Tunable Diode Laser Absorption Spectroscopy Characterization of Impulse Hypervelocity CO

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Tunable Diode Laser Absorption

Spectroscopy Characterization of Impulse Hypervelocity CO 2 Flows

A Thesis submitted for the degree of

DOCTORAT EN SCIENCES DE L’INGENIEUR Jason Matthew Meyers

UNIVERSITE LIBRE DE BRUXELLES F

ACULTE DES

´ S

CIENCES

A

PPLIQUEES

´

VON KARMAN INSITUTE FOR FLUID DYNAMICS A

ERONAUTICS AND

A

EROSPACE

D

EPARTMENT

THESIS COMMITTEE

G´erard Degrez (ULB - Promoter) Doug Fletcher (VKI - Supervisor)

Frank Dubois (ULB - Thesis Committee President) Michel Herman (ULB)

Ajmal Khan Mohamed (ONERA) Uwe Koch (DLR)

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Tunable Diode Laser Absorption Spectroscopy Characterization of Impulse Hypervelocity CO2Flows

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Tunable Diode Laser Absorption

Spectroscopy Characterization of Impulse Hypervelocity CO 2 Flows

A Thesis submitted for the degree of DOCTORAT EN SCIENCES DE L’INGENIEUR

Jason Matthew Meyers

UNIVERSITE LIBRE DE BRUXELLES F

ACULTE DES

´ S

CIENCES

A

PPLIQUEES

´

VON KARMAN INSITUTE FOR FLUID DYNAMICS A

ERONAUTICS AND

A

EROSPACE

D

EPARTMENT

THESIS COMMITTEE G´erard Degrez (ULB - Promoter) Doug Fletcher (VKI - Supervisor)

Frank Dubois (ULB - Thesis Committee President) Michel Herman (ULB)

Ajmal Khan Mohamed (ONERA) Uwe Koch (DLR)

September 9, 2009

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To John Sanford Todd

I almost made it home in time...

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SUMMARY

Tunable diode laser absorption spectroscopy using an external cavity diode laser operating in the infra-red has been developed to monitor CO2 in the freestream of the Longshot hypervelocity facility at the Von Karman Institute for Fluid Dynamics. The Longshot facility offers a unique European facility for ground testing and numerical validation applications, however, some of the traditional data rebuilding aspects are in question. A non-intrusive absorption sensor could significantly aid in improving the knowledge of freestream static values thereby improving the models used in data rebuilding and numerical simulation. The design of such a sensor also expands the spectroscopic capabilities of the Von Karman Institute.

The absorption sensor is designed around the single P12 (00001)→(30013) ro- vibrational transition near 1.6µm (6218.09cm⁻¹specifically) which yields relatively weak direct absorption levels at about 3.5% per meter for typical Long- shot freestream conditions. However, when handled carefully, adequate signal- to-noise can be acquired to exploit significant flow information. By being able to operate in this range, total sensor cost can be easily an a factor of two or more cheaper than sensors designed for the deeper infrared. All sensor ele- ments were mounted to a compact portable optics bench utilizing single-mode optical fibers to allow for quick installation at different facilities by eliminat- ing tedious optical realigning. Scans at 600Hz were performed over 20ms of the 40ms test time to extract core static temperature, pressure and velocity.

These results are compared with the current state of the Longshot data rebuild method.

The non-uniform flow properties of the shear layer and test cabin rested gas accumulation was of an initial concern. The temperature and density gradi- ents along with significant radial velocity components could result in DLAS temperature, pressure and velocity that are significantly different than that of the target freestream inviscid core values. Fortunately, with the proper selec- tion of the P12 rotational number, this effect could be more or less ignored as the higher temperature and lower density gas of this region is relatively trans- parent.

Ultimately, acquired temperature and density were moderately accurate when compared to Longshot rebuilt results owing primarily to the baseline extrac- tion which poses issues for such low absorption signals. However, the ex- tracted velocity data are quite accurate. This is a definite plus for the sensor as the freestream enthalpy of cold hypersonic facilities is dictated primarily by the kinetic energy contribution. Being able to compare velocity gives insight to the level of vibration non-equilibrium in the flow. The velocity of the DLAS

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and the Longshot rebuild are quite close. This adds more weight to the argu- ment that vibrational excitation is very low (if present at all) in the free stream and that the van de der Waals equation of state usage and constant specific heat assumption might be an adequate model for the data rebuild after all.

