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AMERICAN PUBLIC UNIVERSITY SYSTEM

Charles Town, West Virginia

HIGH LIFT ENTRY VEHICLE DESIGN FOR

LANDING HIGH MASS PAYLOADS ON MARS

A thesis submitted in partial fulfillment of the

requirements for the degree of

MASTER OF SCIENCE

in

SPACE STUDIES

by

John Clayton

Department Approval Date:

______________

The author hereby grants the American Public University System the right to display thesecontents for educational purposes.



The author assumes total responsibility for meeting the requirements set by United StatesCopyright Law for the inclusion of any materials that are not the author’s creation or in the public domain.



DEDICATION I dedicate this thesis to my wife and mother. Without their patience, understanding, support, and

most of all love, the completion of this work would not have been possible.



ABSTRACT OF THE THESIS

HIGH LIFT ENTRY VEHICLE DESIGN FOR

LANDING HIGH MASS PAYLOADS ON MARS

by

John Clayton

American Public University System, ________, 2011

Charles Town, West Virginia

Professor Robert Thrower, Thesis Professor

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TABLE OF CONTENTS

CHAPTER PAGE

I. INTRODUCTION ...........................................................................................1 II. APPROACH …................................................................................................5

Atmosphere ......................................................................................................5Elevation ...........................................................................................................8Terrain ..............................................................................................................9Landing Sites ....................................................................................................9

II. MARS ENVIRONMENT ................................................................................5

Atmosphere ......................................................................................................5Elevation ...........................................................................................................8Terrain ..............................................................................................................9Landing Sites ....................................................................................................9

III. MISSION ASSUMPTIONS ..........................................................................24

Launch System …............................................................................................30Reusability …...................................................................................................27Interplanetary Transfer ….................................................................................25Mars Arrival ....................................................................................................27

Aerocapture vs Direct Descent….....................................................................34Landing ….......................................................................................................34

III. VEHICLE DESIGN .......................................................................................24

X-33 Scaled …...........….................................................................................25TPS Requirements ...........................................................................................27Propulsive Requirements …..............................................................................27

IV. SIMULATIONS ............................................................................................34

Aerocapture ....................................................................................................35Descent ….......................................................................................................36Landing ….......................................................................................................38

V. CONCLUSIONS AND FUTURE WORK ....................................................49 LIST OF REFERENCES ...........................................................................................60



APPENDICES ...........................................................................................................66



LIST OF TABLES TABLE PAGE 1. Physical Education Teacher Demographic Data .......................................................15

2. Current University Student Demographic Data .........................................................17 3. Number of High or Low Value Orientations for Respondents ...................................25 4. Teacher Value Orientation Profile by Gender ...........................................................28 5. Teacher Value Orientation Profile by Academic Rank ..............................................33

6. Teacher Value Orientation Profile by Teaching Experience .......................................39 7. Student Value Orientation Profile by Gender ............................................................41 8. Student Value Orientation Profile by Academic Major ..............................................45 9. Student Value Orientation Profile in Different Year at University ................................51



LIST OF FIGURES FIGURE PAGE 1. Physical Education Teacher Demographic Data ........................................................15

2. Current University Student Demographic Data .........................................................17 3. Number of High or Low Value Orientations for Respondents ...................................25 4. Teacher Value Orientation Profile by Gender ...........................................................28 5. Teacher Value Orientation Profile by Academic Rank ..............................................33

6. Teacher Value Orientation Profile by Teaching Experience .......................................39 7. Student Value Orientation Profile by Gender ............................................................41



LIST OF SYMBOLS SYMBOL



LIST OF ACRONYMS ACRONYM

(counted and numbered)



INTRODUCTION

The Martian atmosphere poses a number of challenges for spacecraft entry, the most

important of which are the low density and shallowness, both of which limit the aerodynamic

braking available for entering spacecraft. Further complicating the problem is the variability of

the atmospheric thickness, and the possibility of dust storms.

Up to the present, all Mars lander missions have use variations of the entry, descent,

and landing (EDL) technology created for the Viking missions. That technology consists of a

capsule shape with jettison-able heat shield for atmospheric entry, a parachute to reduce velocity

from supersonic to subsonic, and a propulsive final landing phase. That technology is limited

to low landed masses in areas of low elevation and does not scale up, as larger masses do not

decelerate to velocities suitable for parachute deployment in sufficient time.

