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Mars Sample Return Mission Using Multifunctional Design, Consumable Structures and In-Situ Propellant Production Project

GULLYVER( G U L L Y V i s i t and E a r t h R e t u r n )Csaba Boros Smetanov hj 288/17, 92901 Dunajsk Streda - Slovakia, Europe borospace@yahoo.com

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Abstract

The goal of this mission is to transport of various samples from the Mars to the Earth using a lander. It is to be launched using only one launch vehicle - ATLAS V, in the year 2016, and return directly back to Earth in 2018. The primary landing area is a crater with gullies within +/-30 of the southern or northern midlatitudes. This is a territory with high scientific potential one of those which were in the relatively recent times much more geomorphically active then previously thought and if formed by groundwater or by other unknown liquid seepage, subsurface organisms could have been brought to the surface. The back-up landing site is Elysium Planitia near Athabasca Vallis. The presence of a possible frozen sea protected by a layer of volcanic ash, with occurrence of methane and ancient volcanic or geo/hydrothermal activity is possible there. Past life likehood landing site may be represented by Meridiani Planum and Melas Chasma with layered deposits ( microfossils form ). Consisting of the cruise stage and the lander, the spacecraft will be deduced from the previously planned Mars Science Laboratory. The entry, descent and landing of the unit are practically copying the landing of the probe MSL, but without the use of Sky crane which is replaced by classical landing approach as Viking and Phoenix landers. In fact, this landing option was selected because the MSL landing system is designed for higher weight rover fine landing, where the descent system is then become useless, while with the MSR lander, it is necessary to use the whole landed mass as a stabile launch platform for the safe launch of the MARV. For higher reliability and feasibility of current MSR mission success the MARV consists of three stages. The first stage is a hybrid rocket motor with self consumable structure as solid fuel with LOX oxidizer produced by ISPP technology from the Martian atmosphere thus representing the most risky part of mission. The remaining two stages are presenting proven and reliable technologies using classical solid rocket motors from STAR series making the launch of the Earth Return Vehicle to Martian orbit possible in case of first stage malfunction. The ERV can survive a couple of years on Mars orbit while waiting for the rescue mission. Because of this possible scenario the ERV contain a simple docking module for the possible Mars orbit rendezvous with some future small trans-Earth injection stage. The advantage of this concept is the exact ERV attitude measuring by onboard IMU and Star Trackers for orbit elements determination. The whole conception of the mission is characterized by the utilization of proven, reliable and feasible units from the past and planned missions, except of the multifunctional oxidizer tank of the first stage of MARV before landing the LOX tank is filled with high pressure Helium as a pressurized medium for the descent engines with hydrazine fuel and possible NTO oxidizer. Basically, this mission can be considered as a Groundbreaking Mars Sample Return mission similar to Soviet Luna Moon Sample Return missions with simple drilling equipments. The goal of this mission is to become acceptable for Direct Return advocates as well as for Mars Orbit Rendezvous ones while keeping the required maximum cost cap of two billion dollars.

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CONTENTSEXECUTIVE SUMMARY 1. 2. 3. INTRODUTION.. 5 SCIENTIFIC OBJECTIVES.. 5 LANDING SITES 6 Engineering Constraints Science Requirements Primary Landing Site Back-up Landing Site 4. MISSION DESIGN Mission Overview Launch Vehicle 5. SPACECRAFT SYSTEMS.. 16 Design Philosophy 16 Spacecraft 17 Cruise Stage 17 Entry, Descent and Landing . 18 Descent Stage Lander . 20 Mars Ascent and Return Vehicle (MARV) .. 26 First Stage . 29 Second Stage . 42 Third Stage . 45 Earth Return Vehicle (ERV) . 48 MARV Flight Trajectory Analysis . 51 Electric Power System . 57 ISPP . 58 Telecommunication .. 62 6. 7. 8. 9. MARS PROTECTION 62 13

EARTH PROTECTION... 63 MISSION COST ESTIMATION. 64 CONCLUSION. 65 67 90-94 4

APPENDIXES References

1.INTRODUTIONThe next challenge for the mankind in this century will be beyond doubt the landing of the first people on Mars and the consecutive colonization of the planet. The biggest problem is with the resources. Without available resources on Mars, this mission will be costly, slow and in case of unexpected situations much bigger troubles can be expected and the solubility of such will be determined with the long distance from the Earth and also in the impossibility of any improvisation because the limited presence of resources. Unlike the ancient and brave sailors/colonists, we must know what we are going into, we have to know if there is any exploitable, mainly watery resource on the Mars and of the presence of life on the planet. The positive answers would make the very first expeditions much simpler also from the technical point of view, but nevertheless from the view of public interest. The vanguards of this great mission would be the reception of the samples from the areas of the planet previously considered as suitable, at best with the use of the local resources. Such project would greatly help us to answer the critical questions mentioned above and might speed the preparations for the first manned Mars expedition.

2. SCIENTIFIC OBJECTIVESActually, we may dividing our aims to two basic cathegory looking for recent possible life and past life on Mars. For sustain life is liquid water mandatory. Hence, water is the primary focal point in site selection for the search for present life1. Sites where liquid water may may have been present recently at the surface include numerous gullies,1,7,8,9 the ends of geologically recent outflow channels 1,10,11 and sites of recent glaciation1. However, we have other possible biosignature of present life on Mars gas anomalies, for example metabolic end products e.g. methane4,5 might indicate the presence of active local microbial communities1. Biosignatures of past life exist in several possible form hydrothermal deposits, evaporates, concretions( blueberries), carbon-bearing rocks and soils, layered sedimentary rocks, some C, S, Fe redox state1. The primer objective of this mission is the transport of samples weighting max.0, 25kg from the landing area to the Earth together with the exploration of the landing area with help of various scientific devices. The next priority is to test the ISPP right on the surf ace of the Mars, e.g. production of sufficient amount of liquid oxygen for the MARV module from the carbon dioxide taken from the planets atmosphere. The goal of the GULLYVER mission will be similar as a WOBBLE mission13 elucidate the origin of the gullies observed in Martian valley and crater walls. Primary science package will be low and high resolution images, IR and Raman spectrometry and fluorescence analysis will determine the composition of the rocks and soil. The above mentioned sampling shall cover the landing area completely, taking the first ones already during the approach in a specified height from the surface of the planet together with the recording of the pressure and the temperature values. The next sampling of the atmosphere will be provided at the place of the landing during the sol and also during the night. The reason for that approach is for example the recording of the distribution of the dust particles in the sol/night cycle. On the surface a small meteorological station will be installed to measure 5

and record the temperature and the speed of the wind in different heights over the surface of the planet, like at the MPF mission. This metmast (meteorological station) will providing the useful data about wind direction and velocity to give safe balloon launch condition. There will be delivered to the Earth surface and subsurface samples also from the ground of the landing area. The surface and subsurface sampling will be provided by mechanical arm, (scoop and ultrasonic corer). Dust samples will be taken from ISPP filtering system (for CO2 acquisition compressor). For gullies investigation the lander may use several different approach: - tethered balloon with sample acquisition system with camera (to closeup imaging of the interesting terrain), some scientific instruments and balloon anchoring system. - tethered rocket/penetrator anchoring system it may be similar then above one, only for anchoring we shall using a rocket13. The lander will be equipped with instruments to define the mechanical properties of the ground which will be important for the further manned missions, and with seismometers and heat sensors to measure the heat flows. Astrobiology Experiment Package In order to obtain realistic and truth information about delivered samples from Martian surface, respectively subsurface, we will need some preliminary sample analyzing equipment similar can be easily adopted from ACP/HUYGHENS or cometary COSAC/ROSETTA missions. We would analyze some volatile possible products e.g. methane, hydrogen peroxide, water, or organic and inorganic compounds (mainly depletion of 13C in the isotopic composition of the collected organic matter). It may be based on MSL project SAM package (SAM Sample Analysis on Mars) 2. This very useful device may be implemented to this mission after increasing the overall weight the lander with cost increasing, too.

3. LANDING SITESMars is a dry and windy planet for the first sight, showing no clear visible remarks of water (except gullies) to be present, respectively signs of ongoing or ancient life. For the robotic mission with the hope of success, the landing site must present a compromise between such factors as the brokenness of terrain, the usable amount of the solar energy, the appearance of the dust storms and the needed precision of the landing. Because of such, the possible landing sites must meet the following sometimes even antagonistic criteria: Engineering Constraints: the landing site must be located in between +/-30 latitude because of sufficient sunshine for the solar cells12 the landing must be done long enough before beginning of the global sand storms the terrain must be flat enough to allow the easy landing of the controlled probe on the surface approximately for 2 km diameter circle, with the surface clereance max 0.35m( in the next decade I expected better landing accuracy e.g. minimally ten times of current MSL planned mission). the landing must be done at the clearest possible weather conditions as for the later more precise evaluation of the landing area also for the recharging of the batteries using solar panels 6

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the landing area shouldnt be too dusty, because of the sensible devices, solar panels and possible radar altimeter malfunction for the correct operation of the MARV, the lander shall not land on a terrain with bigger slope than 10 for the most effective use of the landing weight of the lander, the landing area must lay at least 0 km below the MOLA level, the landing parachutes are working much more effective in more dense atmosphere the bottom of the crater shall be relatively flat

Science Constraints: A key objective of this mission is collect samples for return to Earth. First step in the process is to target sites where liquid water was present and could have provided a clement environment for possible life. Idea prefers of possible current life finding for primary landing zone and possible past life in fossils form in ancient layers for back-up zones where aqueous sediments were deposited.

