1 - Purdue University



Crew Return Vehicle

1 Overview

1 Abstract – Kenneth Kirk

The purpose of this section is to take an in depth look at the Crew Return Vehicle (CRV) for our space mission to Mars. In order to achieve this purpose, the aerodynamics team looked into the following areas: human factors, design (interior & exterior), aerodynamics (parachutes, stability, and thermodynamics), power, and the launch escape system (LES). The two main decision factors were to choose between an aeroshell and a winged orbital space plane (OSP). The main difference between the two is that the chosen OSP has wings and the capsule doesn’t, nor is it a lifting body. After deciding that an aeroshell is the way to go, we then began to run analysis on what type of capsule to use in order to return the crew back to Earth. A modified version of the Apollo capsule is the design choice chosen, in which a complete analysis on the entire CRV mission began. The analysis began by the basis of the human factors requirements, since the human’s lives are the most prominent area of interest.

2 Introduction – Kenneth Kirk

The main focus of this section is to completely design a crew return vehicle to finish off the mission to Mars. A CRV is a vehicle that returns a crew as the title signifies, which in this mission is back to Earth. This vehicle also serves as a space ambulance and a lifeboat for the crew. There is no definite answer on what is the most optimal design (ie. whether it is a winged vehicle, a lifting body, or a capsule), but a completely modified Apollo capsule has the most appeal.

3 Background – Kenneth Kirk

Orbital Space Planes of both the wing and capsule type have been used for years as a viable transport vehicle. The question is, is it better to use a capsule or a winged vehicle for the crew return vehicle? The answer to that question may be unknown, but the lower cost, the simple design modifications, and the landing space are among the many wonderful aspects that make the capsule design captivating.

To explain this in further detail, the lower cost is a direct result of having many capsule designs already created that only needs interior design modifications and updates in the materials. The low cost also stems from not having to redesign the capsule itself as a whole. A capsule is proven in its reliability and likelihood of returning the crew, whereas if we choose to use a winged vehicle, the entire OSP needs redesigned in order to make sure that it is absolutely safe and reusable. According to much research many believe that a fully safe and reusable winged spacecraft could take 15-20 years, which is out of the scope of this project timeline in which the mission Homer occurs.

The design modifications are simple since the capsule only has to be re-sized based on the mission and the human factor constraints for the amount of time in which the humans will spend inside the CRV. The new technological advances alone from the 1960’s till now in the computer department will cut down the mass by a fourth from 500 kg to 130 kg. Such mass reductions will also cut down the cost of the entire crew return vehicle.

The vast landing site allows for the capsule design to land anywhere along the coastal areas in the Pacific and Atlantic Ocean. Although the capsule has an enormous landing area, this provides a slight problem for the recovery teams to know the exact location of the capsule in these large masses of water. The capsule design is also promising due to its natural ability to stay afloat due to its design, the crew has enough consumable storage space, and there is recovery equipment as well. A winged vehicle, although, may not necessarily have a place to land in case of an emergency situation if the landing strips are booked with activity from normal day-to-day aircraft. The hazard of no place to land causes major problems since there are a limited number of sites where this winged vehicle may land.

2 Human Factors- Rebecca Karnes

The Crew Return Vehicle (CRV) will serve the purpose of transferring the crew from the transport module back to Earth. The following section contains details about the human factors component included in the vehicle. Displayed in Table 5.1 below is a listing of the CRV payload components, their single masses, quantities, total masses, and volumes. We have included items from crew seats and personals, to the guidance control computer and Martian soil sample. Due to the fact that the crew will only occupy the vehicle for a brief time, the consumable mass is very low unlike the habitation module. The second nominal mass is that of the Martian sample. Since only 1 kg of actual soil will be returned in the sample return vehicle the mass is only slightly greater due to storage and preservation equipment. The crew was limited by the human factors team to only bring a mass of 15 kg of personal belongings for the entire mission. Values for the seats, consumables, and guidance control computer were all taken from the Human Spaceflight book. [i] Historical weights for Apollo return spaceflight suits were used to determine mass of space suits. The total mass of the human factors components in the CRV comes to 689.4 kg with a total volume of 22.26 m3.

