1 

NASA/TM—2014–218366 

 

Conceptual Launch Vehicle and 

Spacecract Design for Risk Assessment 

Samira A. Motiwala 

Stanford University, Stanford, CA, 94305 

Donovan L. Mathias 

NASA Ames Research Center, Moffett Field, CA 94035 

Christopher J. Mattenberger 

Science and Technology Corporation, Moffett Field, CA 94035 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

June 2014
NASA STI Program ... in Profile 

Since its founding, NASA has been dedicated 

to the advancement of aeronautics and space science. 

The NASA scientific and technical information (STI) 

program plays a key part in helping NASA maintain 

this important role. 

The NASA STI program operates under the auspices 

of the Agency Chief Information Officer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NTRS Registered and its public interface, the 

NASA Technical Reports Server, thus providing one 

of the largest collections of aeronautical and space 

science STI in the world. Results are published in both 

non-NASA channels and by NASA in the NASA STI 

Report Series, which includes the following report 

types: 

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase of 

research that present the results of NASA 

Programs and include extensive data or theoretical 

analysis. Includes compilations of significant scientific and technical data 

and information deemed to be of continuing 

reference value. NASA counter-part of peerreviewed formal professional papers but has less 

stringent limitations on manuscript length and 

extent of graphic presentations. 

• TECHNICAL MEMORANDUM. 

Scientific and technical findings that are 

preliminary or of specialized interest, 

e.g., quick release reports, working 

papers, and bibliographies that contain minimal 

annotation. Does not contain extensive analysis. 

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees. 

• CONFERENCE PUBLICATION. 

Collected papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or 

co-sponsored by NASA. 

• SPECIAL PUBLICATION. 

Scientific, technical, or historical information 

from NASA programs, projects, and missions, 

often concerned with subjects having substantial 

public interest. 

• TECHNICAL TRANSLATION. 

English-language translations of foreign 

scientific and technical material pertinent to 

NASA’s mission. 

Specialized services also include organizing 

and publishing research results, distributing 

specialized research announcements and feeds, 

providing information desk and personal search 

support, and enabling data exchange services. 

For more information about the NASA STI program, 

see the following: 

• Access the NASA STI program home page at 

http://www.sti.nasa.gov 

• E-mail your question to help@sti.nasa.gov 

• Phone the NASA STI Information Desk at 

757-864-9658 

• Write to: 

NASA STI Information Desk 

Mail Stop 148 

NASA Langley Research Center 

Hampton, VA 23681-2199 

 

This page is required and contains approved text that cannot be changed. 
3 

NASA/TM—2014–218366 

 

Conceptual Launch Vehicle and 

Spacecraft Design for Risk Assessment 

Samira A. Motiwala 

Stanford University, Stanford, CA, 94305 

Donovan L. Mathias 

NASA Ames Research Center, Moffett Field, CA 

Christopher J. Mattenberger 

Science and Technology Corporation, Moffett Field, CA 94035 

June 2014
Acknowledgments 

The student author would like to thank her mentors, Donovan Mathias and Chris 

Mattenberger, as well as Ken Gee, Lorien Wheeler, and Susie Go for their extensive help 

and patience. The student author would also like to acknowledge the California Space 

Grant Consortium and the NASA Education Associates Program (EAP) for their financial 

support. 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

Available from: 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

This report is also available in electronic form at 

http://www.ntrs.nasa.gov

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 
5 

Table of Contents 

Nomenclature ..................................................................................................................6

1. Introduction .............................................................................................................. 6

2. Preliminary Spacecraft Sizing ................................................................................ 6

2.1. Mass and Power Estimation ......................................................................................... 7

2.2. Geometry ........................................................................................................................ 9

3. Launch Vehicle Sizing and Trajectory Analysis ................................................. 10

3.1. Design Assumptions ................................................................................................... 10

3.2. Optimal Staging ........................................................................................................... 11

3.3. Inboard Sizing .............................................................................................................. 11

3.4. Initial Two-Stage-to-Orbit Trajectory Analysis ......................................................... 12

3.5. Optimal Trajectory Analysis ....................................................................................... 14

3.6. Final Launch Vehicle Design ...................................................................................... 16

4. Concept of Operations .......................................................................................... 17

5. Spacecraft Subsystems Layout ............................................................................ 17

5.1. Structures and Mechanisms ...................................................................................... 18

5.2. Thermal Control System ............................................................................................. 20

5.3. Environmental Control and Life Support System (ECLSS) ..................................... 22

5.4. Power ............................................................................................................................ 24

5.5. Avionics ....................................................................................................................... 26

5.6. Propulsion .................................................................................................................... 28

5.7. Crew Accommodations .............................................................................................. 30

5.8. Entry, Descent, and Landing (EDL) ........................................................................... 30

5.9. Spacecraft Subsystem Summary .............................................................................. 31

6. Safety Design and Risk Analysis ......................................................................... 31

6.1. Spacecraft Reliability .................................................................................................. 32

6.2. Launch Vehicle Reliability .......................................................................................... 34

7. Summary ................................................................................................................. 35

References ....................................................................................................................35
6 

Nomenclature

γ = flight path angle, correlation factor 

γe = flight path angle at entry interface 

Δv = velocity increment 

ε = structural fraction coefficient 

λ = failure rate 

μ = standard gravitational parameter 

ν = true anomaly 

π = payload fraction 

ω = argument of perigee 

Ω = right ascension of ascending node 

a = acceleration, semi-major axis 

Ae = exit area 

At = throat area 

c = effective velocity 

cm = center of mass 

cp = center of pressure 

d = diameter 

D = drag, mission duration 

e = eccentricity 

E = modulus of elasticity 

fcu = ultimate compressive stress 

Fcr = elastic buckling stress 

Ftu = ultimate tensile stress 

Fty = yield tensile stress 

g = gravitational acceleration 

g0 = gravitational acceleration at sea level 

h = altitude/height 

he = altitude at entry interface 

hfg = latent heat of vaporization 

hi = altitude of initial circular orbit or 

 parking orbit 

I = moment of inertia 

Isp = specific impulse 

J = cost function 

LH2 = liquid hydrogen 

LOX/LO2 = liquid oxygen 

m = mass 

m0 = total mass 

mbo = burnout mass 

mGLOM = gross lift-off mass 

minert = inert mass 

mprop = propellant mass 

mstage = stage mass 

Me = expendable mass 

Mmax = peak bending moment 

n = load factor 

pu = ultimate internal pressure 

py = yield internal pressure 

Pavg = average power 

Pc = chamber pressure, 

 compressive axial force 

Ppeak = peak power 

q = dynamic pressure 

Q = heat rejection rate 

r = radius 

req, RE = Earth equatorial radius 

R = reliability 

RP-1 = Rocket Propellant-1 (kerosene) 

t = time 

t1, t2, t3 = thickness 

T = thrust 

T/W = thrust-to-weight ratio 

u = thrust directional control 

v = velocity 

Vce = local circular velocity at 

 entry interface 

x = downrange distance 

1. Introduction 

One of the most challenging aspects of developing human space launch and exploration 

systems is minimizing and mitigating the many potential risk factors to ensure the safest possible 

design while also meeting the required cost, weight, and performance criteria. In order to 

accomplish this, effective risk analyses and trade studies are needed to identify key risk drivers, 

dependencies, and sensitivities as the design evolves. The Engineering Risk Assessment (ERA) 

team at NASA Ames Research Center (ARC) develops advanced risk analysis approaches, 

models, and tools to provide such meaningful risk and reliability data throughout vehicle 

development. The goal of the project presented in this memorandum is to design a generic launch 
7 

vehicle and spacecraft architecture that can be used to develop and demonstrate these new risk 

analysis techniques without relying on other proprietary or sensitive vehicle designs. 

To accomplish this, initial spacecraft and launch vehicle (LV) designs were established using 

historical sizing relationships for a mission delivering four crewmembers and equipment to the 

International Space Station (ISS). Mass-estimating relationships (MERs) were used to size the 

crew capsule and launch vehicle, and a combination of optimization techniques and iterative 

design processes were employed to determine a possible two-stage-to-orbit (TSTO) launch 

trajectory into a 350-kilometer orbit. Primary subsystems were also designed for the crewed 

capsule architecture, based on a 24-hour on-orbit mission with a 7-day contingency. Safety 

analysis was also performed to identify major risks to crew survivability and assess the system’s 

overall reliability. These procedures and analyses validate that the architecture’s basic design and 

performance are reasonable to be used for risk trade studies. While the vehicle designs presented 

are not intended to represent a viable architecture, they will provide a valuable initial platform 

for developing and demonstrating innovative risk assessment capabilities. 

2. Preliminary Spacecraft Sizing 

An initial spacecraft design begins with a rough assessment of the requirements and scope of 

its intended mission. NASA is currently administering its Commercial Crew Program (CCP), a 

multiphase program intended to help the US industry develop space transportation systems to 

safely launch astronauts and cargo to the ISS and other low-Earth orbit destinations1

. The generic 

space vehicle system designed here will apply the same requirements defined for the CCP, which 

include delivering four crew members and their equipment to the ISS while ensuring safety in 

the event of an emergency on the pad or during launch and ascent. In addition, the spacecraft is 

expected to demonstrate its ability to serve as a 24-hour safe haven during an on-orbit 

emergency1

, but the design presented here will include a 7-day contingency as well. 

