerebellum, and postural muscles help to work against gravity every day on Earth. Once in space, this balance is not the same anymore. The brain could have increased sensitivity. When the organs are unbalanced, it leads to different sensations for the astronauts. [22]

Figure 16 Neuro Mapping with the help of MRI for Astronauts

Day and night changes, weightlessness, cosmic rays etc. mess up functional tasks for

the brain resulting in problems ranging from simple boredom and fatigue to acute stress reactions, profound depression, and overt psychosis. Sleep loss, work overload, altered immune response also causes imbalance in brain impulses.

F. Psychological Health Astronauts should be able survive in a wide variety of conditions in a space mission, either expected or unexpected. Problems like a mineral loss, bone density loss, and other issues and challenges are directly related to astronaut`s psychological problems. Anecdotal and empirical evidence from NASA and external studies indicate that the likelihood that a behavioural concern or psychiatric disorder will occur during a mission increases with mission length. The Behavioural Health and Performance element focus on the challenges of keeping a crew productive and healthy in an isolated and confined environment, where day and night cycles differ from the standard 24-hour Earth time, workloads can be heavy, and conflicts may arise. The Element conducts research in various analogy and flight environments to develop tools and technologies for treating symptoms related to stress, lack of sleep, inadequate team cohesion, and crew performance. The end result is to optimize the adaptation of the individual and crew to the space environment and maintain motivation, morale, productivity, cohesion, and communication [22].

Scientists working with the Behavioural Health and Performance Element develop tools and technologies to monitor crew activities, performance, and behavioural health. They also develop countermeasures to mitigate problems related to stress, lack of sleep, team cohesion, or crew performance. They focus on strategies such as education and training, self-


assessment tools, lighting exposure for alertness/fatigue, models to predict performance, individualized strategies for medication use, team training and cohesion measures, early detection and intervention technologies, and unobtrusive monitoring tools that can be used by astronauts and flight surgeons for self-assessment purposes. The end result is to optimize the health and performance of the crew, as well as the adaptation of the individual to the space environment. The Element also focuses on post- flight re-adaptation to terrestrial lifestyle and activities countermeasures [23].

Isolation from earth causes behavioural issues and reduces their performance in the mission. Spacecraft noise levels disrupt sleep, increase stress and tension, and can result in temporary or even permanent hearing loss. Control system generates most of the noise The design limits of most work environments range from 63 to 68 decibels (dBA). These noise levels are far greater than noise levels generated on earth. Sometimes such noises lead to hearing loss [22].

G. Heat or Heavy Radiation Shield

When a Shuttle enters outer space atmosphere it is exposed to heavy radiation. The heavy weight of the space shuttle imposes more effect and damages it. To overcome these problems NASA implemented many techniques that could withstand high-temperature environments. They produced a thermal protection layer which is reusable, light weight and low cost. When Space Shuttle re-enters shock layer provides potential for additional heating. The surface of vehicle collides with air molecules as they break apart. Interfacing material/coating have a low propensity to augment the reaction as surface acted as a catalyst. Glass materials have low capacity and allowed the surface of the Orbiter to reject a majority of the chemical energy. Nasa was influenced by Atomic recombination, arc jet measurements effected over a range of surface temperatures for both nitrogen and oxygen recombination. This improved more confidence in Thermal Protection System. Many thermal protection system materials have been used on the orbiter tiles, advanced flexible reusable surface insulation, reinforced carbon-carbon, and flexible reusable surface insulation. These materials used high emissivity coatings. NASA also used silica tiles or fibrous insulation but these cannot withstand very high temperatures [24]. For optimum shielding for space crafts, the major consideration is interaction between the nuclear and atomic interaction and average distance of penetration for a cosmic particle to make it stop before reaching the spacecraft interior parts. Present usage is Aluminium (ρ = 2.7), whereas polyethylene is (ρ=0.96), thickness of shielding increase as payload to carry also increases [25].


Figure 17: Cosmic ray fragmenting Nucleus, releasing secondary radiation

Fig 18: A repeating molecule of polyethylene. The bonding angles are at 110° since the carbon atom is

tetrahedral.