[pic]

Placement of the CRV components can have a drastic effect on the stability of the vehicle upon re-entry, as later discussed in the aerodynamics stability section. Therefore, it is necessary to determine an acceptable layout of the human factors components placed into the CRV. Fig. 5.1 displays a CATIA model shown from a top view of the crew return vehicle and its human factors components.[ii] A bottom view of the components can be seen in Fig. 5.22

[pic]

Fig. 5.1 Crew Return Vehicle layout top view2

[pic]

Fig. 5.2 Crew Return Vehicle layout bottom view2

3 Design – Kenneth Kirk

Nomenclature

R = base radius

rs = bottom radius

rf = nose radius

b = height

θ = half angle

In order to achieve the overall mission, research and many different analyses on the CRV were performed in the following areas:

• Aerodynamic Stability

• Earth Descent and Landing

• Entry vehicle design (interior and exterior)

• Failure modes for parachutes and launch abort

1 Exterior – Kenneth Kirk

The modified CRV design is based off the Apollo capsule using the geometry taken from ref.[iii] in the appendix. In Fig. 5.3 below, the modified CRV dimensions are listed.

2 Interior – Kenneth Kirk

The modified configuration drawings in fig. show the internal components and the correct corresponding volumes. From bottom to top, the green box stands for all of the scientific materials. The orange box includes all of the Martian samples that the rover collects and the brown boxes represent the crew personal equipment. The round object above the seats is the main computer control system.

The basic configuration was modified from the configuration drawings in ref [iv], which may be seen below in Fig. 5.5.

3 Design Analysis – Kenneth Kirk

In order to get the general shape of the crew return vehicle we defined a few capsule sizing factors in order to obtain the necessary geometry. Table 5.2 contains the dimensional constraints for the following areas of interest: the half angle, height, max radius, base corner radius, and nose cone radius. In this table, the measurements were based off constraints in reference to each other. For example, if the height of the Apollo modified design was at 80 ft, then the base corner radius is 6 ft and the nose cone radius is 5 ft due to the constraints. These dimensional constraints are the optimum for each design through CFD analysis as may be seen in appendix A, and historical wind tunnel data (ref. 3).

Table 5.2: CRV dimensional constraints for sizing

|CRV Sizing Constraints |

| |Apollo |Biconic |Ellipsled |

|Angle (°) |20-33 |0-15 |0-10 |

|Height |.8-1.85 |2.85-5.9 |2.6-6.0 |

|Body Radius |1.195 |1.21 |1.25 |

|Base Radius |.05-.12 |.8-1.88 |.75-1.9 |

|Nose Radius |.05-.25 |.75-2.38 |.75-2.41 |

In order to visualize the table and the explanation above, Fig. 5.6 is dimensionally constrained for each property. The values are as follows: rs = base radius, R = body radius, rf = nose radius, b = height and theta = half angle of capsule.

[pic]

Fig. 5.6: CRV sizing constraints

After the sizings are set, calculations of the volumes and the surface areas of each segment were performed in order to achieve the optimal design based on the human factors constraints and that of the mission. For further accuracy, the entire drawing is modeled in CATIA, for more accurate volumes and surface areas. These values are available in Table 5.3 below for each of the capsule sections and as a whole.

Table 5.3 CRV Dimensions for Volume and Surface Area

|CRV Dimensions |

| |Volume |Surface Area |

|Body |12.3 m3 |12.6 m3 |

|Base |2.4 m3 |2.5 m3 |

|Total |14.7 m3 |15.1 m3 |

Now that the complete exterior is finished, the interior components are organized inside the capsule where precise mass measurements are also determined based on the room available inside the modified crew return vehicle. In order to calculate masses such as the control and recovery systems, these masses are just taken from the Apollo systems from ref.[v]. The breakdown for the entire component masses are in Table 5.4 below.