2.1. Mass and Power Estimation 

Conceptual design of this spacecraft begins with estimating the pressurized volume required 

by the crew, which can be determined from historical guidelines based on mission duration, as 

shown in Figure 1. 

Figure 1. Pressurized volume as a function of mission duration with historical reference cases2. 
8 

Applying this guideline to a 7-day mission yields roughly 4 m3 of pressurized volume per 

person, or 16 m3 of total pressurized volume. After the crew compartment volume is determined, 

a first estimate of the vehicle’s burnout mass (mbo) can be calculated based on historical data for 

human spacecraft using the following equation3

: 

            (1) 

The total dry mass of the spacecraft module consists of the burnout mass and the entry, 

descent, and landing (EDL) mass, which can be estimated as 15-18% of the burnout mass4. 

Propellant mass (mprop) may be determined from the ideal rocket equation: 

 



   (2) 

The velocity increment, Δv, required for this mission consists of the velocity increment 

needed to deorbit as well as a 50% margin for emergency maneuvers, orbital corrections, and 

orbital decay. A standard storable liquid propulsion system using monomethylhydrazine and 

nitrogen tetroxide (MMH/NTO) is assumed, with a specific impulse of 310 seconds. As a general 

guideline, the mass of the propulsion system without the propellant (tanks, lines, engines, etc.) 

can be estimated as 15% of the propellant mass4

 and is included back in the original burnout 

mass. This process repeats for a few iterations until a near convergent solution is reached. 

Deorbit Δv can be determined from the following equation5

: 

  



   



 





 (3) 

where 

     



 (4) 

Parameters required to estimate deorbit Δv (flight path angle and altitudes) were taken from 

Space Shuttle values5 as a best estimate, and the total Δv required (with a 50% margin) was 

estimated to be about 197.5 m/s. 

After estimating the spacecraft’s total dry mass, each of its subsystems need to be defined 

and allocated a percentage of the dry mass (using MERs) to obtain a first-order mass estimate. 

These subsystems include: structure and mechanisms, thermal control, environmental control and 

life support system (ECLSS), power, avionics/communications, attitude determination and 

control, crew accommodations, propulsion, EDL, and the payload (astronaut crew and suits). 

Payload is assumed to be 90 kg (~200 lb) for each person and 50 kg (~110 lb) for each suit 

system6

. 

After the inert mass has been determined, a design margin of 25% is added to all subsystems 

to enable room for mass growth, and wet mass (propellant and other consumables) is added to 

the mass budget. Finally, the LAS mass is included, and is estimated to be roughly 70% of the 
9 

inert mass* (with margin). The total mass of the spacecraft with LAS was determined to be about 

13,800 kg. Table 1 provides a mass breakdown of the spacecraft’s main subsystems and other 

components. 

Table 1. Preliminary mass budget for spacecraft design. 

Subsystem % Dry Mass Mass (kg) With 25% Margins (kg) 

Structure 22 1,251 1,564 

Mechanisms 8 455 569 

Thermal Control 9 512 640 

ECLSS 8 455 569 

Power 15 853 1,066 

Avionics 10 569 711 

ADCS/GNC 2 114 142 

Crew Accommodations 8 455 569 

EDL 18 1,023 1,279 

TOTAL DRY (W/O PROPULSION) 100 5,686 7,107 

Propulsion - 50 - 

Payload (Crew with Suits) - 560 - 

TOTAL DRY (INERT) - 7,667 

Main Propellant - 332 - 

Propellant for ADCS - 213 

Consumables - 224 - 

TOTAL SPACECRAFT - 8,436 

Launch Abort System - 5,367 - 

GROSS MASS 13,803 

Power budgets for space vehicles vary depending on mission requirements, but general 

guidelines4

 for estimating the average and peak power required are: 

          (5) 

   (6) 

These estimates yield average and peak power requirements of 3 kW and 5.25 kW, 

respectively. 

2.2. Geometry 

The spacecraft command module (CM) is assumed to be a conical frustum with the 

corresponding parameters shown in Figure 2. The service module (SM) is estimated as a 

cylindrical shell with the same large diameter and a height of 3 meters. Inertia properties can be 

found in Table 11(a) in section 5.5. 

 

* Based on 2 previous existing abort systems and guideline provided by INTROS7
10 

Figure 2. Command module geometry. 

3. Launch Vehicle Sizing and Trajectory Analysis 

Sizing a launch vehicle can involve many different approaches and employ a variety of massestimating techniques. The architecture developed here uses a top-down approach involving 

optimal staging to size the inert and propellant stage masses (based on assumptions), designing 

the inert stage components (tanks, engines, etc.), and then comparing their masses to estimated 

ones, analyzing a rough trajectory performance to determine feasibility, and iterating multiple 

times until a closed, satisfactory solution is reached. 

3.1. Design Assumptions 

A TSTO launch vehicle is assumed since it is the most common and practical design. A 

single-stage vehicle is not practical for achieving higher burnout velocities, and three (or more) 

stages provide unnecessary extra performance for added weight8

 and risk penalties from 

increased complexity. The launch vehicle must generate enough v to meet mission 

requirements and place the spacecraft into orbit while accounting for gravity and drag losses and 

possible launch site gains. The spacecraft’s speed at orbital insertion must match the speed 

required to maintain the selected orbit. It is assumed that the vehicle will launch from Cape 

Canaveral (28.5° latitude) into a 350-km circular orbit at an inclination of 51.6 degrees to dock 

with the ISS, and will therefore require a total design v of about 9.5 km/s8

. 

Other assumptions include propellant types (fuels and oxidizers), specific impulses, structural 

mass fractions, thrust-to-weight ratios, and number of engines for each stage. Table 2

summarizes the launch vehicle assumption parameters. 

Table 2. TSTO launch vehicle assumptions.

Stage Propellant Specific Impulse (s) Structural Mass Fraction* T/W* # Engines 

1 Lox/RP-1 350 0.06 1.3 4 

2 Lox/LH2 450 0.1 0.75 2 

*Initial assumption (changes with iterations)
11 

3.2. Optimal Staging 

Sizing each stage while minimizing the total LV mass requires an optimal staging technique 

that utilizes the Lagrange multiplier method, a mathematical optimization strategy for solving the 

local maxima and minima of a given bivariate function subject to certain constraints†

. This 

technique involves defining the specific impulse and structural mass fraction of each stage, 

payload mass, and design Δv to yield the minimum-mass N-stage vehicle to carry that payload 

into orbit. This method is in contrast to the restricted staging assumption that all stages of a 

tandem-stacked vehicle are similar and should be designed with the same specific impulse, 

payload ratio, structural mass fraction, and consequently, mass ratio9

. 

3.3. Inboard Sizing 

After an initial sizing of all stages is complete, the LV inboard profile is considered. This 

includes deciding how all LV components (engines, tanks, attachments, etc.) will fit into a given 

envelope. Sizing procedures involve choosing the LV geometry, dimensionalizing the 

components, and using MERs for all components to determine the vehicle’s total expected inert 

mass8

. 

Three iterations were initially carried out to achieve a positive design margin between 

expected stage inert masses (from optimal staging) and calculated stage inert masses (from 

inboard sizing). The first iteration assumed spherical tanks (which require a large diameter) and 

yielded a large negative design margin. The second iteration scaled the LV diameter down to 4 m 

while resizing the propellant tanks (to cylindrical shells with spherical domes) to fit. This 

improved the margin significantly because it reduced the required fairing surface area, thus 

reducing mass. However, the margin was still negative, and the diameter can only be decreased 

so much before the LV is considered infeasible due to structural complications. The third 

iteration experimented with different structural mass fractions‡

 for each stage until a positive 

design margin was achieved. Generally, higher structural mass fractions yield heavier vehicles, 

so these fractions were chosen to be as low as possible while still yielding a design margin of 

about 10%. 

Engines were sized by assuming a telescopic, crank-down nozzle for each engine, with an 

80% nozzle length and 15° cone divergent, 30° convergent half-angles. Guidelines exist8

 for 

estimating the characteristic chamber length (0.75 m) and the ratio of the chamber length to 

diameter (typically 0.9 m). In addition, typical parameters associated with different fuel and 

oxidizer combinations can be found in Table 38

. 

Table 3. Typical engine parameters. 

Parameter Stage 1 LOX/RP-1 Stage 2 LOX/LH2

Chamber Pressure, Pc (psi) 3,000 2,500 

Expansion Ratio (Ae/At) 30 150 

Theoretical c (m/s) 1,825 2,325 

Delivered c (m/s) 1,789 2,279 

 

† See Curtis9

 section 11.6.1 (“Lagrange Multiplier”) for more information

‡ Typical structural mass fractions range from 0.038 to 0.45, based on historical data10
12 

3.4. Initial Two-Stage-to-Orbit Trajectory Analysis 

The LV trajectory analysis was initially conducted with a simple gravity turn trajectory, 

where the launch vehicle takes off vertically and gradually transitions into horizontal flight at 

orbital insertion8-9. The ascent was separated into three phases: first stage propellant burn and 

separation, second stage ignition and LAS ejection 30 seconds later, and second stage propellant 

burn and separation prior to orbital insertion. Performance assumptions include constant thrust 

during burn with no steering losses. A five-second coast period was also assumed between 

propellant burnout and stage separation (for both stages), with no thrust during the coast. 