NASA scientists have newly invented a polyethylene- based composite material called RXF1 [26], which is stronger than aluminium (3 times the tensile strength) and lighter than aluminium (2.6 times lighter), is 50% better at shielding solar flares and 15% better at shielding cosmic rays. RXF1 can even deflect micro meteorites. This material is flexible fabric, easy to mould, cut and vacuum pumped, compressed and autoclaved while using in making body of spacecraft. Research is going on for implementing this kind of materials, even this material can be used to shield astronauts and crew quarters. Dr. Cucinotta and his NASA colleagues while experimenting differences between aluminium and polyethylene materials they no significance difference. So using RXF1 can be implemented for long journeys like Mars missions [26]. The drawback is that polyethylene is very flammable and may melt in direct sunlight for long periods. Research is on-going to resolve these issues. In this present scenario, as we have limited data on astronauts condition who underwent through zero gravity for a long period. Analyzing different health conditions and reducing its effects which occurs during space voyage is impossible. Instead, these health effects on astronauts can be reduced or prevented by allowing astronauts into a deep sleep during space travel.


8 Deep Sleep Mechanism-Hibernation System

It’s a long journey to Mars, therefore, we propose a solution to keep the crew in an inactive state during the transit between Earth and Mars. Placing the crew in an inactive state reduces the cost of habitat supplies such as Food, water etc. – reducing its total volume, and also eliminates significant psycho-social aspects. The idea of deep sleep is often posted in the science fiction, where long-duration voyages are completed successfully. Although means for a full cryo-preservation and restoration remains a long way off still. However, recent medical progress has successfully advanced ability to induce in deep sleep states, also known as “torpor”, which induces reduced metabolism rates in humans over extended periods of time. The proposed system combines two common medical procedures to induce and maintain the crew in torpor state. In medicine, to induce torpor in humans, we do it through Therapeutic Hypothermia (TH) – a medical treatment in which patient’s body temperature is lowered to 32°C to 34°C (89°F to 93°F) slowing down the body’s metabolism rate. While being in the torpor state, the passengers can be fed intravenously by nutritional fluids delivered via a central venous catheter and can be monitored with a robotic hand and body sensors. The process of providing the passenger with the required nutrients intravenously is also called as Total Parenteral Nutrition (TPN). Therefore, with the combination of TH and TPN, a group of astronauts can be kept in torpor for the long duration journey required for Earth-to-Mars transit [28].

8.1 Advantages of Torpor

The crew, in an inactive hibernation state, has many advantages as compared to a traditional settlement-class mission. With the crew in hibernation, there is a drastic decrease in the total pressurized volume required for habitation and living quarters required. In addition to this, ancillary crew accommodation (e.g. food galley, cooking and eating supplies, exercise equipment, etc.) can be eliminated. Moreover, a person in torpor state has a reduced metabolism rate and, therefore, requires less amount of consumable food, water, and oxygen [27]. Once the vehicle leaves the Earth’s orbit, the crew can go in a deep sleep through the 4 – 6 months of transit to Mars, and be only woken once they have reached in Mars orbit. In addition to this, having the space crew inactive and stationary during the mission will also provide flexibility in designing the habitat that can help solve or lessen a number of health issues curbing in long-duration spaceflight. The elimination of these risks can be achieved by rotating the habitat element to induce a low-gravity environment inside the crew cabin. This concept of creating an artificial gravity environment isn’t new – spinning habitats have been considered from the earliest days of human space exploration, and have been featured in engineering design as well. In contrast, for an active crew, this often means, creating a habitat where in the crew can ‘stand up’ in the induced gravity field. Whereas, with a stationary crew, the crew can be ‘lying down’ in the induced gravity thus decreasing the size requirements of the habitat. Another design constraint will be with the building of shielding in the capsule. For an active crew, shielding will be obligatory for the entire habitable volume whereas, better protection for radiation can be built for the stationary crew in the voyage of the mission. This can be done by placing additional radiation shielding wherever necessary [30].