[pic]

Upon completion of the entire mass distribution layout throughout the CRV, important features such as finding the center of gravity, cg, and the best entry angle were taken into account. The cg location from our Catia file is at .5 meters from the base of the CRV found from the CRV Catia file.

4 Aerodynamics

1 Parachute Recovery System – Heather Dunn

Nomenclature

CD = parachute drag coefficient

Cx = opening force coefficient

d = diameter of parachute

Fx = peak opening force

g = gravity constant

ls = length of suspension lines

M = Mach number

mcrv = mass of Crew Return Vehicle

mpar = mass of parachute

q = dynamic pressure

S = surface area of parachute

Vpar = volume of parachute packed

vt = terminal velocity

Wc = mass of canopy materials per unit area

Wl = mass of suspension lines per unit length

X1 = opening-force-reduction factor

z = number of suspension lines

ρ = air density

ρpack = parachute packing density

Since the vehicle that will return the crew to Earth is blunt rather than winged, a water landing is necessary. There are many different types of recovery systems that are suitable for water landings including parachutes, parafoils, ballutes, and reverse thrusters. Since the driving force in this mission is low mass, we choose a parachute recovery system because parachutes typically have lower masses than other aerodynamic deceleration devices.

The parachute system consists of two drogue chutes and three main parachutes. Since main parachutes cannot deploy at high speeds, we use drogue chutes for the initial deceleration because they characteristically have high strength, good stability, and small opening forces. We choose a ribbon shape for the drogues because this shape gives high drag, low angle of oscillation, and small opening forces.[vi] We choose a conical shape for the main parachutes because of the shape’s high drag coefficient, which decreases the required parachute area.

1 Parachute Sizing – Heather Dunn

The Crew Return Vehicle (CRV) enters the Earth’s atmosphere at a velocity of approximately Mach 30. Because density increases as altitude decreases, the increasing air resistance during descent causes the capsule to decelerate. Drogues may be deployed at velocities less than 300 m/s to stabilize and decelerate the vehicle. In our case, the drogues deploy when the capsule slows to 135 m/s at an altitude of 7,000 m. The drogues need to slow the 6,993 kg vehicle to 80 m/s, allowing for main parachute deployment. The total area required for this deceleration is calculated from Eq. 5–1.

|[pic] |5–1 |

When the capsule reaches the terminal velocity of 80 m/s at approximately 3,000 m altitude, the main parachutes deploy. For water landings, we need an impact deceleration less than six times the Earth’s gravitational force (6g).[vii] Historically, landing speeds for a three-person spacecraft are less than 9 m/s to achieve tolerable impact deceleration. Because the return capsule for this mission is larger than return capsules from earlier missions, we limit landing speed to 8 m/s. However, parachute systems are designed with redundancy so that if one parachute fails, the vehicle lands safely with only two functioning parachutes. We set the landing speed to 6.5 m/s for a three-parachute landing, ensuring that the landing speed for a two-parachute landing is less than 8 m/s. With the parachute area from Table 5.5, the vehicle lands safely at 7.96 m/s with only two functioning parachutes. Fig. 5.7 shows how the required parachute area varies with different vehicle masses and landing velocities. Fig. 5.8 and Fig. 5.9 show altitude vs. time profiles for the parachute deployment sequence and landing.