Aerodynamic assumptions included a drag coefficient of 0.2 and a reference area calculated from 

the 4-m diameter cross-section. 

Figure 3 illustrates a simplified free-body diagram of the launch vehicle in flight. A vertical 

liftoff is followed by a small “kick” in the flight path angle after the launch vehicle clears the 

tower, which is usually around 100–150 m9

. 

Figure 3. Simplified launch vehicle free-body diagram for gravity-turn trajectory. 

Equations (7) to (10) are the primary coupled equations of motion8-9 to propagate and iterate 

until a satisfactory convergent solution is reached. These equations account for the 

aforementioned losses and v boost obtained from launching due east in order to achieve the 7.7 

km/s circular orbit velocity at insertion. 



      (7) 

 

    



  (8) 



 



   (9) 



   (10) 
13 

Figure 4 and Figure 5 display the primary results obtained from this basic trajectory analysis. 

The performance parameters presented include the flight path angle (degrees), acceleration 

(m/s2

), velocity (x 10 km/s), altitude (km), and dynamic pressure (x 0.01 Pa). The 2D spatial 

trajectory of altitude vs. downrange distance is also presented. The initial kick from the flight 

path angle was determined to be 0.3 degrees (i  89.7 degrees) and the total time of flight was 

estimated to be about 550 seconds. The gravity-turn trajectory served as a precursor to optimal 

trajectory analysis, where a good first guess is required to achieve a desired solution. 

Figure 4. Two-stage-to-orbit performance results. 

Figure 5. 2D spatial trajectory results. 


14 

3.5. Optimal Trajectory Analysis 

In practice, trajectories are simulated by utilizing multiple maneuver techniques to achieve 

the best trajectory for a given cost function (such as minimizing fuel or maximizing final 

velocity). Launch trajectory optimization involves obtaining the LV dynamic states and controls 

that optimize the chosen performance index while satisfying imposed trajectory constraints8. End 

conditions (orbital parameters) are specified and the optimization problem is deconstructed into 

predetermined launch phases that are linked together to avoid discontinuities in the trajectory (no 

coast periods are involved). 

Multiple trajectory optimization software programs exist for analyzing these trajectories. 

GPOPS (General Pseudospectral OPtimal control Software)—a 3-DOF, MATLAB-based, opensource program that uses hp-adaptive pseudospectral methods to solve multiple-phase optimal 

control problems—was utilized to solve this ascension problem11. The “Multiple-Stage Launch 

Vehicle Ascent Example” from GPOPS was modified from a two-stage vehicle with nine strapon solid rocket boosters to the currently investigated vehicle. 

The GPOP dynamic model uses the equations of motion for a non-lifting point mass 

ascending over a spherical rotating planet in Cartesian Earth-centered inertial (ECI) coordinates: 

   (11) 

   

    

    (12) 

   



 (13) 

Boundary constraints include initial position and velocity conditions based on launch site and 

gross liftoff mass. End constraints were defined by orbital elements (see Table 4) that describe 

the payload’s final expected location in orbit. The program implemented two path constraints: 

one to keep the vehicle’s altitude above the surface of the Earth and another to guarantee the 

control vector (thrust direction) magnitude remains at unit length. Lastly, linkage conditions set 

constraints to force state variables to remain continuous during phase changes and account for 

mass ejections. The optimal control problem is then to determine the control that minimizes the 

cost function: 

  

 (14) 

Where the superscript (3) represents the third (final) phase, and tf represents the final time at 

the end of flight. 

Table 4. End-of-flight orbital parameter constraints. 

Orbital Parameter Value 

af (km) 6,748 

ef 0.004 

if (deg) 51.6 

Ωf (deg) 206.4 

ωf (deg) 103 

νf (deg) unconstrained 
15 

The stage masses and thrust-to-weight ratios obtained from the gravity-turn trajectory 

analysis were used as initial inputs for the optimal control trajectory problem. A number of test 

cases were executed with variable parameters changed until a final satisfactory solution was 

obtained. Figure 6 displays time histories of the altitude, velocity, mass, and control performance 

parameters, and Figure 7 shows the 2D spatial trajectory of altitude vs. downrange distance. 

Figure 6. Time histories of (a) launch altitude, (b) launch speed, 

(c) launch mass, and (d) directional thrust control. 

Figure 7. 2D spatial trajectory results. 
16 

These results are not completely satisfactory because dynamic loading of the vehicle during 

max-q (maximum dynamic pressure) and maximum acceleration need to be considered for 

structural and human factor purposes. It is recommended that a launch vehicle experience only a 

few hundred pounds per square foot (psf) of dynamic loading, whereas the calculated results are 

around 1,100 psf during max-q. Human-rated missions strive to keep the peak acceleration at or 

below 3 g for safety purposes, but results indicate a peak acceleration of about 3.57 g. These 

issues are typically resolved by allowing the thrust to vary and throttling back before reaching 

max-q. Unfortunately, the GPOPS code does not enable thrust variation. Future work will 

include the use of more complex optimization programs, such as OTIS12 (Optimal Trajectories 

by Implicit Simulation), which enable more constraint flexibility and parameter variations 

(including thrust). 

3.6. Final Launch Vehicle Design 

Figure 8 summarizes the final launch vehicle design§

. 

Figure 8. Final launch vehicle design summary. 

 

§ Sketch was created with INTROS7

Item Stage 

1 2 

Prop LO2/RP-1 LO2/LH2

T 3280800 N 393369 N 

Pc 20680 kPa 17240 kPa 

MR 2.47:1 6.0:1 

Isp 350 s 450 s 

 0.11 0.16 

 0.177 0.147 

mp 172954 kg 33331 kg 

m0 194330 kg 39680 kg 

mGLOM 247814 kg 53483 kg 

T/W 1.35 0.75 

V 4290 m/s 5209 m/s 

# engines 4 2 

	



• 

"


#
/+*-

• 
,*-.

• 

*'


)



$	





)

+ ''

+ ''

+ ,0

* '(

, +)

/ ,.

, ,(

- ,

() /

,* 0

)

)

!(
17 

4. Concept of Operations 

Figure 9 depicts the mission concept of operations. During ascent, the first stage separates 

after about 180 seconds of burn time, followed by LAS jettison 30 seconds after second-stage 

ignition, and finally second-stage separation and orbit insertion after about 380 seconds of 

second-stage burn time. Under nominal conditions, the spacecraft orbits Earth for 24 hours and 

makes any orbital corrections needed to rendezvous and dock with the space station. The 

spacecraft remains docked for approximately six months, with routine one-hour checkouts every 

week. Once undocked, the spacecraft prepares for re-entry by performing orbital maneuvers and 

a “cold soak” to reject any heat accumulated during the docked period. The spacecraft then 

initiates a deorbit burn with the main engine to reach the entry interface altitude. The service 

module is jettisoned after the burn and the spacecraft prepares to ballistically re-enter the 

atmosphere. As the spacecraft begins its approach for landing, the parafoil deploys and can be 

steered towards a safe ocean splash zone or ground landing area. 

Figure 9. Concept of operations. 

5. Spacecraft Subsystems Layout 

After approximating the mass for each spacecraft subsystem, a more detailed analysis is 

performed to generate a subsystem mass equipment list, determine if the total subsystem mass is 

within the pre-allocated mass envelope, and ensure closure of these subsystems. Each subsystem 

is designed to the component level for identifying major risks to the mission and crew and aiding 

in reliability estimates. Most of the subsystem designs are based on procedures outlined in 

Human Spaceflight4

 unless otherwise noted. 





Liftoff 

Cape Canaveral 

t = 0 sec 

1st Stage Separation 

Alt = 80 km 

t = 181 sec 

LAS Jettison 

Alt = 117 km 

t = 211 sec 

2nd Stage Separation 

and Orbit Insertion 

Alt = 350 km 

t = 560 sec 

Earth Orbit 

TOF: 24 hrs 

ISS Dock 

Duration: 6 months 

Undock, Prep & Cold Soak 

TOF: 3 hrs 

De-Orbit Burn 

60 min to touchdown 

Alt: 300 km 

 Entry Interface & SM Jettison 

30 min to touchdown 

Alt: 200 km 

 Parafoil Deployment 

2 min to touchdown 

Alt: 5 km 

 Land 

Ocean splashdown or 

ground landing 
18 

5.1. Structures and Mechanisms 

The structures subsystem is essential to the integrity of the pressurized crew compartment 

and protects the crew from various loads during the mission. It is often divided into three 

categories: primary (backbone), secondary (appendages), and tertiary (small) structures. The first 

two categories are the most challenging and influential in the design. Since structures have been 

historically reliable in human-rated space missions, a detailed mission loads analysis is not 

necessary for selecting the material and estimating the capsule mass. An exact architectural 

layout is beyond the scope of this project. 

Since welding is the most reliable method to seal a pressurized structure, composites and 

sandwich construction will not be used for the spacecraft shell. Steel will also not be used since it 

is susceptible to instability under compressive loads. Aluminum alloys are easily welded and 

capable of surviving extreme temperatures in space. Aluminum 6061-T6 in particular is less 

expensive and more readily available than some of its other aluminum alloy counterparts. 

Allowable tensile-ultimate and compressive-yield stresses for this material are, respectively, 160 

MPa and 100 MPa with high-grade machine welds (290 MPa and 240 MPa without welds), and 

safety factors for ultimate and yield conditions are 2 and 1.5, respectively. 

