Chapter 5 Crew Return Vehicle

Chapter 5 – Crew Return Vehicle

5Crew Return Vehicle

5.1Overview

5.1.1Abstract – 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.

5.1.2Introduction – 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.

5.1.3Background – 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 Homeroccurs.

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.

5.2Human 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 5.1below 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.[1] 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.

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.5.1displays a CATIA model shown from a top view of the crew return vehicle and its human factors components.[2] A bottom view of the components can be seen in Fig.5.22

Fig.5.1Crew Return Vehicle layout top view2

Fig.5.2Crew Return Vehicle layout bottom view2

5.3Design – 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

5.3.1Exterior – Kenneth Kirk

The modified CRV design is based off the Apollo capsule using the geometry taken from ref.[3] in the appendix. In Fig. 5.3below, the modified CRV dimensions are listed.

5.3.2Interior – 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 [4], which may be seen below in Fig. 5.5.

5.3.3Design 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. 5.6is 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.

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 5.3below for each of the capsule sections and as a whole.

Table 5.3CRV 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.[5]. The breakdown for the entire component masses are in Table 5.4below.

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.

5.4Aerodynamics

5.4.1Parachute 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.[6] We choose a conical shape for the main parachutes because of the shape’s high drag coefficient, which decreases the required parachute area.

5.4.1.1Parachute 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.

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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).[7] 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.

5.4.1.2Parachute 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).[8],[9]

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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.

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5.4.2Stability- 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.[10] 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.[11] 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 thecenter 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.

Static margin is determined using Equation 1-1below. The equation gives static margin as a function of center of gravity and angle of attack in a percentage value.[12] The static margin must be positive (+) for stability.

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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.