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ACKNOWLEDGMENTS

Though my name is printed on the title page of this work, none of these efforts herein would have been possible without the support of many friends and colleagues. I will try my best to feign an attempt to thank all the many people who deserve credit.

First and foremost I want to thank my family for being so supportive of my being on the other side of the Atlantic for such a long period. I have missed out on so many events, some happy and some sad, that I had no control to be a part of. For your understanding of my absence, I am truly grateful.

I must also thank the Von Karman Institute and Professor Mario Carbonaro for supporting me through this work. Of course the degree was made possible thanks to the efforts of my promoter Professor G´erard Degrez at ULB.

Thank you for your patience and efforts to help a non-European through the ULB administration process. Thank you Professor Dubois for presiding over my committee. Thank you Professor Herman for your input on NIR spectra as well as your willingness to be present on my committee. Thank you Dr. Mo- hamed and Dr. Koch for making the efforts in traveling to hear me defend my work.

Thank you Doug for your unwaivering support, expertise and push to help me to implement yet another spectroscopic technique at VKI. It is a complete pleasure to say that I worked under your supervision. I cannot wait to see the beard in person.

A tremendous thanks goes to my dear friend and colleague Mica¨el Playez.

Your encyclopedic knowledge and experience helped in volumes through tough times. I cannot even begin to count the number of beers I owe you... wait you don’t drink. Is Coke OK?

To Sebastian, the keeper of beer and Longshot, thank you so much for letting me play with one of VKI’s biggest toys. That being said, where would I be if it were not for Patrick and Jerome- the heart of the Longshot maintenance crew.

To Vincent who not only gave immense technical support but also showed me a thing or two about Belgian culture... dank je wel!!

Thanks Pascal for sparing some of your busy Plasmatron schedule to help run my tests. I think that no amount of climbing can prepare one for your vice- grip-like handshake.

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To Ajmal, thank you for willingly bringing your knowledge and experience not only to my committee as mentioned above, but also to VKI. Rafael, it was a pleasure working with you, though if only for a short period. I can only hope to have the pleasure of working with such good company again.

Thank you Louis for your numerical simulations. The last minute addition of your results helped add significant weight to some of my arguments.

To one of my best buds, Raf, thank you so much for your friendship and introducing me to climbing. Terres Neuves became a sanctuary for me of sorts and I cannot appreciate you enough for introducing me to that world.

Marcos, I hope to be visiting you many more times in beautiful Galicia... Nunca M´ais!!!

Thank you Fatboy’s for my Americana fix. Autumn Sundays really helped me to get away when I needed it.

This work would not have been possible if not for the support of the Air Force Office of Scientific Research (Grant FA9550-07-1-0089, Dr. J. Schmisseur, Tech- nical Monitor), the Centre National d’Etudes Spatiales (MSRO-052, Dr. J.-M.

Charbonnier, Technical Monitor) and the Belgian American Educational Foun- dation (Prof. Emile L. Boulpaep, www.baef.be).

I know there are plenty of others, friends and colleagues alike, whom I did not mention. To those people, I thank you for your forgiveness in my forgetting your mentioning here!

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NOMENCLATURE

Latin Letters

κ Force constant

λ Wavelength

T Ro-vibrational term value θf Half-cone angle of focus

a Voigt “a” parameter

a Wavelength modulation amplitude B Rotational constant

c Speed of light

C_p Specific heat capacity at constant pressure

C_v Specific heat capacity at constant volume

D Correction term for non-rigid rotation

E Energy

e Mass specific internal energy E⁰ Upper-state state energy E⁰⁰ Lower-state energy

F Etalon finesse

F Rotational term

f Focal length

f Force from Hooke’s law model Fv Rotational term value

g Gravitaional force

g_J Statistical weight factor for population distribution of rotational levelsJ G_v Vibrational term value

g_l,i Lower state degeneracy of transitioni h Mass specific enthalpy

h Planck’s constant

H0, H1, H2, H3, ... Harmonics of the measurement I Moment of inertia

I Transmitted intensity IO Reference intensity Iinc Etalon incident intensity Itrans Etalon transmitted intensity