The maximum internal pressure for the crew module design in space is 0.1096 MPa (slightly 

more than one atmosphere), so the design ultimate internal pressure is twice as much, or 0.219 

MPa. Away from the welds, the tensile stress resulting from this pressure cannot exceed the 

allowable ultimate tensile stress of 290 MPa. The required thickness away from the welds for a 

capsule with a 2.5 m (2,500 mm) radius is therefore: 

  



 

   (15) 

This thickness must be increased about 2 cm from each weld to keep the design yield stress 

below the welded allowable stress of 100 MPa: 

  



 

    (16)

The ultimate compressive stress needs to be kept less than the elastic buckling stress. For an 

unpressurized, isotropic, monocoque cylinder (assuming plasticity correction factor of 1.0), the 

elastic buckling stress is given by: 

 



 (17)

The ultimate compressive stress, with 1.4 and 1.5 ultimate factors of safety for inertia loads 

and pressure, respectively, is: 

  



 

   

 (18) 

where the peak bending moment, Mmax, can be estimated from the mass and geometric 

properties of the capsule as well as the load factors assumed for each phase of flight (taken from 

those experienced by the Shuttle). The compressive force, Pc, is simply estimated as the capsule 

weight multiplied by the negative axial load factor. 
19 

Assuming the correlation factor, γ, used to match theory with results is 0.5, using the 

modulus of elasticity for aluminum 6061-T6 (69,000 MPa), and rearranging the inequality to 

determine the required thickness for stability gives: 

  



 

     

 (19) 

The capsule thickness results for launch and normal landing flight modes are summarized in 

Table 5. Taking the larger thickness of 2.13 mm and substituting it in Equation (17) to determine 

the ultimate compressive stress gives: 

  

   

    (20) 

This is well under the material’s proportional limit. Since t3 > t1, the capsule shell is critical 

for stability and must be at least 2.13 mm thick away from the welds and 4.11 mm thick near the 

welds. 

Table 5. Estimated capsule thickness to prevent buckling under launch and landing loads. 

Mission Phase

Limit Load Factors 

(from Shuttle) Min. 

Pressure 

(MPa)

Limit 

Compressive 

Axial Force, Pc

(N)

Limit 

Moment, Mmax

(106 mmN)

Required 

Thickness, t3

nx ny nz (mm)

Launch -3.2 1.4 2.5 -0.00345 264,800 104 2.13 

Normal Landing -2.0 1.5 -4.2 -0.00345 165,500 161 1.99 

The mass of the primary structure itself is the product of the material density and the volume 

of the structural module shell. Since the thickness of the capsule varies, the product of the 

aluminum’s area density with the module surface area can be used to estimate the mass. The area 

density of metallic structures typically ranges from 5 to 15 kg/m2

, so a conservative density of 10 

kg/m2

 is chosen, which gives a total of 1,021 kg for the primary shell structure (command and 

service modules). 

A launch vehicle adapter (LVA) is needed to connect the LV structure to the spacecraft 

structure and enable spacecraft separation. LVA mass is a strong function of spacecraft mass, as 

shown in Figure 10. Using the trend-line equation from this figure, the total estimated mass for 

the LVA is about 1,092 kg. This is a conservative estimate for the adapter; the LVA for the 

Orion mission is about 580 kg, which is just a little more than half this mass estimate for a 

heavier spacecraft. Including hatches, a docking adapter, and smaller items such as tolerances, 

joints, thermal coatings, etc. (estimated at about 40% of the primary structure), the total mass 

estimate for the subsystem is about 2,700 kg. Table 6 gives a mass breakdown for the structures 

and mechanisms subsystem. 
20 

Figure 10. Launch vehicle adapter mass13. 

Table 6. Structures and mechanisms breakdown. 

Item Mass (kg)

Primary Structure 1,021 

Hatches (Side and Forward/Docking) 130 

CM-SM Umbilical System 5 

Docking Adapter 50 

Launch Vehicle Adapter 1,092 

Tolerances, Joints, Coatings, Etc. 400 

TOTAL 2,698 

5.2. Thermal Control System 

The thermal control system (TCS) maintains normal operating temperatures for subsystems 

and equipment, as well as comfortable temperatures for the crew. Two types of systems are used: 

active or passive systems. Crewed spacecraft require higher power levels and temperatures to 

accommodate human comfort levels, so active controls need to be implemented. The TCS is 

different from the thermal protection system, which shields the spacecraft from extreme heat 

sources and sinks. 

Sizing the thermal system involves adding all the heat loads aboard the spacecraft. However, 

the largest contributions to these loads are from electrical equipment. Therefore, at this stage it is 

adequate to assume that the total heat load for thermal control is equal to the capacity of the 

electrical power system, which was previously estimated to be 3 kW. Adding a margin of 20%, 

the required total heat load for thermal control is assumed to be 3.6 kW. 

Thermal control systems consist of three primary elements: heat acquisition, transport, and 

rejection. The heat acquisition subsystem collects heat from a component via heat-transferring 

devices (such as a cold plate). The heat transport subsystem carries heat from the acquisition 

subsystem to be dissipated. The heat rejection subsystem expels excess heat into deep space by 

radiation or dissipates it with short-duration heat sinks. The decision to use heat sinks or 

radiators depends on the mission duration. Heat sinks are typically used for extremely short 

durations (3 days or less for near-Earth orbits), while radiators are typically used for longer 

durations. For this design, heat sinks will be used during pre-launch and post-landing phases and 

radiators will be used during on-orbit operations. In addition, refrigerants such as ammonia 

boilers will be used to handle re-entry loads. 
21 

Another element of thermal control is heat generation, which consists of heaters that are used 

to heat subsystems during the docking period. Heat acquisition subsystems consist of heat 

exchangers and cold plates, and heat transport subsystems include pumps (with accumulator), 

plumbing/valves, instruments and controls, and liquid coolant. A fixed-surface radiator was 

selected for on-orbit heat rejection and its mass can be estimated as 5.3 kg/m2 for a two-sided 

panel. The approximate mass of an expendable heat sink (Me) is a function of the required heatrejection rate, mission duration, and latent heat of vaporization (a fluid property): 

  (21) 

During pre-launch operations for the Shuttle, up to 250 watts of power were available to 

perform payload checkouts14. This power requirement will be used for pre-launch and postlanding heat rejection rates. An ammonia boiler similar to that used by the Shuttle will be used 

during re-entry. The mass of the Shuttle’s ammonia boiler system is about 72 kg (including 44 

kg of ammonia) to reject up to 33 kW of thermal energy briefly before, during, and after landing. 

This mass estimate will be conservative for the spacecraft described here considering the 

relatively low heat rejection loads experienced during this flight phase. 

Passive thermal control will consist of multi-layer insulation (MLI), which contains many 

thin reflective layers separated by vacuum to help protect the capsule from solar heating and 

reduce heat loss through thermal radiation. Advanced Flexible Reusable Surface Insulation 

(AFRSI) blankets15 used on the Shuttle are considered here to estimate the mass of the MLI, 

although the technology itself may be outdated and a new, more effective insulation blanket 

could possibly be used. 

Table 7 summarizes the mass, power, and volume requirements for the thermal control 

hardware, which are stored in the command module. Hardware is sized using MER values based 

on capacity (per kW of heat transfer) or surface area size. Redundancies are included for heat 

acquisition and heat transport subsystems to achieve a reliable thermal control system. The total 

mass is about 600 kg. 

Table 7. Thermal control system requirements. 

Item Mass (kg) Power (W) Volume 

(m3) 

Hardware for Active Thermal Control 

Heat Generation Heaters (3 redundant) 4 50 - 

Heat Acquisition Heat Exchanger Pair (2 redundant sets) 3.9 0 0.0001 

Coldplates (3 x 1 kW, with redundant set) 72 0 0.168 

Heat Transport 

Pumps with Accumulator (with redundant set) 34.56 82.8 0.1224 

Redundant Plumbing & Valves 31.97 negligible - 

Instruments & Control 17.98 negligible negligible 

Fluids 17.98 - - 

Heat Rejection 

Radiator (27.5 m2

 fixed, 2-sided) 145.6 negligible 0.55 

Heat Sinks 2.16 negligible - 

Ammonia Boiler 72 negligible - 

Hardware for Passive Thermal Control 

Multi-Layer Insulation 200 0 0.7427 

TOTAL 602 132.8 1.58 
22 

A schematic of the TCS is shown in Figure 11. Two loops are included for redundancy: the 

primary (inner) loop and the secondary (outer) loop. 

Figure 11. TCS schematic. 

NASA has recently been developing an advanced TCS technology concept called the 

Sublimator Driven Coldplate16 (SDC), which combines all three TCS functions into a single 

piece of hardware. This novel concept has the potential to replace an entire thermal control 

system with just one hardware component that does not consume any power or contain moving 

parts. This type of TCS is ideal for low heat loads and short mission durations, which could 

make it useful for this application. Future work may update the generic spacecraft design to 

consider this technology. 

5.3. Environmental Control and Life Support System (ECLSS) 

Although all spacecraft subsystems are vital for a successful mission, only the ECLSS exists 

to keep the crew alive and comfortable. Primary ECLSS functions typically include atmosphere 

management (monitoring, revitalization, control, and supply); water and waste management; fire 

detection/suppression, and food storage/preparation17. For this vehicle, consumables and other 

crew accommodations (e.g. supplies, equipment, and other provisions) will be considered as a 

separate “crew accommodations” subsystem (see section 5.7) and not as part of the ECLSS. 