Primary Landing Site Possibility of liquid water occurrence with recent lifeGullies are found predominantly in the midlatitudes of both hemispheres in detail. Locations of gully clusters for example on southern hemisphere is located within 30-45S (on northern hemisphere those values are similar), elevation (-5000 to +3000m), slope (>10), and orientation (83,8 on pole-facing slopes)8. In 14 is described several problems of possibility of MSL landing near some crater wall with gully associated with small dimensions of those craters, high altitude and planetary protection issues(RTG generators likehood of water contamination by radioactive nuclear material in lander malfunction possibilty). For GULLYVER mission it will be additional requirements maximum latitude +/- 30 for solar power minimal efficience for ISPP system, thermal condition and telecommunication. The preferred location of those landing sites may be on places with minimal elevation - eg. ~ 310 340 ( Southwest of Meridiani Planum) and cca 50 to 100 ( northern part of Hellas basin)of East longitude. Furthemore, on Northern hemisphere we have more area with this restriction. Summary of Restrictions: +/- 30 Latitude Diameter of floor of target crater must be > 2 km Clearance of rock abundation could be < 35 cm Slope elevation must be < 10 and oriented to right direction for MARV lift-off right condition - The landing site must be below 0 km MOLA level - Because of landing site will be inside the crater, shadding must be minimized As is visible, because of these restriction the EDL system of curent mission will need utilization of slightly different landing scheme than MSL mission probably additional subsonic parachute, more fuel for descent engines for horizontal cruise and avoiding some obstacles on actual terrain with using the hazard landing avoidance system see picture below Fig.1..

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Fig. 1 Example of landing site with gullies, Crater with Latitude 18, East Longitude 117 ( Northern hemisphere). Taken by MRO (PSP_003902_1980). As we see, the best landing place circle would be inside the yellow circle with diameter less than 1 km, but it is likely out of limit of current spacecraft EDL technology. Next possibility for landing would be the outer part of crater near to rim, but more attractive are the layers deposit on the floor inside the crater (mainly for cable car malfunction case see below). Next possible landing sites with gullies are61: - PSP_004010_1500, Latitude (Centered) -29,9, Longitude (East) 54,0 - PSP_001666_1530, Latitude (Centered) - 26,8, Longitude (East) -34,7 - PSP_003565_1495, Latitude (Centered) - 29,8, Longitude (East) 322,6

Namely, MSR mission with precision landing capability would very beneficial for example, to land near to the previous mission rover (Astrobiology Field Laboratory), lander (Deep Drill66-72 etc.), it could have a short-range rover or an UAV helicopter64,65 for sample pack acquisition and transport to the MARV1.

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Back-up landing site Methane occurrence landing site with contingency of recent lifeThe back-up landing site will be Elysium Planitia (Fig.2). This area is the second largest volcanic region on Mars after Tharsis Montes. It is measuring about 1700 by 2400 km in size. Located on Elysium Planitia are the volcanoes Elysium Mons, Hecates Tholus and Albor Tholus. For the first time was this region studying by geologists through images returned by Mars Express spacecraft in 2005. They suggested that a frozen sea once existed on Elysium Planitia. This was determined from patterns seen in the ground near Athabasca Vallis a region at 5N and 150E about 1180 km south of Elysium Mons. Presumption is that there was an eruption of water from deep underground and a subsequent catastrophic flood that covered a wide area. This frozen see is roughly 45 m deep. This is suggested by plate-like surface features and may still harbor buried blocks of ice, protected by a layer of volcanic ash (see Fig.3).

Fig.2 Back-up landing site location map showing the location of the image below

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Fig.3 Mars water in frozen form may be located on this flat plain that is part of Elysium Planitia. The above image was taken by the Mars Express spacecraft this scene is approximately a few tens of kilometers across (rectangle on map above on latitude 5N and longitude 150E, image: ESA/DLR/F U Berlin/G Neukum) Figure below was taken by HIRISE camera of MRO spacecraft the image is centered at 5, 5N and 152, 2E.

The next recent finding by Mars Express has increased the interest of this region the spacecrafts Planetary Fourier Spectrometer has detected elevated concentrations of methane in the atmosphere over the Elysium Planitia. The similar Fourier Transform Spectrometer at the Canada-France-Hawaii Telescope by Vladimir A.Krasnopolsky team5 observed 10

a spectrum of Mars and discovered the methane in the Martian atmosphere and next two more teams had measured similar results. This suggested the most likely interpretation is that the methane is a product of the metabolism of indigenous microbes. On the other hand, alternative explanation is that the methane is a product of geothermal activity4 or comets. This region is near Athabasca Valles which was the candidate landing site for MER missions16. Lot of flood channels emanate from the Cerberus Fossae contain the bestpreserved examples of streamlined mesas, terraces and aqueous bedforms known on Mars. Such features can be interpreted as due the floods of water. The myriad channels of this site display distinctive thermal signatures despite the presence of a bright layer of dust covering the region this will be important constraint for feasibility of lander`s dust sensitive equipments. For those reasons this region would be a good candidate as a landing site for the first MSR mission. From engineering constraint view, this landing site will need smaller effort relative to propulsion, landing accuracy and sunlight power accumulation for ISPP system. For those reasons, descent engine will be fueled with hydrazine monopropellant (with smaller specific impulse) and main balloon shall be replaced with small Sojourner-like rover deployment via robotic arm on the Martian surface like cancelled 2001 Mars Surveyor Project15.

Past life likehood landing siteThe best candidate of possible past life finding is Meridiani Planum region see Fig.4. Sedimentary rocks exposed in this site record aqueous and eolian deposition in ancient dune environment that were arid acidic, and oxidizing 17.Orbital observations suggest that the combination of sulfate minerals and hematite found in Meridiani rocks may be unusual on the Martian surface. Because sulfates and iron oxides can preserve detailed geochemical records of environmental history as well as chemical, textural and microfossils18 signatures of biological activity, Meridiani Planum is an attractive candidate for Mars sample return mission. Because we have a Opportunity rover on current site to date, I suggest this place shall be as a next back-up landing zone for this MSR mission due the enormous useful taken science data from this territory.

Fig.4

Meridiani Planum image is centered at 0, 2N and 5, 8E (MRO)

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Next interesting back-up landing zone is Melas Chasma (Fig.5), a rugged canyon region in the central part of Valles Marineris, contains various interior deposits and other features whose origin in some cases remains enigmatic17. Unfortunately, those sites were cancelled for recent MER missions due the potential very rough Martian surface on those places.

Fig.5

Layers in Melas Chasma location is 11, 3S and 286, 3E (MRO)

Hence, EDL system for Melas Chasma Landing Site will be constrained with more precisious landing accuracy( like primary landing site) but with the same other conditions as a first back-up site Elysium Planitia.

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4. MISSION DESIGN Mission OverviewIn order of cost saving for the whole mission, I decided to choose the well know procedure of the transport of the GULLYVER MSR probe to the Mars using the launch vehicle ATLAS V 541 together with the common logistical infrastructure and mission control planned with the missions MSL for example. The given year of the start opens the possibility of the flight with the minimum of the energy needed for the start (therefore to rise the payload weight) and provides enough time to prepare the whole mission. The required launch energy C3 = 8,9km2s-2, and 8 km2s-2, respectively - see Table -119.Nr. Trajectory Type C3 DLA V DAA Launch Date Arrival Date

1 Earth to Mars 2 Earth to Mars 3 Mars to Earth 4 Mars to Earth Tab. 1.

1 2 1 2

8,9 8 6,1 6,3

-45,2 -5,3 -3,3 -22,2

5,3 5,4 3,2 3,9

7,1 -22 20,2 38,6

2/20/2016 3/21/2016 3/14/2018 3/14/2018

8/19/2016 1/20/2017 10/12/2018 11/12/2018

Upon arrival at Mars, the MSR aeroshell together with the lander enters directly into the atmosphere like recent MPF, MER and planned MSL missions. No portion of this vehicle is going to be placed onto orbit around Mars hence, there is no expense for the design, construction and operation of an orbiter. An aeroshell with outer diameter of 4,5 m provides aerodynamic deceleration down to a velocity where a parachute can be deployed the Entry, Descent and Landing (EDL) of this vehicle is practically copying the landing of the MSL, but with use of subsonic chute and hazard landing avoidance system. The hazard avoidance system uses descent imagery to ensure that the touchdown occurs on a relatively safe, flat place. After landing, the vehicle transmits the nested descent images back to Earth, deploys the high-gain antenna and solar panels (after dust deposition). After primary systems deployed, lander will deploy the metmast and monitoring atmospheric conditions at the landing site and collect an atmosphere sample. The low resolution camera will sent the panoramatic pictures to Earth to lander attitude check relative to the crater wall with gullies. Tethered Balloon Concept13 Primary goal is to deploy the main balloon with some scientific instruments and the sampling device. For this reason, lander's high resolution camera will take images of the crater walls and identify the position of the landing site to the target gullies with important areas. In the same time the metmast will collect atmospheric data and will communicate the results to the Earth to start the final launch countdown procedure of the main balloon. Every day a pilot balloon will be released, at different times of the day, and low resolution cameras will track its flight (for this purpose we will have a massive approximately one meter diameter high pressure tank filled with cca 8 kg Helium gas this will be the LOX tank of MARVs first stage consumable hybrid rocket motor). The deployment of the main balloon will start after having received the data validation (comparison of Earth simulation data with positive result), This balloon will be tethered with several km-long ultra weight cable from UHMWPE fibers with 1,or 2mm diameter62, with necessary altitude control63 after balloon anchoring on crater wall with gully this one will have a function as a cable car support for sample distribution to the landers Earth Entry Capsule (EEC). Inside of this cable may be some