J Rotational quantum number

k Boltzman constant

L Pathlength

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Lres Length of resonator

M Mach number

me electron mass

n Collisional broadening exponent n Index of refraction

n Number density

N_J Population distribution of rotational levelsJ

n_L Loschmidt number

N A Numerical aperture N IR Near-infrared

P Pressure

Qelect Electronic internal partition function (partition sum)

qelect Electron charge

Qint Total internal partition function (partition sum)

Qrot Rotational internal partition function (partition sum)

Qvib Vibrational partition function (partition sum)

r Displacement from Hooke’s law model r_eq Equilibrium displacement from

Hooke’s law model

S Linestrength

cm⁻²/atm

s Entropy

S∗ Linestrength

cm⁻¹/molecule·cm⁻²

T temperature

T₀ Reference temperature T_etalon Transmitted light fraction

T_ref Reference temperature from HITRAN database

u Velocity

ugas Gas velocity

V Voigt function

V U V Vacuum ultraviolet w Voigt “w” parameter w0 Beam Diameter (1/e) wf Focal waist diameter

X Mole fraction

y Integral variable of Voigt function FWHM Full-width half maximum

HWHM Half-width half maximum

Greek Letters

α Absorption coefficient β Ballistic coefficient

∆λ_res Spacing between resonator modes

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∆ν frequency shift due to Doppler shift δ Phase shift per traversal

∆νC Collisional broadening FWHM value

∆νD Doppler broadening FWHM value

∆νF SR Etalon free spectral range γ Specific heat ratio

Λ Mean free path

λ Wavelength

µ Reduced mass

ν Frequency

ν₀ Line center frequency

νosc Oscillating frequency of a spring model

ω Angular velocity

ωm Modulation frequency φ Broadening coefficient

φC Collisional broadening coefficient φD Doppler broadening coefficient φV Voigt broadening coefficient

ρ Density

τ Spectral transmittance θ Angle of beam w.r.t. flow θ⁰ Internal angle within etalon

θ_v Characteristic vibrational temperature

˜

ν Wavenumber

d Etalon spacing or thickness tau Transmission fraction

Subscripts

0 Stagnation conditions

∞ Free stream conditions i Line transition index

i Lower state

j Specie index

k Upper state

S Shock

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PREFACE

M

^ARTIANatmosphere entry (illustrated in figure 0.1) is a complicated lo- gistics problem fraught with engineering complexities. A vehicle is launched from Earth or an Earth orbit to a destination 2 years away where communica- tion delays are around ten minutes. Once the vehicle arrives at the destination it must be decelerated and captured if the vehicle is an orbiting vehicle or con- tinue on into an entry decent and landing (EDL) phase if the vehicles is to perform surface operations.

Fig. 0.1:Artist’s rendering of Mars entry.

The EDL complexities begin even before a significant atmosphere is present.

It is desired that the entry vehicle high temperature TPS be oriented properly such that the heat shield will be in the region of the high temperature bow shock so as not to compromise the after body TPS at or near continuum conditions before the high temperature effects begin. Initial maneuvering takes place in the rarefied free molecular regime where the mean free path of molecular collisions is large compared to that of the characteristic body length leading to the familiar Navier-Stokes equations becoming incapable of representing flow physics. Because of this significantly rarefied atmosphere, entry vehicles are statically unstable for the free-molecular and much of the transitional flow regime before the continuum assumptions can be made. It is therefore difficult to control a vehicle, especially one without adequate control surfaces, to the appropriate attitude and keep it that way. Once the vehicle is oriented properly it must remain so and as the aerodynamic forces are slight, vehicles are generally stabilized by means of gyroscopic spin. The aerodynamics and atmosphere at this phase must be well-understood enough that this gyroscopic stability is not compromised or the vehicle orientation could be altered once significant aero-

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dynamic forces begin. Significant effort in this rarefied and transition regime has been done by Moss et al. [1]. As the atmospheric density increases, the transitional phase begins until continuum is met. This is where the significant bow shock forms and the tremendous heat loads arise.