Table 8 provides a summary of mass, power, and volume requirements for the ECLSS. The total 

subsystem mass is around 820 kg. 

Cold 

Plate 2 

Fuel Cell 

Heat 

Exchangers 

Radiator Panel 

GSE Heat 

Exchanger 

Heat 

Sinks 

Cold 

Plate 3 

Cabin Heat 

Exchanger 

Suit Loop 

Heat 

Exchanger 

Ammonia 

Boiler 

NH3 

Accumulator 

Accumulator 

Flow Control 

Assembly 

Pumps 

Relief 

Valve 

Relief 

Valve 

Cold 

Plate 1 
23 

Table 8. ECLSS requirements. 

Item Mass 

(kg) 

Power 

(W) 

Volume 

(m3) 

Pressure Control Subsystem (PCS) 

Plumbing (CM) 26.76 - - 

Plumbing (SM) 5.96 - - 

Air Revitalization Subsystem (ARS) 

Stored Tanks (CM) 60.5 0 2.73 

Stored Tanks (SM) 151.2 0 6.83 

LiOH Cartridges (7) 49 12 0.035 

HEPA Filters (2) 17 - 0.053 

CCA 30 - 0.1 

Atmospheric Monitoring Subsystem (AMS) 

Sensors (CM) 3.52 - - 

Sensors (SM) 1.76 - - 

Mass Spectrometer 2.4 100 0 

Temperature and Humidity Control (THC) 

Condensing Heat Exchangers (2) 2 100 0.1 

Heaters (2) 4 50 - 

Water and Waste Management (WWM) 

Stored Water (CM) 57.4 0 0.056 

Stored Water (SM) 71.8 0 0.07 

Multifiltration 40 160 0.16 

VCD 100.0 120 0.4 

Waste Collection 45 45 0.05 

Supplies and Contingency Bags 7.9 0 0 

Fire Detection and Suppression (FD&S), Suit Loop 

Fire Suppressant System 70 - - 

Smoke/CO Detectors (2) 0.8 - - 

Suit Loop 70 - - 

TOTAL 817.0 587 10.58 

The design includes two sets of air generation systems (as shown in Figure 12) and two water 

storage tanks—one of each for the CM and another for the SM. This saves space in the CM and 

reduces risk of mission or crew failure if any of the pressurized tanks rupture. All other ECLSS 

components are stored in the CM. Redundancies are included wherever appropriate and some 

components, although not critical, are added for completeness. 

The atmosphere management subsystem consists of pressure control, air revitalization, 

atmospheric monitoring, and temperature and humidity control. Air revitalization consists of 

ventilation, make-up gases supply (nitrogen and oxygen), and contaminant removal. 

Contaminant removal includes controlling and removing carbon dioxide in the atmosphere to 

keep it from becoming toxic. Early spacecraft designs have used lithium hydroxide (LiOH) 

canisters to absorb CO2, but other options include molecular sieves, solid amine water 

desorption, and electrochemical depolarization concentration. LiOH canisters remove about 2 kg 

of LiOH daily per each person’s CO2 output, which cannot be regenerated, making them only 

ideal for short-duration space missions. Molecular sieves, on the other hand, use synthetic 

zeolites or alumino-silicated metal, which can be regenerated. The design selects the LiOH 

canisters because they provide an advantageous mass savings over the sieves for a relatively 
24 

short mission. The design also includes a contamination control assembly (CCA) to remove 

carbon dioxide and acidic gases inside the cabin, along with HEPA filters to effectively remove 

small particles and bacteria. 

Figure 12. ECLSS air generation system schematic. 

Water management consists of storing water, processing water and urine, and distributing 

water through plumbing, which can be neglected for preliminary design. Water tanks are stored 

on board and the supply comes from pre-stored water, fuel cell byproducts (see Power section 

5.4 for details), and humidity condensate. Multifiltration is used for processing water and vapor 

compression distillation (VCD) is used to collect and process urine. Waste management consists 

of a waste collection system and associated supplies. 

The fire detection and suppression (FD&S) subsystem typically consists of smoke detectors, 

alarms and shutoff systems, portable fire extinguishers, gas masks, and oxygen bottles18. 

5.4. Power 

Power system options vary depending on mission duration and amount of power required. 

Batteries are very effective power sources for unmanned spacecraft but are not feasible for 

manned spacecraft missions that last at least a few days and require several kilowatts of power. 

Solar cells are ideal for lengthy missions (weeks to years) and orbiting stations, but they require 

complicated deployment mechanisms and are not ideal for shorter missions. Nuclear reactors can 

be massive and hazardous to humans, and radioisotope generators tend to be massive and 

expensive. Regenerative fuel cells are an ideal option for this type of mission and are common 

powering methods for Earth-orbiting spacecraft missions because they produce high energy per 



'

&

!!

	!#"

!"

	"

!$

"

"

!	


!	

	

#"

! 
#	#


 





25 

unit of reaction mass. Hydrogen fuel cells also generate water as a byproduct of electrolysis, and 

that can be circulated back into the ECLSS for use. The primary disadvantage of fuel cells is the 

requirement for storing hydrogen and oxygen, but the overall system mass would still be less 

compared to a battery-powered system. 

Power estimates from each subsystem were used to approximate the total baseline and peak 

power required for the mission. Peak power was determined by taking the highest power 

requirement value from the crew accommodations subsystem since its components operate for 

only a small fraction of the mission duration. These estimated power requirements match closely 

to initial estimates. Therefore, the same power requirements (with margin) are assumed. 

Since no MER or power-rating relationship exists for estimating the mass and power supply 

of a fuel cell, these specifications were taken from a specific fuel cell manufacturer. The 

parameters shown in Table 9 were calculated from the manufacturer’s 10.5 kW fuel cell 

datasheet19 for a 7-day contingency mission. Hydrogen and oxygen are the reactants and water is 

the emission to be circulated to the ECLSS. The fuel cell efficiency for this specific model is 

around 51-69% and the durability/lifetime target is about 10,000 hours20. 

Table 9. Fuel cell parameters for a 7-day mission duration. 

Parameter Hydrogen Oxygen Water 

Mass (kg) 2.966 47.05 16.92 

Mass Flow Rate (kg/hr) 0.0141 0.112 0.101 

Tank Volume (m3) 0.367 0.367 - 

Tank Diameter (m) 0.888 0.888 - 

Tank Thickness (mm) 6.44 6.44 - 

Tank Mass (kg) 289 289 - 

Total Mass (kg) 291.97 336.05 16.92 

Power management and distribution components are also considered in the power subsystem 

design. These include the regulators, converters, charge controllers, and wiring required to 

connect the power sources and deliver the appropriate voltage-current levels to all power 

subsystems and loads while enabling the system’s degradation during the mission. Figure 13 

displays a generic power system block diagram with fault protection for the power management 

system. 

The electrical power system can be decomposed into three categories: generation, 

distribution, and storage. Power generation consists of the fuel cell stacks, heaters and heat 

exchangers, and plumbing (pressure regulator and sensor, purge valve, and water separator) used 

to generate the power and control flow from the reactant gases. Distribution involves the cabling, 

coolant, bus, and power control units. The storage involves the reactant gases and water output. 

Two additional fuel cell, heater, and heat exchanger hot spares are included for redundancy. The 

mass of the power distribution system is estimated as 6% of the initial spacecraft mass. Table 10 

gives a breakdown of the fuel cell system masses. The total mass of the power system is about 

1,030 kg. 
26 

Figure 13. Generic power system block diagram. 

Table 10. Fuel cell system mass breakdown. 

Item Mass (kg) Volume (m3) 

Power Generation 

Fuel Cell Stacks (3) 32.1 0.048 

Heaters (6) 6 - 

Heat Exchangers (3) 1.95 - 

Plumbing 6.88 - 

Power Distribution 

Cabling, Coolant, etc. 340 - 

Power Storage 

Hydrogen Tank 292 0.367 

Oxygen Tank 336 0.367 

Water Output 16.9 0.017 

TOTAL 1,032 0.80 

5.5. Avionics 

Avionics are the electronics used on a spacecraft and typically include guidance, navigation, 

and control (GNC); tracking/communications; command and data handling; and crew interface. 

Most avionics items were selected from examples of currently available components since 

electronics continue to rapidly improve, resulting in weight, power, and volume savings. A few 

components were selected from the Shuttle orbiter and assigned a “Shuttle technology factor” to 

compensate for technological advances. 

GNC is the integrated field that studies the combination of sensors, actuators, and algorithms 

to control the vehicle’s attitude and apply the necessary torques to re-orient it to a desired state. It 

is mainly responsible for stabilizing the space vehicle when disturbance torques are encountered. 

For low-Earth orbiting spacecraft, the primary disturbances encountered are atmospheric drag 

and gravity-gradient torques. For this specific mission, where no interplanetary transfers or plane 

changes occur, GNC is only used for orbital correction maneuvers and re-entry control. 