8.2 MEDICAL PERSPECTIVE 8.2.1 Artificially Induced Hibernation Black bears (Ursus americanus) are the most famous example of this type of hibernation. There is a minor decrease in body temperature and physiological function in black bears during hibernation. Thus, they are able to fully arouse over a very short period of time. Three possible approaches for inducing a hibernation, in animals and in humans, have been studied over recent years and are summarized as follows: -

1. Temperature-based. Torpor is achieved by lowering the body core temperature through either conductive cooling (e.g. evaporative gasses in the nasal and oral cavity) or invasive cooling (infusing cooled IV fluids).

2. Chemical/Drug-based. In 2011, Scientists at the University of Alaska successfully induced hibernation by activating adenosine receptors (5’-AMP) in arctic ground squirrels. [31]

3. Brain Synaptic-based. Researches shows significant decreases in the number of dendritic spines along the whole passage of apical dendrites in hibernating creatures. Whether it was an initiating factor for hibernation or result of another metabolic process is still under investigation [32-33].

Temperature-based cooling approaches are the highest medical data available for human hibernation and are also well-understood procedure. While this method will be our approach for inducing torpor, supplemented or enhanced technique in the near future shall be most welcomed. 8.2.2 Therapeutic Hypothermia Definition Therapeutic Hypothermia (TH) is a medical treatment that lowers a patient's body temperature in order to help reduce the risk of ischemic injury to tissue following a period of insufficient blood flow. [37] In simple terms there are four aspects to the hypothermic process and stasis induction: initial body cooling, sedation, nutrition/hydration, and rewarming. Patients are actively cooled to a mild hypothermic state (defined as a core temperature between 32 to 34 °C (89 to 93 °F). TH is primarily used as a staple of Critical Care for newborn infants suffering from fatal hypoxia and for adults suffering from head trauma, stroke, cardiac arrest and neurological injuries. It is a standard procedure used in most of the hospitals for treating patients. [37]


8.2.3 Active Research Chinese studies showed evidence of torpor with increased benefits prolonged TH (up to 14days) without a linear increase in the risk of complication [29]. Recent studies also confirm the data [34]. In addition, the U.S. military is also actively funding this research as to use TH to support the war fighter - with a goal of increasing the time period to transfer an injured militant and person to receive apposite medical attention.

Unfortunately, none of the aforementioned efforts are focused on achieving extended durations or considering applicability to human space flight. 8.3 Mission Requirements While Therapeutic Hypothermia is currently used as a short term medical treatment, the current 14-day period can be extended for a longer period within 10-20 years based on the rapid progress, understanding, and extension of this process. Though longer-term stasis is not without its own difficulties and challenges, the current complications associated with hypothermia therapy may stem from the fact that this therapy is being used as a treatment for people with severe medical complications (e.g. shock, compromised immune systems, heart failure, traumatic injuries, etc.). Due to the current lack of need or rationale for medical treatments to maintain therapeutic hypothermia beyond 10-14 days, longer periods have not been attempted. Therefore, due to the current lack of need for medical treatments to maintain therapeutic hypothermia beyond 10-14 days, longer periods have not been attempted.

Figure 19 Current TH vs Expected for Mars voyage (SpaceWork)


8.4 Total Parenteral Nutrition (TPN) A central venous catheter (CVC) is an indwelling intravenous device that is inserted into a vein of the central vasculature. The tip of the CVC usually rests in the Cavo-Atrial Junction (CAJ). This junction marks the inferior end of the superior vena cava (SVC), the continuation below that point being considered part of the heart. All CVCs placed for the purpose of venous access and being inserted in the upper body will ideally have the tip placed within the superior vena cava at or just above the Cavo-Atrial junction. [37] In ‘Total Parenteral Nutrition’ or TPN, all the nutrition and hydration needs for a person can be provided by a liquid solution and administered through an intravenous (IV) line directly into the body. This aqueous solution contains all nutrients that the body needs to maintain full physiologic function. The solution is fed slowly through an aqueous solution containing all the nutrients that the body will need to maintain full physiologic function will be provided to the crew. This solution will be slowly fed through a permanent IV line in the body over a period of hours. This can also be used in numerous post-surgery and oncological treatments where the individual has digestive issues or cannot process foods normally. Long-term Total Parenteral Nutrition is often used to treat people suffering consequences of an accident, surgery, or digestive disorder. Cancer patients and preterm infants are routinely on TPN for months at a time. Long-term parenteral nutrition requires a tunneled central venous catheter or a peripherally inserted central catheter (PICC). A tunneled catheter is preferable, since infections are more common among patients receiving parenteral nutrition at home through a PICC [36].