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electrical conductive cable, like TSS-1 tethered satellite73 for electrical power distribution from lander to tethered balloon ( diameter of this cable is 2,5mm and weight cca 8,2kg/km fy Cortland Cable Company). The movement of those small sample canisters (from the balloon to the lander) will be secured by gravitational forces only after the balloon is separating from the investigated area on end of this science task. This balloon when released, will become an independent meteorological device for the higher parts of Martian atmosphere observing the temperature, pressure, aerosols distribution in upper atmosphere (of course, this idea may be limited from the communication point of view, if the lander would be as a retranslation station for communication between balloon gondola and Earth due the terrain restrictions dictated by lander position inside the crater). Tethered Rocket/Penetrator Concept13, 20 Advantage of this approach is the better accuracy of rocket impact to the most interesting place of the crater wall with gullies than previous balloon concept. This rocket may be unguided, but some calculations have to be carried out with the required initial conditions of shooting ( angle of shooting, wind direction ), we have wide experience with similar small rocket anti-tank weapons system ( TOW74, MILAN, etc.). As with the balloon concept we will use high resolution, zoom able optical system and an additional laser marker system for target determination75 for the tethered missile, where the warhead will be replaced by penetrator/anchor system the idea of this rocket is visible on Fig.43- Appendix p.68.As in20. The launch of the missile can be done as a soft launch using a loaded spring with the rocket motor ignition in flight, or using a cold start system with a launch tube and pure gas generator to avoiding the contamination of the lander with combustion products ( e.g. hydro carbonate and other products). The penetrator will contain various scientific equipments as seismometers, temperature sensors to measure the temperature of the subsurface layers in depths. Because the penetrator at the impact will generate large amount of heat and causing significant mechanical shock at the same time in the ground of the impact area, the sampling will be provided with the two possible following procedures: right after touching the ground the front of the penetrator will open with help of a drive which turns the potential energy of a torsion spring to a rotational movement (similar system as it is used in the hand borers to tighten the drill ) through the opened front end of the penetrator an ultrasonic borer will shift out with a hollow drill which acts also as the shell for the sample once the drilling is finished the sample is shifted upwards to the rope of the penetrator (together with its shell) where the cable car can transport it to the returning case in the MARV with the procedure described above all the gaseous and volatile products shall remain in the sample too

Or the sampling procedure will be the same as DS-2 penetrators21,22,23 with miniature drilling equipment in a perpendicular position related to the main longitudinal axis of the harpoon/penetrator/anchor system see the Fig.6.

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Fig. 6 DS-2 Probe with penetrator As we have seen, both concepts have pros and cons. If we use a balloon concept we will have to maintain a large-volume gaseous Helium storage tank inside the LOX tank of I. stage of MARV and the possibility for upper atmosphere research. Disadvantages of this concept are the very erratic wind distribution inside the crater and the potential difficulties with the balloon deployment (plaiting the balloon envelope to the some lander instruments). Primary, the target site of the balloon will be a slightly random process too. For this reason I decided to use the tethered, rocket powered penetrator/anchor system for crater wall investigating. For the surface and subsurface sample acquisition the lander will be equipped with a robotic arm (similar, from Phoenix mission) with simple grab mechanism and sample distribution system to the EEC in MARV. During the balloon and robotic arm operation in Mars exploration, upon deploying the main balloon, the LOX production for first stage of MARV will start immediately. The MARV stage is based on hybrid rocket engines with consumable structures as a solid fuel. This In-Situ Propellant Production system must produce approximately 400 kg of liquid oxygen (LOX) for the first stage of MARV during lander operations it will take nearly 500 sols (from eight to ten hours each sol, on those landing sites is sufficient sunlight to fully power the propellant production plant) but this is only true for equatorial conditions. In 12 power trades show (Fig.41/p.57), that on Martian equator the total energy per sol per solar panel unit area varies from cca. 700 Wh/sol-m2(Northern Spring) to cca. 1075 Wh/sol-m2 (Northern Autumn), but on the middle hemisphere +/-30 the situation is worse on the other hand at the end of Northern Autumn this energy equals to cca. 1300 Wh/sol-m2 it is bigger than equatorial one, but during the Northern Spring is decreased to cca. 330 Wh/sol-m2. This energy is one-half of equatorial one. For this reason we will need double surface of the solar panels than it is need when landing on the Martial equator. It means, that LOX production will be more erratic, and energy will be primarily managed for safely hibernating of the lander in case of a dust storm season and therefore a lower energy input period in the same time.

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Launch vehicleThe preferred launch vehicle is ATLAS V24 the same launch vehicle as for proposed MSL mission. The basic configuration is ATLAS V 541 type, which would consist of a liquid propellant first stage with four strap-on solid rocket boosters (SRB), a liquid propellant Centaur second stage, the GULLYVER spacecraft, and the payload fairing. This configuration involve 5-m payload fairing, four SRBs, and single engine Centaur second stage.

5. SPACECRAFT SYSTEMS Design philosophyThe whole conception of the mission is characterized with the utilizing of proven, reliable and feasible units from past and planned mission except of multifunctional oxidizer tank of first stage of MARV before landing the LOX tank is filled with high pressure Helium as a pressure gas for descent engines with hydrazine fuel and for other possible utilization, after ISPP system starting this same one will be as a Dewar`s vessel for LOX storage, and in finish, it would be as a solid fuel for hybrid consumable rocket motor ( HRCS ). This first stage will be designing according to new design philosophy for multiplying of existed parts of spacecraft. We can express the main philosophy of the design with a few words: Total uTilization of tructure Material TOTEM which stands for the total use/utilization of each gram of material deployed to the Mars. In the same way as the ancient cultures expressed their honor to such approach, this philosophy is the headstone of the whole construction philosophy of this MSR mission. In general the practice follows the main ideas: - Each of the structure elements have to withstand some load or have some function - After finishing the activities in certain functional dependence, the given structure will serve as solid fuel in the HRE Naturally not all of the structures will end as solid fuel, but they may have multifunctional character. The given approach was tested at mission Deep Space 1 for example. This philosophy has its argument in the fact that each kg of material deployed to the Mars surface costs in average 1 million USD. We will further distinguish if the use of mechanical and chemical parameters of the material will be done in one step or in series (one after another): Serial Utilization of Material (SUM) the use of the material will be done in series, e.g. first we use the mechanical properties of the given element and later its chemical energy Parallel Utilization of Material (PUM) when the use of the mechanical properties and the chemical energy of the element happens at the same time

Ideally each element will be multifunctional or used in many ways. We can express the given property with the so called coefficient of utility kutility . The given coefficient will have its definition field from 1 to infinity. It still remains a question what effect will this approach have on the reliability of the whole system. At first we may consider that each element of the system with higher value of kutility will have to be more reliable, which puts enormous demand

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on the quality of certain components of the probe. We may consider the given facts as the basic approach in this new field.

SpacecraftThe Mars Sample Return Vehicle consists of the cruise stage and a lander out coming from the planned mission MSL and it will have similar weight and identical size. The flight system, illustrated in Figure 7,8 is currently estimated to weight approximately 3,600 kilograms25. The weight allocation is on the table 2 see below.

CRUISE STAGE

AEROSHELL WITH STOWED LANDER

Fig. 7

GULLYVER Spacecraft in departure period of time Cruise Stage with Aeroshell ( w/ stowed lander body).

Cruise StageThis part of spacecraft provides the full range of necessary services during the several month passage to Mars included propulsion system for trajectory correction maneuvers, solar panels, star and Sun scanners and heat exchangers for the MARV control computer inside the aeroshell structure . The cruise stage measures 4,4m in diameter25 , weight is 600kg26 ( with 80 kg hydrazine25). Solar cell via 6,825 square meter produces electric power.

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CRUISE STAGE

BACKSHELL

STOWED LANDER

HEATSHIELD

Fig,8 Exploded view of GULLYVER Spacecraft

Entry, descent and landing system (EDL)The objective of this system is getting the MARV with ISPP and sampling devices safely onto the surface of the Mars. This system includes four stages to meet this requirement: Aeroshell Re-Entry Capsule Deceleration System Parachute Deployment system Descent stage Landing legs Aeroshell Re-entry Capsule Deceleration System It consists of two main parts Heatshield and Backshell. The heatshield is a carbon/epoxy composite and aluminum honeycomb sandwich construction with SLA-561V ablator. The weight of the heatshield is 382 kg7. The backshell is generally less heat loaded as the heatshield, but it bears the main part of the loads at the start of the carrying rocket and also at the landing of the Lander unit on the Mars. It weights 349kg7.

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Parachute Deployment System Parachute taken from the project MSL without modifications. It is Disc-Gap Band type and has 19,7m diameter. It is scaled geometrically from Viking heritage and constructed using MER technology and materials. This shall be enough for back-up landing zones, but for gullies target site is not sufficient. Fortunately, there are several possibilities to enhance the landing accuracy of the lander. One is using the additional maneuverable subsonic chute with some electrical actuators for attitude control relative to a target site and additional fuel requirement for precision landing. Unfortunately, that mean weight increase for the lander and smaller payload weight, but because of atmospheric variations (change of density, temperature, wind) this is the only way. The subsonic parachute will operate from M 0,8, furthermore the additional testing and developing shall be less expensive, than new supersonic chute testing28. More effective is the precision guidance and control of aeroshell capsule along the descent path in hypervelocities with slightly more L/D ratio30 from Apollo and Shuttle guidance technology( planned MSL mission will use L/D = 0,2427). The remaining fuel for descent engine will be a crucial factor for the successful landing it will need some optimization software with close-up guidance computer29 see the Fig.9. Summary, it will need additional subsonic chute anchored on lander itself, actuators, guidance and control with fuel consumption and optimization in real time at descent.