 Fuel Cells 

(2 hot spare) 

 

Hydrogen/Oxygen 

Tanks 

 Operational 

Heaters 

 Heat 

Exchangers 

 

DC Power 

Distribution 

Unit @28 

VDC 

 ECLSS 

 Electrical 

Control Units 

 

Shunt 

Regulator 

Unit 

 

DC /DC 

Converter & 

Regulator 

 

Fault 

Detection 

Circuit 

 Subsystems/Loads 

 

Internal/

External 

Heaters 

 



 

 

 

 

 

 

 

 

Pressure Regulator 

Purge Valve 

Water Separator 

Pressure Sensor 



 

Line Heaters 

Isolation (Solenoid) 

Valve 

Vent to 

Space 
27 

Table 11 shows: (a) assumptions made to calculate space environment disturbances and (b) 

the resulting disturbance torques. The center of mass and moments of inertia were estimated by 

considering the capsule as a conical frustum with a small top radius (0.1 m), and the vehicle’s 

center of pressure offset was assumed to be 5% of the largest diameter13. Assumptions about 

each disturbance were made (e.g., field strengths and incidence angles) to yield a rough order-ofmagnitude estimate of disturbances**. 

Table 11. (a) Estimated spacecraft parameters (left), and (b) resulting disturbance torques (right). 

Parameter Value Units Disturbance Torque (N-m) 

h 350 km Solar 9.2E-06 

RE 6,378 km Atmospheric 2.9E-02 

E 3.986E+14 m3/s2 Gravity-Gradient 2.4E-03 

dlargest 4 m Magnetic Field 4.0E-05 

cps – cm (5%) 0.2 m 

dmean 2 m 

cm 1.212 m 

Ix 19,368 kg-m2

Iy 19,368 kg-m2

Iz 26,404 kg-m2

Two types of stabilization and control techniques exist: passive and active stabilization. 

Passive control techniques include gravity-gradient stabilization, where the vehicle’s inertia is 

used to keep the spacecraft pointing towards the Earth’s center, and spin-stabilization, where the 

entire spacecraft rotates to keep its angular momentum vector fixed in inertial space. Active 

stabilization involves the use of attitude sensors, attitude actuators, and typically a processor. The 

sensors take measurements to compute the spacecraft’s angular position/velocity and the 

actuators depend on an implemented control law to counter disturbance torques22. 

Passively controlled stabilization systems yield very low attitude accuracies, and since 

human safety is imperative, pointing accuracies must be high. A common, three-axis active 

stabilization system is therefore used for GNC. Typical sensors include an inertial measurement 

unit (IMU), which consists of gyroscopes, accelerometers, and sometimes magnetometers; sun 

sensors or star trackers; and GPS or an inertial navigation system (used for both altitude and 

range determination). 

The command and data handling (C&DH) subsystem is essentially the “brains” of the 

spacecraft and has primary control over all functions and operations. Key components include 

flight computers, software, and data buses. The crew interface subsystem consists of all hardware 

that the crew uses to interact with the spacecraft, including keyboards, monitors, hand 

controllers, and display units. Finally, the communications and tracking subsystem provides twoway communication between crew and mission control. Antennas, transponders, and supporting 

equipment comprise this subsystem. Table 12 provides the breakdown of all avionics equipment 

required for these subsystems. Redundancies are included where appropriate. The total mass for 

the avionics system is about 400 kg. 

 

** Calculations were made from equations in Brown13 and SMAD21
28 

Table 12. Avionics components mass breakdown23-29. 

Item Manufacturer Mass 

(kg) 

Power 

(W) 

Volume 

(m3) 

Guidance, Navigation, Control 

Sensor Monitor Unknown 10 55 - 

Integrated GPS/INS (3) Honeywell 28.5 45 0.016 

Star Tracker Package (2) Ball Aerospace 12 16 0.01 

VNS Lidar Ball Aerospace 12 60 0.018 

Command and Data Handling 

Vehicle Management Computers (4) Honeywell 60 900 0.05 

VMC ISS Interface Cards (4) Generic 0.4 0 0 

Mass Memory Units (4) Unknown 10 50 0.02 

Multiplexers/Demultiplexers (19) Honeywell 152 532 0.19 

Data Bus Isolation Amplifiers (4) Singer Electronics Systems 

Division 6.8 0 0.003 

Data Buses (4) Unknown 28 0 - 

Crew Interface 

Remote Interface Units (2) ACE 1.8 30 0.002 

Keyboards (2) Generic 1.8 0 0.0048 

Monitors (2) Generic 8 0 0.01 

Hand Controllers 

(2 rotational, 2 translational) Generic 8 0 0.002 

Head-Up Displays (2) SAAB 24 35 - 

Communications and Tracking 

Integrated S-Band Transponders (2) Thales Alenia Space 5.2 30 0.015 

Filters/Switch Diplexers Generic 2 - - 

Power Amplifiers (2) Generic 9 0 0 

Low-Noise Filter Generic 0.1 0 0 

S-Band Parabolic Antennas (2) Generic 18 100 0.2 

Switching Mechanism Parts (2) Generic 0.2 0 0 

TOTAL 398 1853 0.54 

5.6. Propulsion 

The propulsion system connects the propellants to both the thrusters and the main engine, 

and consists of the fuel and oxidizer (and their tanks); pressurant; hardware components such as 

valves, filters, and transducers; mounting hardware, insulation, and plumbing; and a hydraulic 

thrust vectoring control system (TVCS)13. MMH and NTO were already selected for the fuel and 

oxidizer, respectively, and helium was selected as the pressurant. MERs were used to size tanks 

for all three consumables. 

The system can be subdivided into two categories: the main propulsion system (MPS) and 

reaction control system (RCS). The latter includes the bi-propellant thrusters used for attitude 

control, and the former covers everything else. The main engine performs the deorbit burn while 

the thrusters are used for orbital corrections and re-entry orientation. Typically, 12 thrusters are 

needed for full three-axis control—four for each of the roll, pitch, and yaw maneuvers17 (two in 

each direction). This design includes 24 thrusters: 12 on the CM and 12 on the SM. The CM 

thrusters serve as diverse backups to the SM thrusters during orbital maneuvers, and only 6 

thrusters are needed for re-entry. This enables a fully redundant RCS. 
29 

A generic propulsion-RCS schematic for this spacecraft is shown in Figure 14. The 

propellant feed lines are connected to the main engine and the 24 thrusters. Table 13 summarizes 

the masses for all items in the propulsion system. The total mass of the system is about 860 kg. 

Figure 14. Propulsion-ADCS schematic. 

Table 13. Propulsion system mass breakdown. 

Item Mass (kg) Volume (m3) 

Main Propulsion System 

Main Engine 100 - 

Hydraulic TVCS 150 - 

Fuel Tank 176.3 0.167 

Oxidizer Tank 245.8 0.167 

Pressurant 27.3 0.036 

Plumbing (Valves, Filters, etc.) 43.35 - 

Mounting Hardware, Insulation, etc. 70 - 

Reaction Control System

Bi-Propellant Thrusters (24) 48 - 

TOTAL 861 0.37 

TO SM 

	 



	 
 

 
!


!


!



!

 

	 





# 

	"
!

 "

"

SM Thrusters 

TO SM 

THRUSTERS 
30 

5.7. Crew Accommodations 

Crew accommodations are the elements and procedures that most directly serve human 

needs, such as food, clothing, hygiene and housekeeping supplies, workstations, maintenance 

hardware, sleep and recreational equipment, and others. The mass, power, and volume required 

for these accommodations are strongly dependent on the number of crewmembers and mission 

duration (often as a function of persons, days, or both). For the short-duration mission presented 

here, accommodations are kept to a minimum and many noncritical resources are excluded. As 

previously mentioned, most of the accommodations are powered for only a fraction of the day (at 

most 8%), so the total power required is just the maximum power among all items. Table 14 

provides a summary of the necessary crew accommodations. Including space suits, the total mass 

is around 640 kg. 

Table 14. Crew accommodations mass breakdown. 

Item Description Mass 

(kg) 

Power 

(W) 

Volume 

(m3) 

Galley and Food Systems Food/warmer, kitchen/oven cleaning 

supplies, sink/spigot, cooking/eating supplies 74.15 0 0.286 

Personal Hygiene Handwash/mouthwash faucet, personal 

hygiene kit, supplies 17.3 0 0.072 

Clothing Clothes 18.4 0 0.128 

Recreational Equipment Personal storage / closet space 40 700 0.020 

Housekeeping Vacuum (prime + 2 spares), disposable 

wipes, trash bags 18.6 400 0.315 

Operational Supplies and 

Restraints Supplies, restraints, and mobility aids 65 0 3.379 

Maintenance Hand tools/accessories, test equipment, 

fixtures, machine tools, gloveboxes, etc. 125 2,000 0.49 

Photography and Film Equipment 50 400 0 

Sleep Accommodations Sleep provisions (bags/beds, restraints, or 

rigid stations) 18 0 0.2 

Health Care Medical/surgical/dental suite 15 1,500 0.25 

Space Suits Suits for crew (4) 200 0 0.4 

TOTAL 641 2,000 5.54 

5.8. Entry, Descent, and Landing (EDL) 

Atmospheric re-entry is when a spacecraft returns from orbit, encounters the planet’s 

atmosphere, and uses the atmosphere to aerodynamically decelerate and prepare for landing. 

Space vehicles re-entering from a parking orbit can do so ballistically—that is, without 

generating any lift—if they are blunt, axisymmetric shapes. 