The next extreme condition is illustrated in an artist’s rendering in figure 0.1. A strong bow shock that generates tremendous heat (thousands of degrees) begins to form as the vehicle arrives to a significant thickness of the Martian atmosphere at extremely high velocity with a significant wake following behind. It is this dissipation of kinetic energy into heat that slows the vehicle to a safe and desirable decent velocity.

Non-equilibrium gases are formed by the heating of the strong bow shock.

The extreme heating rates involved call for considerable thermal protection systems to protect the payload of such entry vehicles. These gases expand as they travel down the body of the vehicle but at such a speed that much of the chemistry and vibrational energy is frozen into the wake. The wake, though not experiencing as high of heating rates as the stagnation region, plays a huge role in vehicle stability. These non-equilibrium effects can significantly alter pressure distribution and heating rates. Thus, consideration from stagnation region down to the wake must be taken that this flow is not in a fully equilibrium state and frozen non-equilibrium models must be used to adequately describe vehicle aerodynamics. Understanding these phenomena are pivotal for the safe landing of such payloads as the Mars Science Laboratory Rover and the Mars Exploration Rovers illustrated in figure 0.2

Fig. 0.2: Artist rendering of the Mars Science Laboratory Rover and one of the Mars Ex- ploration Rovers illustrating the size comparison of future generation vehicles (taken from solarsystem.nasa.gov

Consider figure 0.3 as an example which illustrates the final EDL key phases of the specific case representing the Mars Science Laboratory (MSL) mission where a tethered “sky crane” approach will be used to land the most massive vehicle ever to the Martian surface. This case is presented as it is, currently, the next significant Mars mission scheduled to launch in the 2011 window.

The first phase, hypersonic entry, begins when the entry vehicle first comes into contact with a sensible atmosphere carrying significant kinetic energy. The tremendous heating during this phase can be anywhere between 25W/cm² (Viking missions) to 100W/cm²(Mars pathfinder). The next mission with the MSL is expected to experience a peak heating of 155W/cm². The peak heating

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Altitude

HYPERSONIC ENTRY

SUPERSONIC PARACHUTE

DECENT

SUBSONIC POWERED DECENT

LANDING Entry Interface

Peak Heating Peak Deceleration

Deploy Supersonic Parachute

Heat Shield Separation Deceleration

Backshell Separation

341 sec 8m 309s 800m 247s

~8km

225s 10km 99s

86s 0s

Fig. 0.3: EDL illustration of the MSL mission using a tethered a novel “sky crane” lander technique

phase for the MSL mission is expected to occur at about 86 seconds after the vehicle reaches the entry interface. This peak heating process is then followed by peak deceleration at about 99 seconds into the EDL. The vehicle will then spend about 70 seconds for orientation and decelerate even more to bring the vehicle to the supersonic regime. At 225 seconds a supersonic parachute will be deployed, marking the start of the supersonic decent phase.

Specially designed parachutes are used to slow the vehicle down in the thin CO2 atmosphere. The vehicle will decelerate for 22 more seconds. Once the vehicle is at the appropriate altitude, the lower heat shield is removed exposing the lander/rover vehicle to the Martian atmosphere for the first time. This is expected to occur at about 8km and 247 seconds into the EDL process.

Further deceleration occurs until the appropriate altitude-velocity combi- nation is met to separate the lander/rover vehicle from the backshell at 800m from the surface and 309 seconds after the entry interface. At this point the final decent phase can begin. Earlier missions, such as Viking, incorporated retro rockets for deceleration and crushable legs to absorb most of the energy upon landing. This technique is not favorable for mobile rovers. Mars missions have also seen inflatable airbag approaches (MP) where the vehicle was slowed to nearly a hovering state about 12m above the surface after which the airbags deployed and the vehicle dropped safely to the ground. Future generations of

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landers having larger mass will not be able to utilize this method. The tethered

“sky crane” approach for the upcoming MSL mission is a unique opportunity to evolve techniques for stable landing of larger mass vehicles of larger mass.

The rover is separated from the “sky crane” and slowly lowered via the teth- ers at about 8m. This will “gently” land the vehicle in a near-ready-to-proceed fashion 341 seconds after the entry interface.