Fig.9 Possible landing scenario (using additional subsonic chute and hazard avoidance system)76

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Landing legs The lander is equipped with four landing legs, designated to anchor the lander firmly to the landing position. Common landing systems using rocket engines mounted to the lander directly are working on the principle, that the touch of one of legs with the surface immediately deactivates the engines, which may cause significant shocks to the probe when settling while the rest of the legs are in certain distance from the surface. More sensitive system will be the Sky Crane in MSL mission, but this landing option is due that MSL landing system is designed for higher weight rover fine landing, where descent system is then useless in contrast to MSR lander, where for the MARV safe launch is necessary to use the whole landed mass as a stabile launch platform. The legs configuration is based on Viking technology31, but replacing any heavy metal structure with light-weight carbon/Kevlar reinforced composite with epoxy resin matrix (but with some outer part from metal alloy as plastic deformation zone if the landing loads exceeds bigger values). Landing legs provide: - stable (upright) landing on the Mars surface with energy absorption that minimized the landing shock loads - support for the lander during its operational life - stable platform for cable car system ( between lander and gully) - stabile position of the launch pad (requirement for MARV lift-off to Earth) WEIGHT ALLOCATION OF SPACECRAFT Allocated Mass (kg) Cumulative Mass (kg) Lander Structure (dry) 732,12 732,12 Payload 948,95 1681,07 Descent Stage Propellant 300 1981,07 Supersonic Chute 2177 2002,07 78 77 100 Subsonic Chute (L/D=0,8) 2102,07 + actuators, and thermal batteries Heat Shield 382 2484,07 Backshell 328 2812,07 Cruise Stage (wet) 600 3412,07 Tab.2

Descent stage - landerAs a previously mentioned, the main problem of current and planned robotic missions to Mars is the maximal weight of payload transported to a Martian surface. There had been done a lot of investigation about EDL systems in the Viking era, when not constrained by todays relatively small budgets. From this point of view this mission will be breaking out of Vikings EDL with adding a subsonic chute (with increasing the weight of EDL system). Maximum weight of payload with Vikings technology is approximately 1000 kg28(- 2 km MOLA reference geoid radius). Reference weight for GULLYVER mission is the weight of the planned MSL mission practically, aeroshell body, cruise stage, tank, RCS system, some engines, radar will be the same, except the structure and landing system(legs). The MSL payload 850 kg, the rover will be smaller, than the one for GULLYVER MARV, ISPP system, solar panels, battery and science instruments summary mass will be approx. 1000kg, but the whole mass shall be

20

the same 2069 kg (landed mass plus descent fuel mass)! For those restrictions we may solve those problems in several ways: - using higher Isp bi-propellant fuel/oxidizer for descent engines than current monopropellant/hydrazine system - canceling the cosine thrust loss (main MLE engines are canted out by 25 to avoid a plume interaction with suspended rover under Sky Crane system) - we have no Sky Crane phase and fly away phase, but those phases in MSL mission will be transformed to cruise phase of current MSR mission, when hazard avoidance system active - the metal structure of Sky Crane of MSL will be replaced with compact, light-weight composite structure (the central sarcophagus + frame) - composite landing legs - lower MOLA elevation ( from +1 km MOLA level of MSL to 0 km MOLA one of current mission) - pressure Helium gas for descent engines will be stored inside the LOX tank of I. stage of MARV ( hybrid consumable structure ) Those restrictions are valid for the primary landing site, for back-up landing sites we can using the same propellant (hydrazine) and only one 19,7 m diameter parachute. Because of current MSR lander will land like the past Viking and current Phoenix missions, the plume interaction with ground in last phase of this maneuver will be the most critical event. For this reason I decided to reuse four old, Vikings version of MSL engines MR 80 600 lbf or 2670 N hydrazine engine (with 18 small nozzles) for the last phase of landing, and next four engines will be same type as a MSL mission MLE with single nozzle and new throttle valve32,33,34. Those engines will be activated in the beginning phase of descent and deactivated in last phase of landing. The MLE engines will be replaced by bi-propellant engines for primary landing site (gullies) and next four Vikings engine will remain without change. The oxidizer will be Nitrogen Tetra Oxide (NTO N2O4) and hydrazine N2H4 as fuel remain. The structure of lander consist of the sarcophagus base, which acts also as a Warm Box for internal important parts, as MARV, batteries, ISPP parts, telecommunications etc. This box has sandwich type walls with some reinforcement, ribs, honeycomb structures and aerogel passive insulation. This box is directly connected to the composite frame structure with attachments, joining parts with the backshell structure and with landing legs. Inside of the composite frame are situated the fuel and oxidizer tanks, propulsion, RCS, science instruments, antenna and the solar panels see the Figure 10-14.

21

UPPER DECK and MARV HOLDER BACKSHELL SEPARATION PLACE ( 4x ) AFT CAP ISPP SYSTEM

RCS ROLL HYDRAZINE THRUSTERS ( 4x ) STOWED HIGHGAIN ANTENNA

FORWARD CAP

SARCOPHAGE ( WARM BOX ) SOLAR PANELS STOWED ( 2x ) LANDING LEGS ( STOWED ) 4x ROBOTIC ARM

Fig.10 Top view on the stowed Lander

Fig.11 Lander in powered descent phase 22

PROPELLANT TANKS (4x) FINAL DESCENT ENGINES MR-80 600 (4x)

MAIN RIBS

MAIN DESCENT ENGINES ( MLE or R-40B-900) 4x

SAMPLING DEVICE

Fig.12

Lander with closed and opened legs

23

Fig.13 GULLYVER Lander with deployed solar panels, high-gain antenna and robotic arm

MARV PNEUMATIC ACTUATORS (2x) AFT COVER

FORE CAP

Fig.14 Lander configuration before MARV Lift-off.

24

WEIGHT ANALYSIS OF LANDERComponent Sarchopagus Frame Landing Legs Tilt-up mechanism (pneumatic cylinders, gas generators, hinges,etc.) MLI insulation Hazard Avoidance System SUBTOTAL R-40B 900 lbf MR-80 600lbf Unit s 1 1 4 Mass Unit (kg) 194 110 19 Total Mass (kg) 194 110 76 Mature Index (-) 6 6 6 Continge ncy (%) 30 30 30 Total Weight w/ Cont. (kg) 252,2 143 98,8 Size (mm) Vendor Comments

1

42

42

6

30

54,6

1 1

3,3 4

3,3 4 429,3

5 6

20 30

3,96 5,2 557,76 Shall be throttleable More changes Shall be about 20% lower weight in the future (www.nasa.gov )

4 4

6,8 ~8

27,2 32

4 2

15 5

31,28 33,6

Aerojet Aerojet

Propellant tanks

4

15,99

63,96

2

5

67,16

587,5x722,8 mm

ATK, PSI

Accessories (tubing, valves, control ,etc.) RCS roll MR107B 30 lbf Oxidizer (NTO) Fuel (Hydrazine) Helium Gas SUBTOTAL MARV ISPP system Solar panels High-gain antenna w/ structure Low-gain antenna Battery Telecom Samplig device Robotic arm,camera, accessories(cor er) Metmast Penetrator/Anc hor rocket(w/spring or gas gen., launch tube, etc.) PanCam +laser guider SUBTOTAL TOTAL

1 4 1

18,8 ~2,2 100 200

18,8 8,8 100 200 8,9 459,66

4 1

15 2

24,44 8,98 100 200 8,9 474,36 696,45 Margin included in Tab.XX Aerojet Must be optimized Must be optimized Stored in LOX tank of I.Stage, 15MPa

1

8,9

1 1 2 1 1 2 1 2 1 1

696,45 48 22 7,7 0,95 29 4,8 5,8 8,5 1,0

696,45 48 44 7,7 0,95 58 4,8 11,6 8,5 1,0 6 4 4 4 4 4 6 6 6 30 15 15 15 15 15 30 30 30

62,4 50,6 8,86 1,09 66,7 5,52 15,08 11,05 1,3

1

20

20

6

30

26

1

3,0

3,0 904,00 1792,96

6

30

3,9 948,95 1981,07

Table 3. Weights allocation of lander. For simplicity, lander with bi-prop engines will use the same tanks as a lander with monoprop(hydrazine) - MLE engines.

25

Mars Ascent and Return Vehicle (MARV)The most important component of the whole mission is the MARV. It presents a unique challenge it is to be the smaller carrier rocket so far, together with the fact, that it will be the first automatically launched rocket from the surface of an alien planet (except of Luna 16, 20 and 24 successful samples return from the Moon). The above mentioned facts will be cardinal for design, testing and running of the given device. The whole system is in fact in contrast with the present trends of the development of launch vehicles on the Earth, when decreasing the useful weight! The common design tends for continuous increase of the useful weight, although there are sign of the increasing need for smaller rockets to transport nano-size satellites to LEO and beyond. Therefore we can finish the introduction with the thought, that the development of the MARV will have great impact not just to the planetary research, but also to the small or micro-size launch vehicles development here on Earth. Basically, we have two main approaches for sample return from Mars surface Direct return to Earth and Rendezvous on Mars orbit (MO). The simpler is direct return problem is in only with overall weight of this vehicle (and in the same time with the weight of lander). The merit of next system rendezvous of Mars ascent vehicle (MAV) on Mars orbit, is a more lightweight lander with small rocket and it is more preferable for NASA because of the smaller lander with smaller MAV is less risky (from technology and costs point of view) in failure case. Disadvantage of this system is requirement of automatic rendezvous and docking on Mars orbit without possible Earth assistance. If we would use the ISPP system on Mars surface, it would be the best option with direct return approach with smaller lander weight. This not new idea has a little problem. For example, for manned mission this concept is required for cost saving and some additional ISPP equipment weight is smaller than propellant produced by one. On the other hand, for robotic mission we prefer a low weight, low cost lander MSR mission. In this extreme case we would have a lander with ISPP system weight which is equal as a possible propellant weight, brought from Earth. Next problem is political lot of the specialists doesnt believe, that some ISPP system would work some years (2 years) in inhospitable conditions without failure. We have a possibility to double those systems, but we turn back to a primary problem minimum weight. If we look at this problem from manned mission perspectives, maybe we will need both approaches some specialists prefer Mars orbit rendezvous (MOR), other prefer direct return to Earth (Zubrin, Mars Direct). Generally, manned mission have one basic condition if we want go to anywhere, we must test it before. The best opportunity is MSR mission, because we may imitate the whole manned mission from departure to Mars to the arrival to Earth.