Figure 20 Chest TPN CVC[37] Figure 21 Arm TPN CVC [37]

The Tunneled Catheters: - A tunneled catheter is a long-term catheter (lasting months to years) that exits the skin via a

subcutaneous tunnel. A Dacron cuff on the tunneled portion of the catheter facilitates anchoring of the catheter through granulation and acts as a barrier to infection. Tunneled catheters may be single, double, or triple lumen. Examples of Tunneled Catheters are Hickmans®, Broviac® and permanent hemodialysis catheters (example. Perm-Cath®). [37]


Figure 22 Tunneled Catheters [37]

Using the Harris-Benedict Equation (HBE) with additional adjustments for individual activity level (e.g. active, resting) and stress environment (e.g. recovering) we can have a standard protocol for calculating the dosage. The HBE estimates the basal metabolic rate and daily kilocalories needed and is a function of gender, body weight, and body height. As the TPN requirements are based on the gender of the crew member, the figure 23 shows required TPN mass per person per day based on a male weighing 80kg and female crew slightly low. The torpor values of “likely” and “potential” represent further reductions in daily TPN requirements from the resting rate due to the lower metabolic rate that is expected in the stasis condition.

Figure 23 Dose rate vs Activity Level [27]

Uses of the TPN system: - • Administer intravenous fluids and blood products • Administer medications • Administer hypertonic solutions (Total Parental Nutrition [TPN]), vesicants (i.e.


chemotherapy), irritants (i.e. cloxacillin), and solutions with extreme pH values (i.e. vancomycin).

• Obtain venous blood samples • Provide long term intravenous therapy • Administer large volumes of intravenous fluid quickly • Administer vasopressor or vasodilator therapy (e.g. Dopamine) • Monitor central venous pressure (CVP) • Provide access for transvenous pacemaker or pulmonary artery catheters • Access venous circulation when a patient has difficult or impossible peripheral access

· Provide hemodialysis access [27]

8.5 CREW TORPOR SUPPORT SYSTEMS Body Thermal Management Systems Torpor or stasis state is achieved by decreasing a crewmember’s core temperature to approximately 34°C (93 °F). In order to achieve this state, systems designed should be low mass, low power, and can be automated. Therefore, regardless of the technique, system characteristics must be maintained. The trans-nasal evaporative cooling system can be used in place of a prolonged cooling pads or cooling major vein of the crewmember. For this system, a cannula (small plastic tube) must be inserted into the crewmember’s nasal cavity and deliver a spray of coolant mist that will evaporate directly underneath the brain and base of the skull. As the flow of blood passes through the cooling area, it consecutively reduces the temperature of the rest of the body. Such an example of a TRANS-nasal evaporative system is the RhinoChill™ system created by BeneChill [27]. Implementation The required crew support systems are identified and demonstrated in Figure 24. The basic features and systems required to support an individual would include:

1. Thermal warming pads to act as an additional thermoregulator and to provide

emergency waking support if needed. 2. An ECG system for monitoring and evaluating heart functions and vital signs. 3. A tunneled catheter for IV hydration, TPN administration, and lab draws. A second

line through arm or thigh could be placed as a reserve in case of infection or damage to the other line.

4. Nasal thermal management system for TH if the evaporative approach is used. (example- RhinoChill).

5. Urine collection assembly and drain line.

Additionally, some loose-fit straps and bindings are used to minimize any movement in the habitat and keep the crew member in their respective nook.