All three of these phases have a close symbiosis in the safe landing of a rover on the Martian surface. However, this thesis will focus mainly on the first stage, hypersonic entry. More detailed discussion of this phase and how it can effect subsequent phases following this section. The first point that must be made is that this process is dramatically different for the Martian entry as opposed to Earth entry as shown in figures 0.4 and 0.5. The relative thinness of the Mars atmosphere is readily apparent. At nearly 1/100^ththe density of the Earth atmosphere vehicles tend to go through the hypersonic deceleration phase at much lower altitudes. Even more, this deceleration is not, in many cases, significant enough to lower the vehicle velocity to subsonic levels. Even if the vehicle were decelerated to a subsonic velocity, for the same altitude and same ballistic coefficient the terminal vehicle velocity is about 4 times higher for Mars entry. This problem is further exacerbated because by the time the vehicle is at an acceptable landing velocity the vehicle is so close to the ground that dramatic landing procedures must be incorporated (crushable lander legs, airbags, retrorockets, tethered landing systems,...).

This altitude limiting velocity is so important that, to date, all of the vehicle landing sites have been targeted for less than -1km Mars Orbiter Laser Altimeter (MOLA) determined altitude in order to have sufficient density for EDL operations. The Martian elevation reference for these altitudes is taken to be 6.105mbar which happens correspond to the triple point of H2O.

0 5 10 15 20 25 30 35 40 45 50

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40

Mars Density [kg/m3]

Altitude [km]

Earth

Fig. 0.4:Martian atmosphere density compared to Earth atmosphere.

Non-lifting entry flight paths are described by the ballistic entry equation:

−1 g

du

dt =β⁻¹ρu²

2 (0.1)

whereβ is the ballistic coefficientW/(CDS)which is a very important entry parameter. The lower this value the more the entry vehicle will be slowed as it passes through the atmosphere. An increase in the drag coefficient,CD, or surface area,S(both body shape variables), as well a decrease in vehicle mass throughW will lower the ballistic coefficient value. To what velocity and at

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what rate the vehicle must be slowed to depends on the payload survivability and the vehicle heating rate allowable. If the landing velocity is too great then the lander payload is at risk. If the heating rates are too high or misinterpreted then the entry vehicle could burn up in the entry phase.

Figure 0.5 compares two non-lifting ballistic cases for Earth and Mars entry. This figure illustrates a Mars ballistic entry vehicle with an entry velocity of about 6km/sec and a ballistic coefficient of 100kg/m²compared to that of an Earth ballistic entry vehicle with an entry velocity of about 8km/sec and a ballistic coefficient of 300kg/m². Even with the lighter-weight/higher-drag Mars entry vehicle case, the vehicle still takes quite a long way to slow down.

In fact, only a ballistic coefficient below 50kg/m²can deliver a vehicle to subsonic velocities. But this is at a cost of having to utilize a very small payload combined with a large surface area. The misinterpretation of the ballistic parameter through the unfortunate use of incorrect units is what caused the Mars Climate Observer to fail during an aerobraking orbital insertion process.

Fig. 0.5: Typical ballistic entry for Earth entry and Mars entry, illustrating the dramatic differences in entry trajectories owing to the differences in atmospheric density Figure 0.6 represents the velocity altitude map of past successful Mars lander operations as well as the proposed entry trajectory of the Mars Science Lab- oratory Lander (MSL) which is scheduled for a 2011 launch. The first successful probe/robotic missions were the Viking series of landers whose entry velocity was around 4.7km/s. To achieve this relatively low entry velocity the vehicle was decelerated to an orbital velocity before the EDL phase could begin. This was achieved with retro-rockets, which entail a significant mass penalty during launch. Another concept to insert the vehicle into orbit would be to use aero-breaking where the vehicle skims through a long enough portion of the atmosphere to remove kinetic energy to a point where the orbital velocity is achieved. This technique, which further complicates the entry process, has not yet been demonstrated on larger lander vehicles. Smaller orbiting surveyors, though, have been able to utilize this technique (Mars Global Surveyor and Mars Odyssey). The lighter weight of these vehicles provides the opportunity to utilize the small drag and small control surface forces inherent in the rarefied atmosphere. If a pre-entry orbit can be achieved, which will significantly

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reduce the amount of energy needed to be dissipated during the EDL phase, one should expect about a 3.3 to 5km/s initial entry velocity. But again, this comes at an expensive launch mass penalty. More recent missions have all uti- lized direct entry corridors where velocities are between 5.5 and 7.5km/s. For relatively safe landing velocities the removal of 99.995% to 99.99999% of the intitial entry kinetic energy with respect to the landing site is imperative [2].