26

Fig.15 MARV`s I.stage boost phase. The characteristics of the rocket The main idea of MARV design is We always have some chance this is a three stage rocket, where first stage has a hybrid rocket propulsion with LOX oxidizer (produced by ISPP system) and next two stages are the solid rocket motors with capability reaching the Mars orbit with the ERV if the first stage would suffer a fatal failure or ISPP have some malfunction ( Trajectory Variant B). This is a philosophical boundary between some new untested part of mission and feasible, proven good understood technology (for example the upper two stage solid rocket MAV with rendezvous technology in connection with a lot of work had been done already). Summary: if the MARV flight will be successful, it will be Direct Return to Earth, if no it would be a Mars Rendezvous on Orbit (MRO) the ERV will remain on Mars orbit couple of years until to rescue mission carried out by some Delta II, or similar smaller launch vehicle. I think this concept is a very good compromise for MSR missions, incorporated with both approaches described above with this more feasible and cheaper system. The landing module introduced below, will not contain fuel in the conventional way before landing on the Mars except producing LOX from the Marss atmosphere as oxidizer, the construction of the pressure pots for the liquid oxygen. The MARV is a three stage ascent vehicle, fully controlled in each of its axis with starting weight of 1096,5 kg. The objective for the given launcher is the transport of the ERV to the trajectory leading to the direct Earth return ( Trajectory Variant A Fig.15 ). However, as it was mentioned above, the upper two stages of this vehicle have the capability to deliver the

27

ERV to Mars orbit without the first stage ( Var.B). see the Fig.35/p.53. On Fig.16-18 is the MARV with the description of its main parts.I. STAGE II. STAGE - STAR 13B (4x)

III. STAGE STAR 24

EARTH RETURN VEHICLE (ERV)

EARTH RETURN CAPSULE (ERC)

Fig.16. MARV overview MARV WITH AERODYNAMIC FAIRING

BIOSHIELD ( APPROX: IN THE MIDDLE OF FAIRING )

Fig.17. MARV with visible bioshield wall

28

MARV WITHOUT AERODYNAMIC FAIRING

LITVC SYSTEM OF I. STAGE

MLI BLANKET

OUTER PART OF NOZZLE DEPLOYED

Fig.18 MARV without aerodynamic fairing The second stage of MARV consist of four canted solid rocket motors from STAR series STAR 13B (see App.Fig.49-50). These motors have the same overall impulse, but lower average thrust and longer burn capability. This requirement is due to the fact, that in case of malfunction of first stage, the second stage would have too aggressive thrust level (originally this type has average thrust nearly 4x7600N = 30,4 kNs, with gravitational acceleration on Mars approx. 17g!). It would mean higher velocity loss for upper stages of MARV due the drag in the relatively thin Martian atmosphere. In this case the task for second stage is to deliver the third stage with Earth Return Vehicle to upper atmosphere, where the atmospheric drag is minimal and it will need only tangential addition to the final, orbiting velocity. The third stage is equipped with one solid rocket motor STAR 24 (Fig.51-52). This type is able to get the ERV to Mars orbit (to a very eccentric elliptic orbit). At last, the 9 kg of hydrazine fuel of ERV will be consumed for decreasing, or circularizing the ERV`s final orbit around Mars.

First StageIn consumable rockets field exist some investigation and patents97 103, but only on a theorethical base of possible investigation. I think these structure is very good applicable for hybrid rocket application, where a solid fuel has a potential as a structure application. From burning rate point of view, these structure material burning very slowly in oxidizer flow( we may show some dependence between hardness of material/fuel and burning rate e.g. paraffin burning ten times faster than UHMWPE). In 104 wrote, that the high tensile strength fuel lasted much longer in the burn and started breaking at a web thickness smaller (it is mean 29

less residual unburned solid fuel) this is an advantage of hybrid consumable structure (HRCS).Some reflections about HRCS are in 101,102 see the Appendix. Consisting of the great Dewars vessel in general, functioning at the same time as burning space for the first stage of HRM the here introduced type of the Hybrid consumable structure is the one where the chemical and the mechanical properties of the solid fuel are used at the same time it means that the tank with the LOX is to be burned from outside in the way the mechanical toughness of the pot walls remains. As the pressure of the helium above the LOX is decreasing together with the thickness of the wall, caused by the burning, will the tension of the construction remain almost on a constant level. (However other factors, like temperature stress and others were not considered in advance.) The capacity of this tank is approx 400kg of LOX at -186K while its density is 1140kgm-3. During the LOX production the tank is well insulated with the high vacuum between the walls of the burning chamber and with good thermal insulation above that. The main issue is to prevent the losses of LOX due to its evaporation while storing. When the rocket is working will consume 400kg of LOX and 167kg of solid fuel, the construction material of the Dewar pot, as previously introduced. In sectional view the first stage Fig.19, it has been seen that the burning area consist of two spherical shape where the burning of solid fuel is passing from inside the channel to outside direction, where the solid fuel is the construction of the Dewar vessel with the LOX, too.ROLL THRUSTERS ( GAS GENERATOR EXHAUST GAS GENERATOR GASES) FOR NTO TANK PRESSURIZATION FIXED NOZZLE

PYROLYTIC CRITICAL NOZZLE INSERT

ADDITIONAL POSTCOMBUSTION CHAMBER AFT LOX INJECTOR FORWARD LOX INJECTOR IGNITER

DEPLOYABLE, OUTER PART OF NOZZLE N2O4 (NTO) IN DIAPHRAGM PRESSURE TANK HYBRID CONSUMABLE STRUCTURE (50% EPOXY + 50% UHMWPE) POSSIBLE LOCATION OF HIGH PRESSURE HELIUM TANK

LOX

MAIN INTERSTAGE STRUCTURE

Fig.19 MARV First stage section view

30

As we seen on table 10. in App. we have only several possibilities for hybrid consumable fuel selection. Practically, we may use from the following combinations: UHMWPE fibers in UHMWPE matrix UHMWPE fibers in Epoxy matrix Kevlar fibers in Epoxy matrix, or Combination of the above UHMWPE Fibers in UHMWPE matrix The Dyneema (or Spectra) is a super strong polyethylene fiber which offers maximum strength combined with minimum weight. It is extremely durable and resistant to moisture, UV light and chemicals. The Dyneema fiber itself has 15 times higher tensile strength as the common steel, more than 40% stronger than aramide fibers, but only at lower temperatures as being thermoplastic. At the same time it has higher thermal conductivity, so the unnecessary heat has to be led away (in regenerative way for example) while the engine is working. As the most of thermoplastic materials, they are subject to creep under continuous load . 44 However, the use of PE fibers in composites is impeded by the strong temperature dependence of their properties and the difficulty of obtaining sufficient chemical interaction with polymeric matrices35 (because of its non-polar molecule). There exist several technologies for PE/PE composites (Polyethylene fiber in Polyethylene matrix), when PE fiber is chemically activated and pressed, for example UHMWPE, SK-75, DSM The Netherlands36) with its long molecule chain structure together with Dicumyl Peroxide (DCP) as cross linking agent. Another example is37, when the composite beam specimens are pressed and cured at 125C, with 200 bar pressure, with total pressing time of cca. 30 min. There exist some sintering process too38, but this is more acceptable for PE powder processing. Patent39 invented for the providing of composites on UHMWPE fiber with UHMWPE matrix this material has superior mechanical properties and may be sterilized (Mars protection). Technology is based on matrix powder reheating process (after melting UHMWPE powder is heated above 140C, cooled and repeatedly melted) and we will get a matrix, melted at 10C lower temperature than the original powder. This is critical for the invention, because we can make a PE/PE composite material (with melted UHMWPE resin and unmelted UHMWPE fiber). Maximal cryogenic temperature utilization varies from -150C to -268C (Fluoroloy)40. Its impact resistance (shock absorption capacity) in cryogenic temperature is extraordinary43. UHMWPE Fibers in Epoxy Matrix As mentioned above, in the classical composite structure, but with UHMWPE fibers in epoxy resin matrix, usually we have a problem with bonding the fiber to the matrix structure due to the non-polar characteristic of this material. One method to increase the affinity of PE fibers is plasma etching41. However, there is possible to use a pre impregnated UHMWPE/Epoxy material 42. Hence, that is clear, than mechanical, thermal properties will be worse than above, but from the technological point of view, it is more practical ( e.g. filament winding for pressure vessels) and proven. Kevlar/Epoxy Matrix These aromatic polyamides are part of the nylon family. Aramide fibers have a combination of good tensile strength and modulus they have excellent toughness and impact resistance. However, due to the lack of adhesion to the matrix (similar problem than above

31

material, but smaller), this material has a poor transverse tension and interlaminar shear strength. Finally, the burn characteristic with oxygen is worse (has a higher oxygen index ~ 36), than UHMWPE (18). The thermo dynamical characteristics of fuel materials of the 1st stage The 1st stages oxidizer will be LOX. The solid fuel will be the combination of UHMWPE and epoxy resin. The thermodynamic calculations were done with REAL 2.1 version code the gas flow through the jet were considered as balanced with polytrophic exponent, with equal enthalpy, so these are the theoretical peak values of the possible reachable specific impulse ( for very large engines with nozzle expansion to absolute vacuum). The next calculations had been done with freezing flow adiabatic condition this represents the lowest possible theoretical specific impulse. In reality for the first rough calculations, for smaller rocket engines is the adiabatic, frozen flow calculation enough, but to be more exact, for bigger ones, this equilibrium calculation is needed the real value of Isp, theor will be lying somewhere between those extreme ones. Naturally, the other calculations were corrected with various values of efficiency (of the combustion, of the engine).

32

3800

3600

3400

p=0.5MPa p=1.0MPa 3200 p=1.5MPa p=2MPa p=2.5MPa p=3.0MPa Specific Impulse [Ns/kg] p=3.5MPa 3000 p=4.0MPa p=4.5MPa p=5.0MPa p=5.5MPa p=6.0MPa 2800 p=6.5MPa p=7.0MPa p=7.5MPa p=8.0MPa p=8.5MPa 2600 p=9.0MPa p=9.5 MPa p=10.0MPa

2400

2200

2000 0 0,5 1 1,5 2 2,5 3 3,5 4 4,5 Mixture Ratio O/F

Fig.20 UHMWPE, Epoxy, / LOX Specific impulse of the propellant relative to the mixture ratio for variouschamber pressures equilibrium condition, for various chamber pressures, and ambient pressure 0,001 MPa.