[pic]

[pic]

[pic]

2 Parachute Material, Mass, and Volume – Heather Dunn

The parachutes are composed of a nylon canopy and Kevlar suspension lines. The required suspension line length and number of suspension lines is proportional to the diameter of the parachutes. From the canopy size, number and length of suspension lines, and material properties (Table 5.6), the mass of the total parachute system is 1,148 kg (Eq. 5–2) and the total volume required to store the parachute recovery system is 1.83 m3 (Eq. 5–3).[viii],[ix]

|[pic] |5–2 |

|[pic] |5–3 |

Parachute deployment generates considerable forces that may be critical to mission success. For example, if opening forces exceed the allowable forces of the parachute canopy and suspension line materials, the canopy may be punctured or ripped from the suspension lines, causing the vehicle to land at a catastrophic speed. The opening force created during drogue opening is 66,667 N and the opening force created during main parachute opening is 221,406 N (Eq. 5–4). Similar parachutes used for aircraft deceleration are able to withstand opening forces as large as 405,000 N.7 Therefore, the opening forces generated by our parachutes appear small enough to avoid material failure. Refer to Appendix A for further information on calculations relevant to the parachute recovery system.

|[pic] |5–4 |

2 Stability- Rebecca Karnes

Nomenclature

Xcg = center of gravity along axis

Xref = reference center of gravity

xMRC = x location at moment reference center

zMRC = z location at moment reference center

Lref = reference length

Cmref = pitching moment at the reference CG

Cm = pitching moment coefficient

Cm0 = pitching moment coefficient at moment reference center

CN = ormal force coefficient

a = angle of attack

Longitudinal stability analysis was performed on aerodynamic data of the crew return vehicle to insure the safe homecoming of the crew. It is essential to determine if the vehicle will return to its desired trim angle of attack if it is disturbed.[x] The first step in this analysis was to obtain the aerodynamic data for the selected geometry of the crew return vehicle; this was done using historical Apollo data.[xi] The first major stability analysis was done on the static margin in order to obtain a measure of longitudinal stability. Static margin is determined by how far the center of gravity (cg) is from the neutral point, along vehicle axis. Fig. 5.10 below gives an Apollo capsule module to display the locations of cg and neutral point used to determine static margin. The Fig. also show how the following sections will be referencing forward and aft of the Crew Return Vehicle. The neutral point or aerodynamic center is the zero pitching moment location along the axis, or the location where the pitching moment is independent of angle of attack. A perturbation of the vehicle will result in the vehicle diverging away from trim conditions if the center or gravity location is in front of the neutral point and could lead to tumbling flight. An aft center of gravity location will insure a dampening effect and a disturbance will result in the vehicle resuming towards its trim angle of attack.

[pic]

Static margin is determined using Equation 1-1 below. The equation gives static margin as a function of center of gravity and angle of attack in a percentage value.[xii] The static margin must be positive (+) for stability.

|[pic] |5–5 |

[pic]

Worst case static margin occurs at a trim angle of attack of zero degrees, while the largest static margin values will be found with alpha of 180 degrees. Fig. 5.11 displays the entry orientation of the CRV and the reference angle used. Initial stability analysis is performed in order to determine a desired location or aim for the CRV center of gravity. The placement and layout payload components are determined in order to obtain the most aft center of gravity as possible. The CRV center of gravity is then determined to lie at 0.5 m from the base of the vehicle and a z-axis offset of approximately 0.1% of the length. Analysis is done using aerodynamic data at various angles of attack ranging from 0 to 180 degrees and a speeds ranging from Mach 0.2 to 29.5. Particular attention is paid to super sonic speeds as there is a large difference in aerodynamic properties at subsonic speeds. Details of the stability analysis code can be found in the Matlab code in Appendix A. Fig. 5.12 below shows a plot of the static margin as a percent verse the center of gravity location as a percent of the body length. The plot is for a speed of Mach 29.5 and for angles of attack ranging from 110 to 180 degrees. A red dotted line is drawn at the location of the determined center of gravity. A blue dashed line is placed across the zero static margin axis. This displays the limiting cg locations at the various trim angles and shows that our trim angle must be greater than 115 degrees. A comparison on the effect that speed has on the static margin can be seen by looking at static margin plots for speeds of Mach 4 and 10 in the Appendix A.