This implies that nearly all of the initial 11MJ/kg to 26MJ/kg energy must be dissipated from direct entry missions where initial parking orbits are not uti- lized. Current estimations for large-scale missions involving landing manned, scientific and habitation payloads call for entry masses in the range of 40 to 80 tons and when considering which greatly increases the ballistic coefficient of the entry vehicle making the slowing of the vehicle sufficiently enough quite the challenge.

0 1000 2000 3000 4000 5000 6000 7000 8000

0 10 20 30 40 50 60

9E58E5 7E5 5E5 2E5

4E5 3E5

1E6 6E5

PHX MSL MPF

MER

Altitude [km]

Velocity [m/s]

Viking ^1E5

Re [1/m]

Fig. 0.6:Past successful and future Mars lander entry trajectories

The drastic increase of mass is not the only dramatically changing parameter vexing scientists for future large scale missions. Higher accuracy landing will help to avoid adverse terrain and large scale surface features that could in- hibit the landing system as well as the sensitive equipment and ground sensors on board. Another challenge is to develop EDL systems allowing landings in higher altitude (lower density) environments in proximity to scientifically in- teresting terrain [3]. Table 0.2 is a list of past successful and near future Mars lander missions.

For Earth entry missions it is relatively simple and inexpensive to test ground facility, numerical and flight data as there are many more real flight missions to compare to than for Mars entry. Moreover, Earth atmosphere chemistry is better understood. For Mars entry missions there exist limited flight data, putting a strong emphasis on the accurate modeling and facility testing of such entry systems. Aero-capture and precise control during the hypersonic phase of

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Tab. 0.2:Pertanent data of past successful and future Mars lander missions

EDL Ventry β mentry CD Daeroshell Qpeak L/D

Lander Date [km/s] [kg/m2] [kg] [-] [m] [W/cm2] [-]

Viking 1 1976 4.7 64 992 0.67 3.5 26 0.18

Viking 2 1976 4.7 64 992 0.67 3.5 26 0.18

MPF 1997 7.26 63 584 0.4 2.65 100 0

MER-A 2004 5.4 94 827 0.4 2.65 44 0

MER-B 2004 5.5 94 832 0.48 2.65 44 0

PHX 2008 5.67 70 600 0.67 2.65 58 0.06

MSL 2010 6 115 2800 0.67 4.6 155 0.22

EXOMARS 2015 5.4 681

Future N/A N/A N/A 36000 to N/A N/A N/A N/A

Landers 72000

entry and descent will depend on knowledge atmosphere density and its varia- tions in 20 to 60km altitude range. Current data are not sufficient to meet these operational needs for aero-capture and precision entry/descent [4]. More accurate ground test experiments must be developed to validate numerical models which simulate Mars entry to enable improved Mars entry systems.

European involvement with Mars lander missions will take a significant step when ESA launches the first phase of the Aurora program. The Aurora program has been established by the ESA Directorate of Human Spaceflight, Micro-gravity and Exploration to support future exploration of the Solar Sys- tem via robotic and human means [5]. Aurora seeks to develop a series of robotic missions with a strong technology development content to act as build- ing blocks that will support human space exploration. The program plans long term technology developments to support missions to Mars, the Moon and possibly other near-Earth objects.

European Space Agency’s next Mars lander mission will be Exo-Mars. The Exo-Mars mission has similar goals compared to other Mars missions. Once landed, the surface rover science platform, comparable in size to the two NASA exploration rovers, Spirit and Opportunity, will help to search for past and present evidence of life, to understand the near surface geochemical environment and to better characterize the atmosphere. These goals, if met, will greatly aid future manned missions. Exo-Mars is currently scheduled for launch in 2013, with a possible back-up launch date foreseen in 2015 [6].