33

3400

3200

p=0.5MPa 3000 p=1.0MPa p=1.5MPa p=2MPa p=2.5MPa p=3.0MPa 2800 p=3.5MPa p=4.0MPa p=4.5MPa p=5.0MPa p=5.5MPa 2600 p=6.0MPa p=6.5MPa p=7.0MPa p=7.5MPa p=8.0MPa 2400 p=8.5MPa p=9.0MPa p=9.5 MPa p=10.0MPa 2200

2000

1800 0 0,5 1 1,5 2 2,5 3 3,5 4 4,5

Fig.21 UHMWPE, Epoxy, / LOX Specific impulse of the propellant relative to the mixture ratio for variouschamber pressures frozen flow, for various chamber pressures, and ambient pressure 0,001 MPa.

34

3600

3400

3200

p=0.5MPa p=1.0MPa p=1.5MPa p=2MPa

3000

p=2.5MPa p=3.0MPa p=3.5MPa p=4.0MPa p=4.5MPa p=5.0MPa p=5.5MPa p=6.0MPa p=6.5MPa p=7.0MPa p=7.5MPa

2800

2600

2400

p=8.0MPa p=8.5MPa p=9.0MPa p=9.5 MPa p=10.0MPa

2200

2000

1800 0 0,5 1 1,5 2 2,5 3 3,5 4 4,5

Fig.22 UHMWPE, Epoxy / LOX Specific impulse of the propellant relative to the mixture ratio for variouschamber pressures middle, more realistic theoretical value ( approx. kinetic flow), for various chamber pressures, and ambient pressure 0,001 MPa.

35

The functional and geometrical characteristics of the 1st stage HRE The given engine will work at the following conditions: pSK = var., mOX = const. an O/F =const. (pSK = pressure in combustion chamber, mOX = oxygen feed rate and O/F = Oxidizer to fuel ratio)96. This means that during the function of the HRE the pressure in the combustion chamber will gradually decrease, the oxidizer supply together with the mixing ratio of the oxidizer and the fuel will remain constant. The condition pSK=var means decrease of the pressure during the functional period of the HRE in order to keep the condition of the consumability of the material of the tank as consumable structure, so the burning down of the inner side of the tank will be linked with the pressure decrease and so the tension in the material remain constant. To maintain the given decrease, it is enough to consider the erosion of the critical insert of nozzle of the engine. The conditions mOX=const. a O/F=const will be met with the help of the functional scheme of the given HRE, where the LOX is to be led as to the frontal section of the combustion chamber (primal circuit) as to the mixing chamber (secondary circuit) situated before the nozzle of the HRE. The LOX supply to the primal as to the secondary part will be redistributed with the help of the main valve in the inner section of the LOX tank (changing the flow trough rate in these circuits, while keeping the LOX supply constant). The reason for the given scheme is mainly the fact, that the spherical shape fuel channel, see Fig.19, is very slim approximately 1:80-100, meaning that this HRE will work at large oversupplies of the fuel materials which are to be burned additionally in the additional combustion chamber near to the nozzle of the engine. The procedure described here has an another advantage too when taking in account the wall thicknesses of the tanks and the needed length of the working period of the engine because the given MARV at lower structural height has quite big outer diameter 1.1m, will have bigger aerodynamic drag when descending trough the atmosphere (also if the Martian atmosphere is approx.100x less dense than on the Earth) so it will be more beneficial to use lower traction at longer working period. Although the optimalization of the MARVs trajectory is not the part of this work, we may designate the engine working time to 100 sec. It is clear from this, that when using appropriate wall thicknesses, the burning rate of the solid fuel must be very low, approximately 0.15-0.2mms-1. This is only possible, at very low GO values in the diffusion area of the given solid fuels burning law. The advantage of the given diffusion area, is that the burning rate is independent of the pressure in the combustion chamber and so the calculation model of the HRE combustion will be simpler. The thrust control of the 1st stage is not considered the trajectory changes will be fine tuned with the help of the RCS system of ERV. The parameters of the MARV are in the table 4-8.. Despite the complicatedness of this conception for the first sight, the advantage will be the relatively high initial pressure (10MPa) in the combustion chamber maintaining the optimal mix ratio for the whole range of functioning and also the very sound engines construction number of 0.9. Within the given parameters this engine overcomes even the most modern engines for TPH, particularly in its specific impulse Ispmax Theor = 3500Nskg-1. Practically, combustion efficiency of hybrid rockets are smaller, - we may calculating with 0,94 value. In my opinion the given limitation of the UHMWPE/ Epoxy composite material ( interlaminar shear strength, creep etc. see above) will be possible to eliminate with bigger safety coefficient for the construction above, where the wall thickness from this material will be several times bigger like the needed minimum. Because of the quite low density of the

36

material, such approach wont lead to massive weight increase of the given part. For these reasons of Vis major the outer walls of the pressure pot for LOX will be this material, while its strength will be used right before of the ignition of the first stage only. Then the LOX tank with the use of compressed He will be continuously pressurized up to 15MPa and so the walls loaded. The given load will change when the engine starts working, because the 10MPa of pressure in the combustion chamber (between the given walls) will decrease the differential pressure to 5MPa. In this phase the temperature of the burning will reach approximately to 3600C so the inner thin walled pot from carbon/epoxy composite with Al alloy (or Titanium, respectively stainless steel) liner, will take over the mentioned load, as the Dyneema loses its mechanical properties at this temperature. Resuming all this we can conclude, that the mechanical properties of the Dyneema will be used right before the ignition only and later it will be used as solid fuel for the 1st stage. Because of the high thermal conductivity of the UHMWPE, inner part of the LOX tank will be solved as thermal exchanger, where the gaseous He will pressing itself from the lower to the upper half of the tank, above the LOX level. Therefore the liquid oxygen wont be in direct contact with the very hot surface of the metal liner (especially at the final phase of the engines function), but there will be a layer of inert gas flow He, which will have two functions: working as cooling medium will be used to press the LOX into the combustion chamber of the HRE, where the heat taken from the tank walls will be also used to pre heat the He itself and so for the given amount of oxidizer smaller amount of gas will be necessary Those will be similar to the VAPAK system in QuickResponse approach105, where the pressurizing of the propellants had been done by self-evaporating only with the thermal energy, produced by engine. The outer pot of the LOX tank will be made of the same UHMWPE/epoxy composite, insulation, and outer carbon/Epoxy composite shell this material will be burned slowly, but also taking the pressure load over from the combustion chamber of the HRE. We can find the further properties of the polymer materials, which can be used as hybrid rocket consumable structure, in the table 10 in App.. In certain places of the construction temperature sensors will monitor the thermal field during the HRE activity. The LOX injectors together with the main LOX valve are situated at the lowest and at the uppermost part of the tank, so the intake for the 1st stages HRE shall be problem free for the whole functional period. The ignition will be pyrotechnical (pyrogenic type) with backup charges and independent electrical igniting circle to increase the reliability of the ignition. The pressurized He will be stacked in a toroidal tank at 30 40MPa of pressure, situated between the spherical LOX tank and the upper stages of MARV. In the lower parts of this engine is the secondary LOX injector with the aft mixing chamber. As it is to be seen on the picture nr.19, the control of the 1st stage is done with the Liquid Injection Thrust Vector Control (LITVC) in the main stream of the exhaust gases. The liquid component is N2O4 (NTO)46,47.

37

WEIGHT ANALYSIS OF FIRST STAGE OF MARVComponent Cabling Battery LITVC (dry)53 NTO injectant45,53 Gas Generator Moveable nozzle LOX injectors Main Valve LOX tank internal structure RCS Roll Thrusters Igniter Aft Combustion Chamber Composite outer chamber shell Internal Support Element of Combustion Chamber Internal ablative insulation Outer MLI insulation Nozzle w/ critical insert High pressure Helium tank Outer Fairing Interstage Subtotal (dry) Fuel (UHMWPE/Epoxy LOX Summary weight I.Stage Units 1 2 1 1 2 1 2 1 1 4 1 1 1 20 Mass Unit (kg) 0,65 0,46 5,84 11,2 1,1 1,43 0,38 0,87 0,5 0,22 0,3 1,7 1,6 0,05 Total Mass (kg) 0,65 0,92 5,84 11,2 2,2 1,43 0,76 0,87 0,5 0,88 0,3 1,7 1,6 1,0 Mature Index (-) 4 4 6 6 6 6 6 6 4 6 6 6 Contingency (%) 15 15 30 15 30 30 30 30 30 15 30 30 30 Total Weight w/ Cont. (kg) 0,75 1,06 7,59 11,2 2,53 1,86 0,99 1,13 0,65 1,14 0,39 2,21 2,08 1,3 Size (mm) Vendor Comments

1 1 1 1 1 1 45 1 1

2,4 0,4 2,0 4,1 0,8 1.2 160 400

2,4 0,4 2,0 4,1 0,8 1,2 40,75 160 400 602,55

6 5 6 6 6 6

30 20 30 30 30 30

3,12 0,48 2,6 5,33 1,04 1,56 49,01 160 400 611,35

Table 4. Weights allocation of First Stage of MARV.