[pic]

A plot of the static margin values for the various Mach numbers ranging from 0.2 to 29.5 can be seen below at the angles of attack of 140, 145, and 150 degrees Fig. 5.13. Upon re-entry the CRV will decelerate to a desired speed before drogue parachute and main parachute deployment. It is necessary to determine if the CRV will remain stable through all the varying speeds to ensure crew safety. With the current cg location we have a positive static margin or stable vehicle conditions for the range of speeds that will be critical during re-entry. The plot shows static margin values ranging from 25% to 50%. . Historically static margin values for stable airplanes range from 5% to 40%.

Trim lines are used to represent the cg location at each angle of attack that neutralizes the pitching moment. 12 Equation 1-2 below displays how the plot of xcg vs. zcg determines. The longitudinal stability trim lines.

|[pic] |5–6 |

By setting the moment coefficient equal to zero using an array of xCG points, zCG points are generated at the various angles of attack. Also plotted on Fig. 5.14 below are the points at which static margin becomes unstable. These static margin zero location points place a constraint on the possible cg locations for the speed of Mach 29.5 at the desired trim angles. The important contribution of the z-axis offset can be seen by examining the plot. A cross is also plotted in Fig. 5.14 below to show that the actually determined cg location lies within the bounds set by the static margin and the trim conditions, therefore proving the stability of the crew return vehicle.[xiii]

[pic]

3 CRV Heat Shield – Matthew Branson

Empirical heat methods was used for the CRV heat shield design.[xiv],[xv] Table 5.7 shows the thicknesses of the heat shield.

[pic]

An ablator material was used due to the large amount of heat flux that needs to be dissipated. A Graphite ablator, similar to the ablation material on the Mars Lander, is used. Both carbon foam and glass polyimide honeycomb is used to reduce the mass and keep the inside of the CRV at 300 Kelvin.

5 Power - Ryan Spalding

We choose batteries to provide the power requirements for the crew return vehicle. With the energy density of batteries constantly improving, we have the option of using anticipated performance numbers of future technologies. However, in the interest of presenting a conservative design, we instead decide on the most energy dense type of battery currently in production: lithium thionyl chloride batteries. The choice of the particular battery model of ER14250H as the most energy dense (480 W-hr/kg) is based on the total power capacity of 24 kW-hr needed for the CRV. This power capacity allows for 24 hours of operation at a power level of 1 kW, a level sufficient to run all life support, navigation and communication equipment on the CRV. To meet the specific current and voltage demands of each piece of equipment, the batteries are clustered in series or parallel as needed.

Each ER14250H battery has a nominal voltage of 3.6 V and a nominal current capacity of 1.2 A-hr, giving a total power capacity of 4.32 W-hr.[xvi] To meet the power capacity demands of the CRV, we employ 5600 batteries. This number includes a 1% margin which allows for the small leakage that occurs over time. Each battery has a cylindrical volume of 4.1 cm3 and a mass of 9 grams, giving a total mass and volume for the system of 50.5 kg and 30,000 cm3 respectively.16 These numbers are summarized in Table 5.8. See Appendix E for detailed calculation procedures.

[pic]

6 Launch Escape System – Heather Dunn

One of the most life-threatening aspects of the mission occurs at launch. While on the launch pad or shortly after launch, an exploding rocket or malfunctioning control system may jeopardize not only the success of the mission, but the lives of the astronauts on board. Therefore, we have abort scenario procedures in place to avoid such catastrophic events as loss of life. In the event of an emergency on the launch pad or shortly after launch, a Launch Escape System (LES) carries the CRV away from the launch vehicle. The LES consists of five major parts: Launch Escape Tower, Launch Escape Motor, Tower Jettison Motor, Pitch Control Motor, and Boost Protective Cover.[xvii]