Re-entry heat loads require the spacecraft to have a thermal protection system (TPS) with 

ablative heat shields to protect the capsule. PICA-X, a reusable and less expensive modification 

of the PICA (Phenolic Impregnated Carbon Ablator) TPS material, was selected for the 

spacecraft. PICA is a modern TPS material, developed at NASA ARC, with low density and 

efficient ablative capability at high heat flux. PICA-X, developed more recently by NASA ARC 

and Space Exploration Technologies Corporation (SpaceX), is an improved TPS technology that 

is easier to manufacture30. TPS mass is estimated by using the material’s density and thickness 

over the spacecraft’s exposed surface area. 
31 

After initial re-entry, descent can be controlled through the use of parachutes or parafoils, 

and landing for a blunt capsule involves an ocean splashdown or ground area touchdown. 

Parafoils were chosen for this spacecraft design because they deploy like parachutes but can 

glide and steer toward a landing target. The parafoil system’s mass is a function of its sink rate 

and the vehicle mass. Typical spacecraft sink rates for parafoils are near 7.5 m/s, so this value is 

assumed. For the vehicle mass and assumed sink rate, a parafoil system with a lift-to-drag ratio 

of 3.0 and a density of 1.197 kg/m2

 will have a mass of approximately 300 kg. Lower sink rates 

or higher lift-to-drag ratios yield more massive parafoil systems. 

Table 15 summarizes approximate masses and volumes for both the TPS and parafoils. A 

backup parafoil is included in the design for redundancy since parachutes and parafoils are 

susceptible to getting tangled and can have deployment issues. The total mass is about 1,020 kg. 

Table 15. EDL mass and volume estimations. 

Item Mass (kg) Volume (m3) 

Parafoils (x2) 600 7.784 

TPS (PICA-X) 424 1.5708 

TOTAL 1,024 9.355 

5.9. Spacecraft Subsystem Summary 

The updated mass budget is presented in Table 16. Some items were shifted/rearranged or 

combined into other subsystems. The total mass is estimated to be 13,800 kg, which is a 

difference of about 0.02% from the initial estimated mass budget. 

Table 16. Updated mass budget. 

Subsystem % Mass Mass (kg) 

Structure & Mechanisms 20 2,698 

Thermal Control 4 602 

ECLSS 6 817 

Power 7 1,032 

Avionics 3 398 

Propulsion 6 861 

Crew Accommodations 5 641 

EDL 7 1,024 

Launch Abort System 39 5,367 

Payload (Crew) 3 360 

GROSS MASS 100 13,800 

6. Safety Design and Risk Analysis 

Safety design for spacecraft involves tradeoffs between component mass/volume and 

redundancy. Risk and reliability analyses employ statistical methods to estimate the reliability of 

a system in order to mitigate risk. Two approaches to determining the system reliability are faulttree analysis and subsystem analysis. The former calculates reliability by breaking the mission 

into multiple phases and identifying different “hazards” or events that can jeopardize the crew’s 

safety within each phase. For example, the mission can be decomposed into the launch, orbit, 
32 

and re-entry phases, with the main event being loss of mission (LOM) or loss of crew (LOC) 

(Figure 15). Each phase is a contributing event and all are separated by an “OR” logic gate since 

they are treated as mutually exclusive events. Each of those contributing events would have their 

own contributing sub-events of major failures that can occur in that mission phase. Probabilistic 

risk assessment (PRA) is then performed by determining the reliabilities of each subcomponent 

and working up towards the main event to compute its overall probability. 

Figure 15. Fault tree of main event. 

The subsystem analysis, which is employed here, considers each subsystem separately for all 

mission phases. The total risk exposure time, as well as the system response to failure (spares, 

thresholds, etc.), is determined for every assembly and component of a system. The analysis 

determines each assembly and component’s risk/reliability for the whole mission and then works 

up to the subsystem reliabilities and, eventually, the system reliability. 

6.1. Spacecraft Reliability 

Analyzing systems involves using combinations of basic reliability constructs to estimate a 

unit’s reliability. These constructs apply for units that are in series, in parallel, on standby, or 

cross-linked. The true reliability of a unit can be estimated by: 

    (22) 

where t is the operation time and the failure rate, , can be inputted directly if a component’s 

failure rate is given or can be estimated as the inverse of the mean time between failure (MTBF). 

Most failure rate data were taken from historical databases, the Aerospace Failure Data 

Handbook31, or from the manufacturer’s site if a specific component was selected19,24-25,32-35. 

The risk exposure times are separated into the different mission phases: ascent, pre-docked, 

docked (on and off), post-undocked, and EDL. The pre-docked phase is assumed to be the 24-

hour mission in orbit, and the docked phase is assumed to be a six-month period where the 

spacecraft is docked to the ISS. Another assumption is that critical subsystems/components will 

be turned on once a week for an hour during the docked period for routine checkout (docked on). 

Also, the post-undocked period involves a cold soak and de-orbit burn until the capsule reaches 

proper entry conditions. Table 17 shows the risk exposure time breakdown. 







  	
33 

Table 17. Risk exposure time for each mission phase. 

Mission Phase Risk Exposure Time (hrs) 

Ascent 0.154 

Pre-Docking 24 

Docked (Off) 4,366 

Docked (On) 26 

Post-Undock 3.5 

EDL 0.5 

TOTAL 4,420.15 

The system response to failure for each component includes number of cold/hot spares, abort 

threshold and abort time, LOC threshold, and any LOM/LOC diverse backups (non-identical 

redundancies) for that component. Hot spares are identical redundant backups that are kept 

powered for seamless operation transition if the primary component fails, while cold spares are 

not powered and have to be “switched” following primary component failure. The abort 

threshold is the number of components that should remain before the mission is aborted; if it is 1, 

then the mission aborts when one component remains. The abort time is the time it takes the 

spacecraft to safely land following abort initiation, which could occur at various times 

throughout the mission. Similar to the abort threshold, the LOC threshold is the number of 

components that remain to have a loss of crew. For most components, this number is 0, meaning 

that when all the components of a type fail, the crew will not be expected to survive the mission 

(worst-case scenario). Some components, however, are not critical to crew survivability and do 

not have a LOC threshold. Diverse backups are redundant spares that are non-identical to the 

primary component, meaning they come from different manufacturers and/or use different 

methods to achieve the same function. These are important because of “common cause” failures, 

which occur when a single failure mode affects the operation of multiple devices that would 

otherwise be considered independent (e.g., software, manufacturer defect, or environment). 

The mission is designed such that the ECLSS consumables have separate systems for the 

command and service modules and none of the command module reserves are used prior to 

undocking. If SM consumables experience a failure, the mission aborts and the crew uses CM 

reserves to return to Earth. 

Additional events and hazards are included in the reliability analysis, such as software 

failure, rendezvous/docking failure, hatch opening failure, fire, micrometeoroid and orbital 

debris (MMOD) penetration, bird strike, and space radiation. Table 18 provides a summary of 

the LOM and LOC risk and reliability results for the spacecraft system. Some components were 

shifted to other subsystems because of their failure modes (e.g., parafoils moved to 

‘mechanisms’), and some components with noncritical failures were not modeled. 

The analysis shows that there is a roughly 1 in 52 chance for loss of mission and 1 in 1,038 

chance for loss of crew. Redundancies were included in the system (as represented by previous 

mass breakdowns) to achieve this desired outcome. The TPS was not modeled since that would 

require thermal analysis that is beyond the scope of this project. 
34 

Table 18. LOM and LOC risk and reliability analysis results. 

Subsystem 

Risk Reliability 

LOM LOC LOM LOC 

LV Equipment 1.3473E-05 1.3473E-05 9.9999E-01 9.9999E-01 

Structure and Mechanisms 8.5076E-04 1.7181E-04 9.9915E-01 9.9983E-01 

Power 3.3384E-03 1.0120E-04 9.9666E-01 9.9990E-01 

Propulsion 1.5461E-03 3.0901E-04 9.9845E-01 9.9969E-01 

Avionics 1.8332E-03 4.2607E-05 9.9817E-01 9.9996E-01 

ECLSS 1.0271E-03 6.0440E-05 9.9897E-01 9.9994E-01 

TCS 3.1590E-03 5.2389E-05 9.9684E-01 9.9995E-01 

Crew Accommodations 1.2387E-04 1.2388E-05 9.9988E-01 9.9999E-01 

Events and Hazards 7.5814E-03 2.0049E-04 9.9242E-01 9.9980E-01 

TOTAL 1.9327E-02 9.6344E-04 9.8067E-01 9.9904E-01 

6.2. Launch Vehicle Reliability 

Launch vehicle reliability involves failure modes associated with launch sequence events. 

These include stage separation failure, re-contact during staging, engine startup failure, loss of 

control, structural failure, and tank (pressure vessel) failure. Table 19 shows the risk and 

reliability results for critical LV failure modes. 

The analysis shows that there is a roughly 1 in 85 chance for launch vehicle failure. Failure 

rates for contained/uncontained engine failures were allocated to have the design fit the existing 

class of launch vehicles. A detailed risk analysis is beyond the scope of this report and the work 

presented here is not meant to represent a full or competitive risk assessment. 

Table 19. LV risk and reliability results31,36. 