Figure 24 Crew Support System

1. Thermal pads (warming) controller 2. TPN administrator via chest 3. Alternate TPN administrator 4. Thermal management system inserted through nasal cavity (coolant) 5. Urine collection assembly and drain line

9 Apollo Parachute Landing System The Apollo Parachute Landing System is an extensively engineered, thoroughly tested and complex landing system developed over a span of eight years. The parachute system has the task of stabilizing and decelerating the Apollo spacecraft on the completion of its successful mission to a descent velocity safe enough for a water landing. A system engineering approach was utilized throughout the design of the Parachute Landing system. Extensive testing done on the system, along with the utilization of reliability and systems analysis, made sure that the Apollo spacecraft safely returned the crew back to Earth after a successful mission. [14] The testing methods included unmanned missions, ground tests and airborne program which were


done to test the reliability and safety of the designs, which in turn will ensure the safety of the crew. [13]

The system approach initially started with the ground rule that no single component failure should end up in loss of life or result in a failed mission. It was an arbitrary concept which was replaced by a probabilistic approach. Different sets of function failures that could occur simultaneously and in series were identified, to which a probability approach was applied on each one of them. The failures that had very low probability of occurring were ruled out and those which had a probability of occurrence more than a significant value were thoroughly tested and appropriate design changes were made wherever required. The parachute system thus designed was able to cope with a lot of potential failures without the mission being jeopardized. [14]

6.1 . Parachute System

The final parachute system selected for the Apollo command module is shown in Figure 25. The Apollo Parachute landing system consisted of three main parachutes, three pilot parachutes, two drogue parachutes, a forward heat shield augmentation parachute and related electronic, mechanical and pyrotechnic systems to perform the operations. The recovery sequence was automated with the use of barometric switches, which closed at required atmospheric pressures. There was also an option by which crew could manually initiate the recovery systems using time delay relays. The two drogue parachutes are present for initial stabilization and deceleration. Only one drogue parachute is actually required for the task, but the deployment of the both the drogue chutes in tandem will eliminate the need for an emergency sensor to detect the failure of one of the chutes. The use of an additional chute also helps in faster stabilization of the command module and also helps in the attainment of more favorable main chute conditions. The disconnection of drogue chutes is followed by three pilot chute deployed as main chutes. Only two main chutes are required to attain a rate of descent that favors safe landing, but an additional main chute is provided as a safety measure. Moreover, the additional main parachute eliminated the need for error sensors and also helped in establishing more favorable landing conditions. Detailed testing and analysis found that a fourth main chute was not required to attain the required safety standards. [13-14]

Figure 25 – The Apollo Parachute System [14]


6.2 Functional Description

The Parachute Landing system is aimed at stabilizing and decelerating the Command Module to conditions that are safe for landing for manned missions. The Command Module is installed with 9 parachutes, all of which are utilized during some part of the landing sequence. The parachutes included in the Earth landing system are three main parachutes, three pilot parachutes, two drogue parachutes and a forward heat shield separation augmentation parachute.

The recovery sequence is initiated automatically through the closure of barometric switches or through the function of time-delay relays. A logic diagram illustrating each of the ELS functions is presented in figure 26.

Figure 16- Logic and Redundancy in the Earth Landing System [13,15]

The Apollo parachute recovery sequence is illustrated in Figure 27. The deployment

sequence on a nominal landing starts at 21000 feet (7315m) with the separation of the forward heat shield. A 2.2m diameter parachute is mortar deployed from the forward heat shield immediately. [15] This creates a force that pulls the heat shield away from the Command Module so that a collision is prevented. Two 5m diameter conical ribbon drogue parachutes are mortar deployed after a delay of 1.6 seconds from the ejection of the forward heat shield. The drogue chutes remain reefed for the next ten seconds, after which they are fully deployed and they remain attached to the Command Module until an altitude of 11000 feet (3352m) is reached. At this altitude, the drogue chutes are jettisoned and three 2.2m diameter ring sail type pilot parachutes are mortar deployed. These pilot chutes supply the adequate force required to pull the main parachute pack from its stowage location. The main parachutes are of 25.4m diameter and they are reefed in two stages to a fully open position. The first stage occurs 6 seconds after extraction and 10 seconds later the second stage occurs at approximately 9000 feet (2743m). Upon impact with the water, the main parachutes are released enabling the vehicle to float in the water. Impact velocity with all the three main parachutes in good condition is 9.8 m/s. [13-15]