The Launch Escape Tower is a truss-like structure that is positioned above the CRV as shown in Fig. 5.15. The tower is constructed of 65 mm and 90 mm diameter titanium tubes and is attached to the CRV with explosive bolts. The cylindrical structure above the tower houses three solid-propellant rocket motors. The Launch Escape Motor, located in the bottom of the rocket cylinder, is the largest of the three motors and is responsible for carrying the CRV away from the launch vehicle during an emergency situation. The motor is similar to the Little Joe and Little Joe II motors that were used in the Mercury and Apollo missions and provides about 655,000 N of thrust over a burning time of 4 seconds. The Pitch Control Motor is mounted at the top of the rocket structure and produces 11,000 N of thrust in about 0.5 seconds. This motor provides the initial pitch maneuver we need to change the flight course during abort. The Tower Jettison Motor is located in the middle of the rocket structure. This motor jettisons the Launch Escape Tower and the Boost Protective Cover after an abort or after the first stage when the tower is no longer needed. This motor produces 140,000 N of thrust. The Boost Protective Cover is a honeycomb structure that covers the entire CRV to protect the external surfaces from the heat generated by the Launch Escape Motor during abort. Relevant dimensions of the LES are shown in Fig. 5.16. Table 5.9 shows the mass breakdown of the entire Launch Escape System.

1 Pad Abort – Heather Dunn

There are two main abort scenarios that require the LES. If a rocket engine explodes on the launch pad, we need to perform a pad abort. During a pad abort, the Launch Escape and Pitch Control Motors fire to bring the CRV up and away from the launch vehicle to an altitude of approximately 1200 m. Once the CRV reaches this height, the tower jettison motor fires and the explosive bolts holding the LES to the CRV explode, separating the LES from the CRV. Once the CRV is free, the parachute recovery system operates as in normal reentry. Fig. 5.17 shows the descent of the CRV during this scenario

2 Return to Launch Site Abort – Heather Dunn

The second abort mode for the LES is the Return to Launch Site (RTLS) abort. An RTLS abort can be performed anytime after launch until the first stage rocket reaches the end of its burning time. An RTLS abort is initiated when thrust from an engine is lost or when pitch, roll, or yaw angle go beyond a maximum limit. Once RTLS abort is initiated, the three motors perform as in pad abort and the parachutes deploy for a safe landing. The descent of the capsule during an RTLS abort scenario is similar to the reentry and descent regime discussed in Section 5.4.2. If no abort is initiated before the first stage reaches its maximum height, the LES is automatically jettisoned from the CRV. After this time, aborts may be performed by following procedures similar to normal reentry.

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[i] Larson, Wiley J. and Linda Pranke Hamilton. Human Spaceflight: Mission Analysis and Design. McGraw-Hill Companies Inc. Crawfordsville. 1999.

[ii] Kenneth Kirk, CATIA Model of Crew Return Vehicle, Purdue University Senior Design Course Project 2004

[iii] Romere, Paul O.; and Hillje, Ernest R.: Apollo Comand Module Aerodynamics for OTV Studies. Lyndon B. Johnson Space Center, April 5, 1983.

[iv] NASA History Office, Apollo Configuration Drawings.

[v] Mass Components of the Apollo.

[vi] Kiker, John W. and Carlisle C. Cambell, Jr. “Spacecraft Landing Systems – Design Criteria and Components.” Manned Spacecraft: Engineering Design and Operation, pp. 327-341. New York: Fairchild Publications, 1964.

[vii] Knacke, T.W., Parachute Recovery Systems Design Manual, Chaps. 5, 6, 1st ed., Para Publishing, Santa Barbara, 1992

[viii] Parachute Branch, Equipment Laboratory, Engineering Division, Air Materiel Command, United States Air Force Parachute Handbook, Sections 4-6, Springfield, OH: Carpenter Litho and Printing Co., 1951.

[ix] Brown, William D., “Parachutes”. London: Pitman, 1951.

[x] “Aerodynamic Stability Analysis.” Sample Collection for Investigation of Mars (SCIM)

[xi] Hillje, Ernest R. and Romere, Paul O. “Apollo Command Module Aerodynamics for OTV Studies.” 5 April

1983.