Several technologies have been proposed, and are under

study, to overcome the Mars high mass EDL problem. These

include the use of inflated structures, “ballutes”, either

attached or towed by the craft, to increase the drag profile,

ballistic descent with propulsive landing, and a high lift entry

design which could “fly” in the Mars atmosphere for a longer

descent profile, with a propulsive landing phase. This high

lift approach, originally proposed by Wernher von Braun in

his 1952 book Das Marsprojekt , has not been extensively

studied, and is the focus of this paper.

APPROACH



The objective of this paper is to explore the feasibility of a

high-lift approach to landing high-mass payloads on Mars,

not to perform a trade study of different approaches and

configurations. To that end, a number of assumptions and

design configuration choices were made in order to limit

the scope of the research while still addressing the main

objective.

Entry Vehicle Choice

A number of high-lift designs have been researched and tested in the past, including delta-winged

vehicles such as the Space Shuttle, and lifting body designs dating from the 1960’s. Later lifting body

designs include the X-33 Single Stage to Orbit and X-38 Crew Return Vehicle flight demonstrators. For

this paper, a scaled down version of the X-33, denoted “X-33S”, is chosen as the entry vehicle. The X-

33 has an extensive database of aerodynamic testing and simulation research available that are applicable

to Mars high-lift entry simulations. It has hypersonic aerodynamic characteristics that seem appropriate

for the Mars EDL problem, including lift generation and stability at high angles of attack, and has the

high volumetric efficiency needed for a payload vehicle.

Entry Vehicle Configuration

Many configurations of spacecraft are conceivable for the interplanetary journey from

the Earth to Mars. For this paper, it is assumed that the X-33S is sized to fit a present-day

heavy launch vehicle, such as the SpaceX Falcon 9 Heavy which can place 50,000 kg into LEO.

Restricting the vehicle size to the width of the Falcon 9 Heavy, and allowing for a 20 percent

decrease of LEO capability due to increased drag, places an upper mass limit of about 40,000 kg



on the X-33S, and a maximum width of approximately 13 m, with on-board fuel limited to that

required for Mars entry. It is further assumed that external fuel tanks are launched separately,

are then attached to the X-33S, and provide the delta-V for a Hohmann transfer orbit, using on-

board X-33S rocket engines.

Mars Arrival

Given a Hohmann transfer, the Mars arrival velocity will be approximately 7.3 km/s.

Previous studies show that aerocapture into orbit and subsequent descent, has advantages in

thermal loading and landing precision over a direct descent. For this paper, it is assumed that

an aerocapture into orbit is used, and that the X-33S TPS must be dual-capable of handling

the thermal environment of both aerocapture and descent. The aerocapture loads are restricted

to a maximum of 5 g’s, which is presumed to be the maximum tolerable by de-conditioned

astronauts.

Descent and Landing

The descent profile is proposed to be in three stages, all within a 5 g load limit. The

first phase is a gliding descent from orbit utilizing the high lift characteristics of the X-33S to

achieve maximum deceleration to near the surface. The second phase uses rocket propulsion

to maintain altitude during a Pugachev Cobra maneuver, in which the vehicle pitches up past

vertical, and uses the forward component of the rocket propulsion vector to reduce the horizontal

velocity to zero while still maintaining altitude. The final phase is a vertical landing using rocket

propulsion.

Mars Environment



NASA has developed a model of the Martian atmosphere known as Mars Global

Reference Atmospheric Model, or MarsGRAM. MarsGRAM has been used for the entry

calculations and guidance for all of the recent NASA Mars missions. It models seasonal, dust

storm, time-of-day, and several other parameters. For the goals of this paper, a previous NASA

simplified model of the Mars atmosphere is sufficient, which corresponds to average parameters

of MarsGRAM, and in which the atmospheric density is modeled in two parts by:

ρ = p / (0.1921T ) where p = 0.699e-0.00009h

ρ = density (kg/m3), p = pressure (K-Pa), T = temperature (oK), h = altitude (m)

For 0 ≤ h ≤ 7000: T = -31 - 0.000998h

For 7000 < h ≤ 30500: T = -23.4 - 0.00222h

Incorporating the atmospheric density model into the equation for the speed of sound, a,

at altitude yields:

a = √(⋎RT), ⋎ = 1.29 the specific heat ratio (dimensionless), R = 191.8 J/kg/K

For 0 ≤ h ≤ 7000: a = √(-7670.1 - 0.2469h))