Stage(s) Failure Mode Risk Reliability 

1 Contained Engine Failure 4.00E-03 9.96E-01 

1 Loss of Control (engine out + gimbal) 3.24E-04 9.9968E-01 

1 Loss of Control (avionics + software) 7.41E-05 9.9993E-01 

1 Catastrophic Engine Failure 1.09E-03 9.9891E-01 

1 Structural Failure 3.40E-06 1.0000E+00 

2 Structural Failure 3.40E-06 1.0000E+00 

1 LOX Tank Failure 8.20E-06 9.9999E-01 

1 RP Tank Failure 9.70E-06 9.9999E-01 

2 LOX Tank Failure 8.20E-06 9.9999E-01 

2 LH2 Tank Failure 9.70E-06 9.9999E-01 

1,2 Failure to Stage 5.00E-04 9.9950E-01 

1,2 Re-contact at Staging 0.00E+00 1.0000E+00 

2 Contained Engine Failure 4.50E-03 9.9550E-01 

2 Loss of Control (engine out + gimbal) 3.24E-04 9.9968E-01 

2 Loss of Control (avionics + software) 7.41E-05 9.9993E-01 

2 Uncontained Engine Failure 1.00E-03 9.9900E-01 

TOTAL 1.1873E-02 9.8813E-01 
35 

7. Summary 

This memorandum has presented generic spacecraft and launch vehicle designs developed 

for a mission delivering four crewmembers and equipment to the ISS. The spacecraft subsystems 

were designed to the component level for a 24-hour on-orbit mission with a 7-day contingency, 

and redundancies were implemented to achieve a LOM probability of less than 1 in 50, and a 

LOC probability of less than 1 in 1,000. A TSTO launch vehicle was iteratively designed to 

achieve a 350-km orbit by using a combination of MERs and trajectory analyses. Reliability 

analysis estimates the launch vehicle failure probability to be around 1 in 85. Future work will 

include higher-fidelity launch trajectory analysis, analysis of launch vehicle loads during flight, 

and development of a more detailed architectural layout for the spacecraft. The vehicle designs 

presented are not intended to represent a viable architecture, but rather to provide a nonsensitive, non-proprietary platform for risk analysis studies and development efforts. This 

generic space launch architecture will be used to perform risk trade studies and to develop, test, 

and demonstrate innovative Engineering Risk Assessment capabilities at NASA Ames Research 

Center. These efforts will provide insight into key system risk sensitivities, factors affecting risk 

analysis estimates, and how to most effectively capture the impactful risk information needed to 

design safe human spaceflight vehicles. 

References 

1. "Commercial Crew Program." NASA Facts. National Aeronautics and Space Administration, 2011. 

http://www.nasa.gov/pdf/609181main_12.08.11_CCP.pdf 

2. Man-Systems Integration Standards. NASA STD-3000. National Aeronautics and Space 

Administration. Houston, TX: 1995. 

3. Mass Estimating and Forecasting for Aerospace Vehicles. Systems Definition Branch. National 

Aeronautics and Space Administration. Houston, TX: 1994. 

4. Larson, Wiley J., L. Pranke. Human Spaceflight – Mission Analysis and Design, McGraw-Hill, NY: 

2000. 

5. "Single Impulse De-orbit from Circular and Elliptical Orbits ." Orbital and Celestial Mechanics. C. 

David Eagle, 09 Dec 2012. http://www.cdeagle.com/pdf/deorbit.pdf 

6. "Space Suit Evolution: From Custom Tailored To Off-The-Rack." NASA History Program Office. ILC 

Dover, Inc.: 1994. http://history.nasa.gov/spacesuits.pdf 

7. Lynn, Emory E., INTROS User’s Manual, 2007. 

8. Edberg, Donald. ARO 409: Introduction to Launch Vehicle Design and Systems Engineering (class 

notes). California State Polytechnic University, Pomona. Aerospace Engineering Department. 

Pomona, CA: 2012. 

9. Curtis, Howard D. Orbital Mechanics for Engineering Students. 1st ed. Burlington, MA: Elsevier, 

2005. 

10. "Rocket Mass Characteristics V4.6." EarthLink. Earthlink, Inc., 30 Dec 2009. 

http://home.earthlink.net/~apendragn/atg/coef/Structural_Coef.pdf 

11. Rao, A. V., Benson, D. A., Darby, C., Patterson, M. A., Francolin, C., Sanders, I., and Huntington, G. 

T. 2010. “Algorithm 902: GPOPS, A MATLAB software for solving multiple-phase optimal control 

problems using the Gauss pseudospectral method.” ACM Transactions on Mathematical Software 37, 

2, 22:1–22:39 

12. Riehl, J. P., Paris, S. W., and Sjauw, W. K., “Comparison of Implicit Integration Methods for Solving 

Aerospace Trajectory Optimization Problems,” AIAA 2006-6033, 21-24 Aug. 2006. 

13. Brown, Charles D. Elements of Spacecraft Design. Reston, VA: American Institute of Aeronautics and 

Astronautics, 2002. 
36 

14. Dumoulin, Jim. "Mission Preparation and Prelaunch Operations." NSTS 1988 News Reference 

Manual. NASA Kennedy Space Center: 1988. 

15. Sawko, Paul M., and Howard E. Goldstein. "Performance of uncoated AFRSI blankets during multiple 

Space Shuttle flights." NASA STI/Recon Technical Report N 92 (1992): 29104. 

16. Stephan, Ryan A. "Overview of NASA's Thermal Control System Development for Exploration 

Project." NASA Technical Reports Server. (2010).

17. Edberg, Donald. ARO 491L: Introduction to Spacecraft Design (class notes). California State 

Polytechnic University, Pomona. Aerospace Engineering Department. Pomona, CA: 2011. 

18. “International Space Station: Environmental Control and Life Support System." Boeing Defense, 

Space, and Security. The Boeing Company. http://www.boeing.com/boeing/defensespace/space/spacestation/systems/eclss.page 

19. "Fuel Cell Products: FC Velocity 9SSL." Ballard. Ballard Power Systems, Inc.: 2011. 

http://www.ballard.com/fuel-cell-products/fc-velocity-9ssl.aspx 

20. "PEM Fuel Cell Product Portfolio." Ballard. Ballard Power Systems, Inc.: 2011. 

21. Wertz, James R., David F. Everett, and Jeffrey J. Puschell (Editors). Space Mission Engineering: The 

New SMAD. Hawthorne, CA: Microcosm, Inc., 2011. 

22. de Ruiter, Anton H., Christopher Damaren, and James R. Forbes. Spacecraft Dynamics and Control: 

An Introduction. West Sussex, UK: John Wiley & Sons, Ltd., 2013. 

23. Petty, John I.. "HSF - The Shuttle." Human Spaceflight. National Aeronautics and Space 

Administration, 07 Apr 2002. http://spaceflight.nasa.gov/shuttle/reference/shutref/verboseindex.html 

24. "SIGI - Space Integrated GPS/INS." Honeywell. Honeywell International, Inc.: 2006. 

25. "CT-602 Star Tracker." Ball Aerospace. Ball Aerospace & Technologies Corp. 

http://www.ballaerospace.com/file/media/D0540_CT-602.pdf 

26. Christian, John A., Heather Hinkel, Christopher N. D'Souza, Sean Maguire, and Mogi Patangan. "The 

Sensor Test for Orion RelNav Risk Mitigation (STORRM) Development Test Objective." American 

Institute of Aeronautics and Astronautics: 2011. 

27. "ACE Remote Interface Unit (ACE-RIU)." ASTi Support. Advanced Simulation Technology, Inc.: 2013. 

http://support.asti-usa.com/hardware/ace_riu.html 

28. "Advanced, Light-weight, Low-cost Head-up Display System." SAAB Group. SAAB AB: 2014. 

http://www.saabgroup.com/Global/Documents%20and%20Images/Air/Avionics%20Systems/RIGS/RI

GS%20product%20sheet.pdf 

29. "Integrated S-Band Transponder (ISBT)." Thales Group. Thales Alenia Space. 

https://www.thalesgroup.com/sites/default/files/asset/document/ISBT.pdf 

30. Chambers, Andrew, and Dan Rasky. "NASA + SpaceX Work Together." ASK Magazine. 17 Oct 2010. 

http://appel.nasa.gov/2010/10/17/40s_space-x-html/ 

31. Fragola, Joseph R. Aerospace Failure Data Handbook. Valador, Inc., 2010. 

32. "Flash Mass Memory Unit." SSTL. Surrey Satellite Technology Ltd., n.d. Web. 

http://www.sstl.co.uk/getattachment/115a57bc-8b47-4977-9a8f-1b543d02e6ab/Flash-Mass-Memoryunit--FMMU33. Wieland, P., C. Hutchens, D. Long, and B. Salyer. United States. National Aeronautics and Space 

Administration. Final Report on Life Testing of the Vapor Compression Distillation/Urine Processing 

Assembly (VCD/UPA) at the Marshall Space Flight Center (1993 to 1997). Huntsville, AL: Marshall 

Space Flight Center, 1998. 

34. "Understanding Your Product Through Reliability Modeling." Reliability Ques. Vol. 3.2 Quanterion 

Solutions Incorporated: 2004. 

35. Seedhouse, Erik. SpaceX: Making Commercial Spaceflight a Reality. New York, NY: Springer, 2013. 

36. National Aeronautics and Space Administration. NASA's Exploration Systems Architecture Study: 

Final Report. NASA-TM-2005-214062. November 2005.