[xii] “NASA OSP Gum Drop Shape: Preliminary Aerodynamic Stability Study” Rea, Jeremy and Whitley, Ryan. 31 January 2003

[xiii] “The Apollo Project.” Ball, Marcel and Bartlett, Adrian. 23 January 2004

[xiv]

[xv] Charles D. Brown, Elements of Spacecraft Design, AIAA Education Series, Castle Rock, CO, 2002

[xvi] Power Stream. lip.htm. 2003.

[xvii] Duncan, J., “Launch Escape”, , August 2002.

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Table 5.4 CRV mass component factors

|CRV Mass |

|Communications |100 kg |Personal Storage |60 kg |

|Computer System |130 kg |Power |50.5 kg |

|Control Systems |600 kg |Propellant |80 kg |

|Consumables |10 kg |Recovery Equipment |250 kg |

|Crew |280 kg |Scientific |15 kg |

|Heat Shield |850 kg |Seats & Provisions |440 kg |

|Navigation |500 kg |Space Suits |44 kg |

|Martian Sample |10 kg |Structure |1,600 kg |

|Misc Contingency |200 kg |Telemetry |200 kg |

| | |Total Mass |5,400 kg |

Crew Personals

Martian Sample

[pic]

Fig. 5.5 Apollo Interior Configuration

Table 5.1 CRV Human Factors Components

|Component |Mass (kg) |# |Total Mass (kg) |Total Vol (m3) |

|Seats |35 |4 |140 |10 |

|Consumables |2.6 |4 |10.4 |0.03 |

|Crew Personal |15 |4 |60 |0.16 |

|Scientific |15 |1 |15 |0.04 |

|Martian Sample |5 |2 |10 |0.03 |

|Guidance Control Computer |130 |1 |130 |2 |

|Crew |70 |4 |280 |10 |

|Return/Launch Space Suits |11 |4 |44 |Included with Crew |

|Total |283.6 |24 |689.4 |22.26 |

Table 5.5 Parachute System Properties

|Component |Property |Value |

|Drogues | | |

| |Number |2 |

| |Diameter (each) [m] |5.5 |

| |Area (each) [m2] |23.5 |

| | | |

|Main Parachutes | | |

| |Number |3 |

| |Diameter (each) [m] |33.6 |

| |Area (each) [m2] |888.0 |

[pic]

Fig. 5.7 Parachute Area vs. Vehicle Mass and Landing Velocity

[pic]

Fig. 5.8 Altitude vs. Time for Earth Descent and Landing

[pic]

Fig. 5.9 Earth Descent and Landing Timeline

Table 5.6 Parachute Mass, Volume, and Material Properties

| |Parameter (per each drogue) | |

|Graphite Ablator |0.001 |30 |

|Si-C Composite[xviii] |0.0012 |115 |

|Carbon Foam |0.12 |560 |

|C-C Composite |0.001 |270 |

Table 5.8 Summary of CRV Battery System Mass and Volume Numbers

|Lithium Thionyl Chloride Battery: |Mass (kg) |Volume (cm3) |Nominal Voltage (V) |Nominal Current Capacity |Power Capacity (kW-hr) |

|ER14250H Model | | | |(A-hr) | |

|Each Battery |0.009 |4.1 |3.6 |1.2 |0.0043 |

|Total (5600 Batteries) |50.5 |30000 |3.6 |6720 |24.2 |

[pic]

Fig. 5.-‡ˆ SEQ Figure \* ARABIC \s 1 15 LES Configuration

[pic]

Fig. 5.16 LES Dimensions

Table 5.9 LES Mass Breakdown

|Component |Mass (kg) |

|Launch Escape Tower |517 |

|Launch Escape Motor |2132 |

|Boost Protective Cover |430 |

|Pitch Control Motor |23 |

|Tower Jettison Motor |50 |

|Total |3152 |

[pic]

Fig. 5.17 Pad Abort Descent Profile

[pic]

Fig. 5.3 Dimensional Constraints for modified Crew Return Vehicle

[pic]

Fig. 5.4 Dimensional Constraints for modified Crew Return Vehicle

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