Upper StagesUpper stages of MARV are capable to deploy the ERV to the Martian orbit without first stage also. On the other hand these stages will accelerate the ERV to 4,5 kms-1 approximately, to reach the hypervelocity for direct return to Earth. The above mentioned stages are fully controlled by jet vanes, situated on the exit parts of own nozzles, as seen on Fig.23. It seems to be a little bit as anachronism, but the advantages of such system dominate over the disadvantages48. Advantages are: proven technology, low actuation power, high slew rate, roll control with single nozzle45 and mainly, - application of this control system has low influence to the rocket motor performance or hardware, development and testing and finally, its mean 38

lower cost. Disadvantages are thrust loss of 0,5 to 3% and erosion of jet vanes (limited duration). More significant is the thrust loss, because the combustion temperatures of current solid rocket motors dont exceed temperatures of the LOX/Ethylalcohol combination for example (V-2 German missile example). From the design point of view, these upper stages are characterized by utilizing solid rocket motors of STAR series. The technical parameters of these rocket motors are to be seen in Ap. Due the volumetric constrain of MARV, I decided to use a cluster of four motors consisting of STAR 13B50,51,52 type motors with canted angle of 20 measured from the central, axismetric axis, positioned around the III. Stages nozzle see the Fig.23,24. . In the same time, the axisymetric thrust directions of these motors must get along the weight center of upper stages due to the possible thrust misaligments mainly in end of burning. Attached stabilisators may also help to repair these misaligments.The jet vane control system is characterized by lightweight structure and utilizing of four servomotors for the both stages (one servomotor runs two jet vanes via one gearbox for the first time it is running jet vane for II. stage, then jet vane for III. Stage see the Fig.23,24). The second stage is attached to the first stage via aerodynamic fairing there are four shell structures with STAR 13B solid motors attached, STAR 3-A separation motors, and hardware. Each of the STAR 13B motors has an individual gearbox with one jet vane and clutch interface with servo on the third stage (Fig.25). These motors are attached to the third stage. The third stage of MARV is attached to the first stage via four beams. These beams are characterized as composite, high torque resistant ultra light structures, with internal stowed stabilizators for the eventual rescue part of mission (first stage malfunction case). On aerodynamic fairing are situated the individual covers, or caps for better handling of stage igniters and other devices. The main support structure ring is situated on the perimeter of the combustion chamber of STAR 24 solid rocket motor of the third stage - all forces from stages weight, thrusts and payload are cumulated here. Material will be high-strength Titanium alloy and carbon/Kevlar composite structures. On the upper side of this ring the ERV adapter is attached. To this element some electrical systems, as a cables, electrical distributing elements and thermal batteries for stages control systems are connected. To this adapter the Earth Return Vehicle (ERV) is mounted by explosive joint (V-band). The aerodynamic fairing consist of 4 parts see Fig.27. Approximately in the middle of each part a hermetical biobarrier wall is attached, which divides the internal volume of this fairing to a biohazard area (upper part of fairing with the Earth Entry Capsule (EEC), and lower part with common sealing system. In the upper part of the shell two Sample Distributing Systems (SDS) together with Small Sterilization Chambers (SSC) take place (in mirrored position), which are responsible for ERV protection, to avoid the contamination of this Earth-return part with the possible Martian life (Fig.54).

39

JET VANES WITH SERVOS

SEPARATION ROCKET MOTORS STAR 3A (4x)

SECOND STAGE ADAPTER WITH STOWED STABILIZATORS

CLUSTER FROM STAR 13B SOLID ROCKET MOTORS II. STAGE OF MARV

SUPPORT, MOUNTING RING OF III.STAGE

ERV

ERV ADAPTER

Fig.23 Upper stages of MARV with solid motors (STAR series)

40

Fig. 24

Upper stages of MARV with opened stabilizators and without aerodynamic fairing (Flight plan B)

41

II. Stage

Fig.25

STAR 13B solid rocket motor with attached jet vane, gearbox and clutch.

The STAR 13B ( TE-M-763 )50 incorporates a titanium case developed for the STAR 13 for the propellant and nozzle design of an earlier STAR 13 apogee motor . The technical datasheet is in Appendix. As it was mentioned above, we will need longer burn time than predicted for this type of motor with lower average thrust (40s burn instead of 15-16s and thrust cca. 2,900 N, instead of 7,600 N). ATK has the experience to modify their basic motor designs and can design completely new motors at minimal risk to support specific flight applications49, 50. Of course, this configuration shall be optimal for given motor, therefore I decreased the specific impulse by about 5% in following trajectory calculations - see Flight Trajectory Analysis of MARV, variant B, Fig.36-40. This decreased performance is valid mainly for trajectory variant B, where II. Stage will operate from surface conditions to approx. 50-70 km height, so not in the ideal conditions of the vacuum of deep space. In this case, the vehicle will be in configuration with opened stabilizators, see Fig.35. This is needed, because although the second stage is fully controlled, after the 40s of burning a coast phase will follow (from end burn cca 4 km to 50-70km height). In this phase, e.g. to the ignition of the third stage these stabilizators have to stabilize the flight of the vehicle in atmosphere.

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Fig.26 Base structure with stabilizators

43

Fig.27 Compressed pictures about MARV stages separation and fairing drop down. Upper figure can represent a fairings drop down for A flight path, and next picture shows B flight variant. For simplicity, lower picture will be representative.

44

WEIGHT ANALYSIS OF SECOND STAGE OF MARVComponent Cabling Battery TVC Jet Vanes Gearbox Structure Pyrotechnics Separation Solid Rocket Motor STAR 3A Adapter for Sep. motor Base Structure Stabilizator Aerodynamic Fairing Sample Distributing Systems (SDS) Small Sterilization Chamber (SSC) Thermal Cover MLI STAR 13B solid rocket motor other Summary weight of II.Stage Units 1 1 4 4 4 8 4 Mass Unit (kg) 0,45 0,4 0,2 0,25 0,11 0,04 0,89 Total Mass (kg) 0,45 0,4 0,8 1,0 0,44 0,32 3,56 Mature Index (-) 4 4 6 6 6 4 2 Contingency (%) 15 15 30 30 30 25 5 Total Weight w/ Cont. (kg) 0,52 0,46 1,04 1,3 0,57 0,4 3,74 Size (mm) Vendor Comments

ATK

4 1 4 4 2 2

0,2 1,47 0,4 0,8 0,35 0,38

0,8 1,47 1,6 3,2 0,7 0,76

6 6 6 6 6 6

30 30 30 30 30 30

1,04 1,91 2,08 4,16 0,91 0,99

1 4

0,26 47,04

0,26 188,16

5 2

20 5

0,31 191,92 ATK

203,92

211,35

Table 5. Weights allocation of second Stage of MARV.

III. StageThird stage consist from STAR 2451,50 solid rocket motor- see App.Fig.51-52 , jet vanes control system, adapter and Earth Return Vehicle. In flight path type A the jet vanes control system will control the stage in whole burn time ( cca 30 s), because the flight path for direct return must be more precisious than flight type B (emergency mode), when the final low weight of stage with ERV is more important, than accuracy of orbit insertion. It is mean, that jet vanes control system (JVCS) will be functional in beginning phase of STAR 24 burn, stabilizing it, and finally ensuring the rotation of stage, before JVCS will be jetissoned. After ERV separation from empty stage, this one will canceling the rotate motion by own RCS system see the Fig.35, 28. The weight of third stage including ERV is 269,4 kg STAR 24 weight is 218 kg, ERV 44 kg, JVCS system 7kg, and adapter weight is 0,4kg.

45

Sample Distributing System (SDS) with Small Sterilization Chamber(SSC), as was mentioned, are attached on aerodynamic fairing, which will drop down before third stage ignition.The Earth Return Vehicle (ERV) will be jettisoned by compressed springs inside the adapter.III. STAGE FLIGHT CONFIGURATION (VAR.A,B) III.STAGE FLIGHT CONFIGURATION (VAR.B)

JET VANES CONTROL SYSTEM (JVCS) MODULE SEPARATION

ERV ADAPTER

EARTH RETURN VEHICLE (ERV) SEPARATION FROM EMPTY III.STAGE

NOZZLE REINFORCEMENT (DUE JVCS ADDING)

Fig.28 Third stage functions, from left to right direction: Stage burning, JVCS jetissoning, ERV separating.

46

Fig. 29 MARV`s last stage burning WEIGHT ANALYSIS OF THIRD STAGE OF MARVComponent Units Mass Unit (kg) Total Mass (kg) Mature Index (-) Contingency (%) Total Weight w/ Cont. (kg) 0,52 0,46 1,3 3,64 1,04 0,31 1,43 0,52 0,31 222,56 232,09 ATK Size (mm) Vendor Comments

Cabling 1 0,45 0,45 4 15 Battery 1 0,4 0,4 4 15 TVC Jet 4 0,25 1,0 6 30 Vanes Gearbox 4 0,7 2,8 6 30 Structure 4 0,2 0,8 6 30 Pyrotechnics 6 0,04 0,24 4 25 Support Ring 1 1,1 1,1 6 30 ERV adapter + 1 0,4 0,4 6 30 V-band Thermal 1 0,26 0,26 5 20 Cover MLI STAR 24 solid 1 218,2 218,2 1 2 rocket motor other Summary 225,65 weight I.Stage Table 6. Weights allocation of Third Stage of MARV.

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Guidance and Control of the MARV during ascent The guidance system for MARV shall bring this vehicle onto the desired direct hyperbolic path to Earth within a given accuracy (while minimizing the fuel consumption) and ensuring that the path constraints are well met during the flight. During the first stage function is vehicle designed for the 3-axes attitude control by the using the LITVC ( Liquid Injection Thrust Vector Control) in rocket motor divergent part of nozzle see the Fig.16-19. for the pitch and yaw will be employed the thrust vector control by NTO injecting to a divergent part of nozzle and roll will be control by the exhaust from gas generators system which is used for NTO tank pressurizing. The second and third stage will be managed by using the jet vanes in exhaust plumes, as was mentioned above (see Fig.23). The guidance may be based on a calculus of variations approach similar to that used for the final phase of the Apollo Earth return guidance56.

Earth Return Vehicle (ERV)EEC

CRUISE STAGE

Fig.30 Earth Return Vehicle consist from two parts Earth Entry Capsule (EEC) and Cruise Stage (CS).

Earth Entry Capsule (EEC)

The Mars Sample Return (MSR) mission will attempt to return the first samples from another planet. The final phase of the mission requires an Earth entry, descent, and landing capsule, which is responsible for transporting the samples safely through Earth`s atmosphere to a recoverable location on the surface.This capsule will has a two very important tasks: - isolating of samples from Earth contaminants, - eliminating the potential for terrestrial contamination.