The engineering challenges for Exo-Mars mission is to develop a rover (as seen in figure 0.7) with a mobility of at least several kilometers. Relatively deep drilling capabilities are also on the design slate giving the rover access to soil samples up to 2 meters below the surface. These capabilities will be coupled with a wide array of automatic components for sample preparation and testing with the on-board scientific instrumentation. Characterization of the water/geochemical environment in the shallow 2m subsurface is also of high interest. This mission will also study surface terrain for hazards potential for future missions as only an extremely small area of the planet’s surface has been adequately surveyed for landing opportunities [5].

In conclusion, understanding the physics of the initial entry phase is imperative for the delivery of a landing system to a safe velocity at the lower altitude descent and landing phases of the EDL process. For any entry problem, the chemistry and the subsequently effected heating and aerodynamic character-

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Fig. 0.7: Artist rendering of the Exo-Mars lander as it rolls away from the base platform atop its deflated airbag system [5]

istics can significantly influence the vehicle entry performance. The thin Mars atmosphere creates an environment that requires a more rapid hypersonic deceleration process than that of Earth entry compounding the problem. This is because the supersonic deceleration and subsonic landing phases that occur at much lower altitudes leaving significant requirements for high drag vehicles to reduce the velocity to acceptable levels to avoid damage or mission failure during landing. Thus, sufficiently understanding the initial phase of the EDL, the entry problem, where most of the energy is to be dissipated, is pivotal for safe decent and landing operations for future missions. This is especially necessary if the future calls for increases in landing mass, improvements in landing accuracy and increases in landing elevation are to be met.

The emphasis of this thesis is to help to develop more accurate tools to define thermodynamic conditions in these ground test facilities. This will help to more accurately monitor ground testing facilities thereby improving the numerical accuracy and confidence in CFD simulations.

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1. INTRODUCTION

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1.1 The Hypersonic Regime

H

^YPERSONICaerodynamics is generally classified as a flow regime where the Mach number exceeds 5. This is, of course, a rough generality as the true nature of hypersonics is a flow regime where the enthalpy is significant enough that inviscid, perfect gas aerodynamics fail to describe the nature of the flow. Hypersonics is, perhaps, better described as the flow regime where non-equilibrium thermo-chemical effects and real gas phenomena require at- tention. Implications of these effects are discussed in the next section. Thus, a better definition of the hypersonic regime is a regime characterized where one or more of the effects described in the next subsection have an influence that can not be neglected. The understanding of this flight regime is based upon theory reinforced by limited testing and computational simulation abilities.

Flight tests, which can provide useful information about these phenomena, are, unfortunately, prohibitively expensive. Data from flight tests have come primarily from entry data from vehicles like Shuttle and the Mars landers. There are only very limited data from these tests due to infrequent operation stemming from the tremendously expensive nature of such flights. These missions usually serve an alternate purpose than the study of hypersonic aerodynamics which leads to a sharing of a large portion of the cost to other scientific endeavors. However, there are stand alone hypersonic test beds- some that have been flown and others in development. NASA’s Hypersonic Bound- ary Layer Transition (HyBoLT) rocket’s mission is to obtain unique high-speed flight data for fundamental boundary layer transition flow physics. This vehicle unfortunately failed on launch. The NASA X-43 program is an air-breathing self-propelled hypersonic vehicle technology demonstrator. Tests have been few with the last occurring in 2004. The vehicle has proven many technolo- gies including SCRAMjet propulsion, materials able to withstand significant thermal loads, vehicle control and stability, etc. The HyShot and HyCause pro- grams from the University of Queensland have been doing SCRAMjet propulsion flight test research since 2001. The test vehicles are launched in the Aus- tralian desert atop a two-stage Orion booster which accelerates the vehicle to the edge of the atmosphere. After a few orientation maneuvers the vehicle bal- listically accelerates toward the surface. At the point where significant Mach number is achieved, the engines are started and fired until a peak Mach number is reached (between 7 to 10). The data on board the vehicle are relayed to several monitoring stations. The EXPerimental Re-entry Testbed (EXPERT) is an ESA project with the objective to study, in flight, the phenomenon of transition, catalicity and oxidation, real gas effects on shock-wave boundary layer in- teractions, microaerothermodynamics and blackout. It is a highly collaborative effort between several research institutes including the Von Karman Institute.