Figure 27 – Apollo Parachute Sequence [15]

6.3. PARACHUTE SYSTEM DETAILS

Several novel design approaches were made in the design of the Apollo Parachute System. The requirement of independent parachute deployment, coupled with large command module oscillations, which in turn creates the possibility of contact between the parachute risers and the hot rear heat shield, and the increase in the weight of the command module weight without any increase in compartment volume or allowable parachute cluster loads resulted in various innovative design features in parachute packing, storage and shape retention.

6.3.1 Drogue Parachute Mortar Assembly

The drogue parachute risers end in steel cables which are attached to the Command Module by means of the flower-pot. The steel cables prevent riser damage due to impact with hot heat shield when the command module oscillates. The setup of the two drogue parachute mortars in one of the four parachute compartment bays is shown in Figure 28. [14]


Figure 28 – Drogue Parachute Bay [14]

The drogue parachute mortar assembly used in the Apollo parachute system, which incorporates a steel cable storage and cartridge orifice design, is shown in Figure 29. The riser ends are secured to prevent riser kinking and the risers are coiled under tension without twisting the ends, and the risers are cast in urethane foam. The light foam disintegrates upon deployment and the risers stretch without kinking by releasing the tension they were wound with.

Figure 29- Drogue parachute mortar assembly [14]


The mortar arrangement of Apollo Command Module is shown in Figure 30. The mortar is

fired from the tumbling Command Module. Dual mortars are used for ensuring reliability, faster command module stabilization and for removing the need for error sensing mechanisms for parachute failure.

Figure 30- Mortar Arrangement of Apollo Command Module [14]

The steel cables are surrounded by lead tubing which is aimed at minimizing the abrasion

between Titanium flower pot collar and the cables, which is the resultant of the steel cables bending and rolling over the flower pot during drogue parachute deployment. This protective system for steel cables is shown in Figure 31. [14]

Figure 31- Steel Cable Protection Using Lead Tubing [14]


6.3.2. Mortar Orifice Design

Figure 32- Mortar Orifice Details [14]

It was observed that increased reaction forces produced upon mortar firing due to an

increase in command module weight and the resultant increase in drogue parachute size and weight could not be tolerated by the Command Module. To solve this problem, an eroding hybrid orifice was designed which maintained the required muzzle velocity for the heavier parachute without increasing the reaction loads of the powerful cartridges. Figure 32 shows the design and pressure characteristics of hybrid and standard orifices. The design of a standard orifice is shown in the upper left hand side of Figure 32. It can be observed from the pressure time plot that the pressure varies a lot with time in the standard orifice’s case. The design of the hybrid orifice is also shown in Figure 32. It has aluminum and brass inserts which burn away progressively and allows more gas to enter the mortar tube. The temperature level of the expanding gas is kept high using the aluminum inserts. From the pressure time graph of the hybrid orifice, it can be noticed that a reasonably constant mortar tube pressure is maintained and the mortar reaction loads are maintained within allowable limits. [14]

6.3.3 Main Parachute Reefing System

The main parachute reefing system in the Apollo parachute landing system was a two stage process. The main parachute reefing system employed in the Apollo parachute landing system is shown in Figure 33. The first stage makes use of two reefing lines, with each line having its individual set of two reefing cutters. This made sure that premature severance of one line will not disreef. The parachute and the separation of both lines are required to disreef the parachute into the second stage. The second stage has one reefing line with two cutters, and it has been noticed through considerable testing and analysis that the rupture of the second stage reefing line will not result in the destruction of the parachute.