For 7000 < h ≤ 30500: a = √(-5789.7 - 0.5492h))

The atmospheric model parameters are shown in the following graphs:



VEHICLE PARAMETERS

X-33

The X-33 is a lifting body design that was intended as a technology demonstrator for

Single Stage to Orbit (SSTO) operations. Although cancelled due to budgetary and technology

problems, a number of studies were made of its aerodynamic performance, including scale model

flights. The results of those studies form the basis for the vehicle design considered in this

paper.

The X-33 planform reference geometry and aerodynamic parameters are given in Table

X.X.

Parameter Value Unit

Empty Mass, m 28,440 kg

Reference Surface Area, S ref 149.39 m2

Ref. Aerodynamic Span, b 11.16 m



Ref. Aerodynamic Chord, c 19.26 m

Ref. Aspect Ratio, AR 0.86 -

Sweep 70 degrees

The X-33 aerodynamic coefficients are modeled by second order polynomials from Mach

0.4 through Mach 20, as a function of the angle-of-attack ⍺ up to a maximum of 50o, and the

Mach number M :

C L(α, M ) = -0.0005225α2 + 0.03506α - 0.04857 M + 0.1577

C D(α, M ) = 0.0001432α2 + 0.00558α - 0.01048 M + 0.2204

X-33 Scaled

The relationship between the aerodynamic parameters of a scaled down model and a full

size vehicle depend on a number of factors, principally whether the flow is incompressible or

compressible, and the gravitational field. Those scaling factors are shown in Table X.X.

Scaling Factor Incompressible Flow Value Compressible Flow Value

Linear dimension n n

Relative density (m/ ρl 3) 1 1

Mach number 1 1

Froude number (V 2/lg ) 1 1

Angle of attack 1 Dependent on Froude scalingLinear acceleration 1 1

Weight, mass n3/σ n3/σ

Moment of inertia n5/σ n5/σ



Linear velocity n1/2 n1/2

Angular velocity 1/n1/2 1/n1/2

Time n1/2 n1/2

Reynolds number n1.5v/v0 n1.5v/v0

For unaccelerated flight conditions, the lift coefficient is given by

C L = W / ½ ρV2S = 2mg / ρV2S

Note that for Mars, g Mars = 0.38 g , and so to be able to apply the X-33 aerodynamic data

to the Mars environment directly, the X-33S reference area must have a scale factor of 0.38 also,which implies a linear scale factor of 0.62. The difference in Reynolds number has a primary

effect on the boundary layer separation characteristics at high angle of attack conditions, and is

highly configuration dependent. Because the main effect is that flow separation of the model

is delayed to a higher angle of attack in the model, validity of the X-33 data for the X-33S near

the maximum angle of attack is a reasonable assumption. The X-33S reference geometry and

aerodynamic parameters are given in table X.X, for a linear scale factor of 0.62, which gives a

vehicle size approximately correct for the Falcon 9 Heavy launch vehicle.

Parameter Value Unit

Reference Surface Area, S ref 149.39,56.77

m2

Ref. Aerodynamic Span, b 6.92 m

Ref. Aerodynamic Chord, c 11.94 m

Ref. Aspect Ratio, AR 0.86 -

Sweep 70 degrees



The X-33 L/D range from approximately 1.25 to 4.0 for a Mach range of 0.1 to 30 respectively.

TPS Requirements

Propulsive Requirements

The entry profile to maximum the effect of lift for deceleration ends in flight near the

surface at a velocity where aerodynamic lift is insufficient to maintain altitude???

Rocket propulsion is then used to maintain altitude while using the Pugachev Cobra

maneuver to decrease the velocity to zero. Determining the transition point is the subject of the

following paragraphs.

Plotting these coefficients shows that L/D becomes less than 1 at Mach 0.2 and



Assuming that an altitude of 1000 m provides sufficient terrain clearance for the duration

of the Pugachev Cobra maneuver, the

At an altitude of 1000 m, the speed of sound is 243 m/s.