48

The impact of these requirement on development and design of an Earth Entry Capsule is significant. The design of Earth Reentry Capsule is derived from60, 630 mm diameter with weight approximately 7 kg without parachute deceleration system. On account of its simplicity, the chuteless EEC is quite scaleable, decreasing in mass in nearly linear fashion with sample size this is very useful mainly for current project, because the overall system is limited by useful weight transporting by the lander. To a date lot of work has been done in this field57,58,59,, but unfortunately, the weight of this one is continually increased ( from weight in earliest design approx. 7kg to nearly 45 kg in current state-of-the-art). The samples, in such a design, are packaged in two hardened containers, in axisymetric position regarding to rotational axes of EEC, and surrounded by sufficient energy-absorbing material to limit dynamic loading during ground impact see Fig.31:

SAMPLE PLACE (2x) MAIN SPHERICAL CAPSULE

JOINTS W/ CRUISE STAGE (4x) 60 HALF-ANGLE SPHERE CONE FOREBODY

Fig.31 Passive Earth Entry Capsule The Cruise Stage consist from light-weight composite structure, solar panels, avionics, RCS hydrazine thrusters, low-gain antenna, rechargeable battery, Star Trackers, Sun Senzors, etc. see the fig. 37. Practically, this design is derived from Lockheed-Martin ISPP MSR project 60, with the same weight budget, but to current design is implemented a larger hydrazine fuel tank for 10 kg hydrazine. For Trajectory Correction Maneuvers (TCM) we will need approximately 200m/s velocity budget, and remainig 200-250 m/s will as a redundancy for if finest potential correction maneuver will be required after MARV functioning (the MARV consist mainly from solid rocket motors with relatively wide .spectrum of final velocities these must be corrected). Similar system has been done for PEGASUS small, winged launch vehicle system HAPS (Hydrazine Auxiliary Propulsion System) improves injection accuracy and increases payload capability79. The returning capsule (EEC) belongs to the most important part of the MSR mission it has to provide suitable environment for the Mars samples pressure of 6 to 10 Torr and

49

temperature of -10C ( or more, without refrigerator system) at least. It has to be absolute sterile in order to prevent any accidental contamination of the samples with microorganisms for the Earth (also at very low concentrations they might multiply themselves for 3-4 years). For these reason to have each sample hermetically closed right on the Mars surface. The minimum weight of the samples taken will be 0.25kg. As was mentioned above, in first stage malfunction case (emergency mode, flight variant B) the ERV can survive a couple of years on Mars orbit while waiting for the rescue mission. Because of this possible scenario the ERV contain a simple docking module for the possible Mars orbit rendezvous with some future small trans-Earth injection stage. The advantage of this concept is the exact ERV attitude measuring by onboard IMU and Star Trackers for orbit elements determination.TWO CLUSTER OF FOUR MR111C 4,5N THRUSTERS (AEROJET)

DEPLOYABLE SOLAR PANELS (2x)

ADAPTER RING

FIXED SOLAR ARRAYS DIAPHRAGM HYDRAZINE TANK (ATK 80444-1) FRAME COMPOSITE STRUCTURE

PASSIVE DOCKING MECHANISM (ON ROLL AXIS)

STAR TRACKERS (2x)

Fig. 32

Earth Return Vehicle configuration

During the flight of the capsule from the Mars to the Earth the monitoring of the inner pressure and temperature will be also necessary, so in case of having the capsule dehermetized, it will be possible to modify its trajectory in order to avoid the samples getting in touch with the Earths biosphere. The landing capsule will contain the following: container with the samples, radio transmitter, telemetries, batteries (for one use, with chemical activation), plus radio and light beacon. The possible landing areas will be relatively flat regions where the population density is lower with large opened areas, low temperature (in case of surface contamination the micro organisms will have smaller chance to spread away) and on the other hand if the search for the

50

samples would take longer as predicted, the low temperatures are also more suitable to store the samples. The low temperature has its further advantages too for example because of the higher air density, the free fall of the capsule will be decelerated more intensively by the air the capsule will have lower impact speed. The snow cover might also soften the impact. On the other hand the environmental and meteorological factors may hinder the search for the capsule too. The next important requirement is good tracking possibilities with optical or radar instruments, respectively.

ERV ALLOCATED MASS Total Total Weight w/ Component Mass (kg) Cont. (kg) Power and Cabling 4,546 5,615 Telecom 3,664 4,489 Avionics 2,990 3,273 C&DH 0,404 0,485 Thermal Control 0,8 0,916 Structures, 5,96 6,907 Mechanisms Propulsion 5,039 5,406 Earth Entry Capsule 5,698 6,657 (EEC) GHe pressurant 0,17 0,18 N2H4 8,5 10 Summary weight of 37,77 43,928 ERV

Table 7. Weights allocation of ERV60

Mars Ascent and Return Vehicle Flight AnalysisThe basic reference aerodynamic shape for the Mars Ascent and Return Vehicle (MARV) is dictated primarily by packaging and volumetric constraints imposed by the MSL - like aeroshell. This three stage launch vehicle is compound inside the aeroshell in a lying position. Due the fact, that simply drag coefficient depend on a Mach number of flight, in this first approach I calculated the trajectories without using of angle-of-attack related to trajectory path only with single drag coefficients from CD= 0,5, 0,6, 0,7 ,0,8, 0,9 and 1,0 were taken as a probanly value for zero angle of attack. The trajectory shape is given only by initial elevation angle of the MARV lift-off without active control and only for first stage burning ( there are two variant of trajectories A and B). This trajectory simulation is based on numerically integrating the equations of motion. The Mars atmospheric density model is based on Mars-GRAM (Clancy and hydrostatic equilibrium models) 54 and Mars Pathfinder measured atmospheric density data55 see Fig.33. Only B trajectory type was investigated as a requirement of ERV delivering to Martian Orbit (MO) and for first stage only( drag, etc.).

51

0.02 0.018 0.016 0.014 Density [kg/m3] 0.012 0.01 0.008 0.006 0.004 0.002 0 0 10 20 30 40 50 60 70 80 90 100 Altitude H [km] y = 0.0303e-0.1216x R2 = 0.9963 atm[MPF] atm[Clancy] atm[GRAM] Exponenciln (atm[GRAM])

Fig. 33 Mars atmosphere density related to altitude model. The insisted equation on this figure is more strict to atmospheric density in low altitude (05km) but it is on the safely side of next trajectory calculations. The stage masses of MARV are determined through geometrical and maximum LOX mass requirements related to a strict volume utilizing in this concept from this point of view distribution of stage masses are not optimal(to optimal trajectory). It would be need some optimalizations in the future work. Mass and rocket motor properties of MARV are in Table 8. Stage 1 dry mass [kg] Stage 1 propellant mass [kg] Stage 2 dry mass [kg] Stage 2 propellant mass [kg] Stage 3 dry mass [kg] Stage 3 propellant mass [kg] ERC+cruise stage mass (ERV) [kg] Total MARV lift-off mass [kg] Stage 1 average engine Isp [Ns/kg] Stage 2 average engine Isp [Ns/kg] Stage 3 average engine Isp [Ns/kg] Stage 1 time of function [s] Stage 2 time of function [s] Stage 3 time of function [s] Stage 1 average engine Thrust [N] Stage 2 average engine Thrust [N] Stage 3 average engine Thrust [N] Referential MARV cross area [m2] MARV characteristical velocity [m/s] 49,01 560 47,35 164 32,09 200 44 1096,45 3290 2796 2775 100 40 30 17,920 9929 19,657 1.0 7091,1(+200 m/s ERV)

Table 8. Mass and rocket motor properties of MARV stages. 52

Third Stage Ignition and burning.

ERV is jetissoned and fly away to Earth

Coast phase. The III.stage is stabilized by ERV`s hydrazine RCS system After burnout, the fairings are dropped down with empty II. Stage. ERV RCS system will be activated The II.Stage undelayed ignition and burning to T+140 s.

T + 100s, II. Stage is jetissoned from empty I. Stage of MARV

MARV Lift-Off, I. Stage is burning

Fig. 34.

A type trajectory Direct Return to Earth.

Third Stage Ignition and burning .

ERV is jetissoned and fly away from empty stage. After cca 1700 second, ERV`s RCS will circularize the orbit around Mars

Approx. in T+200 s empty II.stage is jettisoned( in altitude from 50 to 70 km), and last stage will be stabilizing by own ERV`s RCS system.

Due the densier atmosphere, the II..stage will use the stabilizators for stabilizing the ascent vehicle after burnout (in coast phase).

Stabilizators are in the opened position immediately after lift-off. II. Stage is activated. At first, II.stage will be separate from an unfunctional I. stage

Fig. 35.

Emergency, B type trajectory Mars Orbit Insertion (MOI).

53

Velocity of MARV second Stage (B)900

800

700

600

Velocity [m/s]

500

400

300

200

100

0 0 20 40 60 80 100 Time [s] 120 140 160 180 200

Figure 36

MARV second stage ascent velocities with various elevation angle of this stage lift-off and drag coefficient ( CD varies from 0,5 to 1,0 ). It has visible, that these factors have a lower incidence for first, rough calculation on final velocity of this stage.

54

Acceleration of MARV second Stage (B)10 8 6 4 2 0 0 -2 -4 Time [s] 20 40 60 80 100 120 140 160 180 200

Figure 37

Acceleration [1/s.s]

MARV second stage acceleration with various elevation angles of lift-offs and drag coefficients.Drag Force of MARV second Stage (B)

3500

3000

2500

Drag Force [N]

2000

1500

1000

500

0 0 20 40 60 80 100 Time [s] 120 140 160 180 200

Figure 38

Drag force, with various elevation angle of lift-off and drag coefficients.

55

Y-Altitude of MARV second stage (B) for various Drag coefficients and elevations angle (from 60 to 85 deg)100000 90000 80000 70000 Altitude [m] 60000 50000 40000 30000 20000 10000 0 0 20 40 60 80 100 Time [s] 120 140 160 180 200

Figure 39

Altitudes changes with various elevation angle of lift-off and drag coefficients.

Attitude of MARV second Stage for B trajectory for various elevation angle (from 60 to 85 deg) in first 200 sec flight.

90000 80000 70000 Altitude [m] 60000 50000 40000 30000 20000 10000 0 0 20000 40000 60000 80000 100000 Downrange x [m]