Mid-gore reefing was used on all parachutes with cutters attached to canopy radials. Both the first stage lines passed through the same ring opening, called “Siamese Rings”. The slack part of the second stage reefing line is gathered with a drawstring of the same length as the first stage, thus avoiding problems of storing and securing the second stage line. [14]

Figure 33- Main parachute reefing [14]

6.3.4 Main Parachute Retention System

Figure 34- Main Parachute Retention System [14]

The main parachute deployment bags are kept in a truncated cone segment and are stored in

a highly compressed form throughout the mission and storage with a specific gap between apex cover and packed parachute, in order to protect from the intense heat during reentry. The parachute packing density was maintained at 0.0245 lb. /cubic inch. The parachute bag is


packed in the truncated cone form for one year without growth by combined pressure and vacuum packing and storage of the bag in a wooden compartment former. Daisy chain retainers are used to restrain the bags in the CM parachute bays on three sides. There are no intermittent flaps in the connection between the retention system and the deployment bag. The main parachute retention system is shown in Figure 34. [14]

7 Ground Segment

Table 5 shows the ground segments proposed to communicate with the spacecraft. Table 5 Ground Segment [3] Features Specification

S-bandsea-borneterminals LocatedinPacificOceantomonitorPS4andsatelliteseparation.

JPL(JetPropulsionLaboratory)DeepSpaceNetwork

LocatedatCanberra,Goldstone,andMadridwith34m&70m(200kW)antenna,uplinkingat20KWpower.

ISTRACGroundNetworks TTCOperationsduringGeocentricphase.

IndianDeepSpaceNetwork(IDSN) LocatedinBangalorewith18m&32mantenna,uplinkingat20KWpower.

IndiaSpaceScienceDataCentre(ISSDC) Forarchivinganddisseminationofpayloaddata.

8 Mission Cost Estimation

Table 6 shows detailed segment cost estimation required to successfully complete the

mission. Approximately total required budget comes around 82.09 million USD. Table 6 Mission Cost Estimation

SLNO SystemDescription CostinIndianCrores(Rs) CostinUSAMillions(USD)

1 SpaceSegment 315.00 47.01 2 GroundSegment 70.00 10.45 3 ProjectManagement/Contingency 45.00 6.72 4 ProgramElements 10.00 1.49 5 LaunchCost(PSLV-XL) 110.00 16.42 6 Landing 150.00 22.35 Total 700croresRs 104.44millionUSD


Conclusion The evolution of human civilization has reached a stage that colonizing other planets is imperative to its survival. Colonizing Mars is the first and possibly the most obvious stem in the scheme. The paper has described at length about the innovative, cost efficient and safe mars flyby human mission that it proposes. Various aspects of the mission have been covered in detail starting from the mission sequence to trajectory planning and even aspects of human sustenance in such a mission. The proposed launch vehicle is the PSLV-XL of ISRO. The paper describes how its alternate four stage propulsion system will carry a spacecraft with two astronauts on board and put it into the Earth`s elliptical orbit, eventually escaping its gravitational pull and making its way to Mars. Its return has been described using the Inspirational Mars Free Return Trajectory which uses Hofmann Transfer. To make the mission description through and exhaustive, an Innovative deep sleep mechanism has been described for the astronauts onboard. The mission is designed to last approximately 253 days for phase consisting of flying by Mars and its free return trajectory will last approximately 273 days reaching back to the Earth’s sphere of influence. Description of the composite material heat shield a parachute landing system will be used to safely land humans on the Earth’s ocean have been given. Thus it can be concluded that given the depth and exhaustiveness of the research that is done and documented in this paper, the proper implementation for the Mars Flyby mission will be a groundbreaking project in the history of planet Exploration.

ACKNOWLEDGMENT The authors would like to thank Mechatronics and Intelligence Systems Research Lab,

Department of Mechanical Engineering, Amrita University, Amritapuri campus for providing support to carry out the research and experiments. We would also like to thank Shri. G. Nagesh, project director, Chandrayaan-2 of Indian Space Research Origination (ISRO), India to actively clearing our doubt with the mission. Thanks to Srinivas, Scientist, ISRO who explained full details about the high success rate PSLV-Rocket.

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