The aerodynamic coefficients during supersonic flight in compressible flow of a scaled

model are approximately the same as the full size vehicle. The X-33 aerodynamic coefficients

during supersonic flight will therefore approximate those of the X-33S, and are modeled by

second order polynomials from Mach 1 through Mach 20, as a function of the angle-of-attack ⍺

up to a maximum of 50o, and the Mach number M :

ANALYSIS

Aerocapture

Descent



Landing

VEHICLE PARAMETERS

X-33 Scaled

TPS Requirements

Propulsive Requirements

CONCLUSIONS AND FUTURE WORK

REFERENCES

Bibliography

Appendices



Purpose Statement

This thesis will propose and explore the feasibility of a high lift Mars entry vehicle

design to allow precision landing of high mass payloads. It will include a review the current

literature for the status of key technology areas, including prior terrestrial and Mars EDL

systems, approach navigation, thermal protection systems, direct EDL versus aerocapture,

atmospheric entry dynamics, propulsive landing dynamics, and characterize an optimal vehicle

architecture. Data from prior terrestrial and Mars missions will be used in the evaluation of these

areas and vehicle design, along with software to simulate and evaluate approach, atmospheric

entry, and landing dynamics.Statement of the Problem

Significance of the Research

This thesis will help determine whether a high lift Mars entry design is a feasible option

for supporting a manned mission to Mars.

Initial Research

The author has reviewed literature in the technology areas related to this thesis. The

results of this initial research will guide the thesis, and include the following initial observations.

Current Mars EDL limits. Proven Mars entry, descent, and landing systems are

bounded by the Viking parachute system at around 900 kg landed mass with a landing accuracy

of 20 km. The Viking system uses disk-band supersonic parachutes that are qualified to 19.5m

diameter, and Mach 2.2 opening velocities. This system can be extended to marginally higher



masses through qualification of larger parachutes, stronger materials for higher deployment

velocities, and higher temperature materials.



Mars atmospheric entry. A fundamental design trade-off is whether to perform

the EDL directly from the interplanetary approach, or to first enter orbit with an aerocapture

maneuver followed by a subsequent EDL. This trade-off has been well studied, and the result

for capsule entry systems has been in favor of direct EDL. This is largely because the different

thermal protection requirements for aerocapture lead to the necessity of having two separate

heat shields, one for each phase. Newer thermal protection systems (TPS) may make feasible

the use of a single TPS for both aerocapture and subsequent EDL. Advantages of aerocapture

over direct EDL include better characterization of atmospheric conditions prior to EDL, and

more precise EDL targeting. A major consideration of the entry design for manned vehiclesis the need to stay below a 5-g maximum loading for astronauts de-conditioned by the long

interplanetary transit time.

Prior high lift entry vehicle designs. The Earth’s upper atmosphere provides an analog

for testing a Mars entry vehicle. A number of high lift vehicles have been designed for Earth

reentry, including lifting bodies, the space shuttle, and the X-33. These designs provide a wealth

of data on hypersonic aerodynamics, including stability and thermal protection, that is applicable

to Mars EDL.

Landing. This thesis will explore the feasibility of a high lift vehicle that “flies” as long

as possible through the lower Martian atmosphere for deceleration from the hypersonic entry.

Once the angle of attack lift limits are reached, rocket propulsion will provide the additional

lift vector needed to maintain altitude, and provide the final deceleration through a “Pugachev

Cobra” maneuver. That maneuver involves increasing the angle of attack past vertical for final

deceleration, then a return to vertical for landing.



Data Analysis

Aerocapture flight dynamics analysis will be performed using the fourth order Runge

Kutta algorithm of Gallais. Entry dynamics and thermal analysis will be performed using the

equations of Gallais and Wiesel. High lift hypersonic flight dynamics analysis will be based on

Newtonian flow as described by Bertin, and will derive from legacy vehicle data. Aerocapture

and entry TPS analysis will be based on legacy mission data and recent test data. Landing

analysis will derive from legacy mission data and the Pugachev Cobra flight regime. Matlab

will be used for computational analysis and visualization.

Tentative Table of Contents

Chapter 1 - Mars EDL Requirements

Chapter 2 - Legacy High Lift Entry Designs

Chapter 3 - Proposed Mars High Lift Design

Chapter 4 - Aerocapture Analysis

Chapter 5 - Entry Analysis

Chapter 6 - Landing Analysis

Chapter 7 - Conclusion

References



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