The scheduled test flight of this vehicle is currently set for 2010.

Ground testing in hypersonic or sub-sonic high-enthalpy (plasma based and non-plasma based) facilities are significantly cheaper. However, these facilities operate in a narrow window and no one facility is capable of duplicating the full flight environment. Thus, a ground test campaign requires many tests from various facilities to accurately define the flight envelope or flight path that an actual flight test could describe. Ground testing can be used on flight vehicle test models to test vehicle performance or specific hypersonic phenomena (once the facility is calibrated). Numerical validation is another use of these

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facilities which can be done with or without a test model installed provided that the same real gas effects can be generated in the facility as they occur in reality during entry or flight. Numerical code can be calibrated with these facilities as they are very capable of providing the non-thermal equilibrium and non-chemical equilibrium effects inherent in hypersonic flight. More specific details about various ground facility types with their operation principles will be discussed in section 1.2.

Numerics are normally the cheapest and easiest option to incorporate yielding the capability to map the entire entry corridor if the physics are modeled appropriately. However, appropriately modeling the physics is quite a challenge as the thermal non-equilibrium and chemical non-equilibrium effects inherent of hypersonic flows can make even already complex problems, such as viscous laminar to turbulent transition simulations, much more complex. The tools must be validated with physical data. Flight tests are generally limited but provide real flight information. Ground tests are the typical validation tool as they can be implemented more easily and controlled to specific phenomena of interest more accurately. Since these ground-test and flight test results are limited in the hypersonic regime, CFD tends to have a greater impact in the hypersonic regime then in the subsonic or supersonic ones. Thus, CFD, if cor- rectly modeled, is an extremely powerful research, design and development tool.

There obviously is a strong symbiosis between flight testing, ground testing and numerical simulation of hypersonics that cannot be ignored. Figure 1.1 illustrates that these three fields are needed to complete the entire picture. CFD alone is not advanced enough to model every mission accurately and ground tests alone cannot map the entire flight corridor. The future success of understanding the hypersonic regime and the success of real flight vehicles depends heavily on an iterative evolution of flight tests, numerical tool development and ground test validation and analysis. This report will focus on the ground test portion with results aimed at aiding future facility improvement and numerical tool development.

1.1.1 Characteristics of Hypersonic Flows

There are several important physical phenomenon that are associated with hypersonic aerodynamics. These are as follows:

Thin Shock Layer

The area between the shock and the vehicle of a super- or hypersonic vehicle is known as the shock layer. As Mach number increases, this shock layer region becomes thinner as the shock angle decreases. High-temperature effects such as chemical reactions will reduce this shock angle further. This thin shock layer for some vehicle shapes can merge with the thick developing hypersonic boundary layer. This phenomenon can be taken advantage of leading to the simple and straightforward thin shock layer theory.

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GROUND TEST FLIGHT

CFD

Fig. 1.1: CFD, ground tests and flight data all rely on one another and are all needed to complete the entire picture of hypersonic flight

Substantial Entropy Layer

Consider a blunt body vehicle traveling at hypersonic velocity. A substantial bow shock will form ahead of the body compressing and heating the flow.

Parallel pre-shock streamlines pass through the curved bow shock and are no longer parallel. This creates vorticity in the region behind curved portions of the shock. It is well known, according to the theorem of Crocco (equation 1.1), that the presence of vorticity is associated with the existence of entropy.

~

u×(∇ ×~u) =∇h0−T∇s (1.1)

Total enthalpy across a shock, curvature or not, does not vary(∇h0 = 0). Thus, if vorticity is created as the flow crosses a curved shock(~u×(∇ ×~u)6= 0), then there exists an entropy gradient. Near the stagnation streamline there is a higher curvature gradient creating a strong entropy zone in the stagnation region of the vehicle. This entropy-laden flow travels downstream following the streamlines close to the vehicles surface interacting with the boundary layer.

This interaction can lead to difficulties in analyzing the boundary layer along the vehicle properly .

Viscous Interaction

Traveling at high velocities means that there is a large amount of kinetic energy for the flow with respect to the vehicle. This flow near the surface will be slowed by friction, converting the kinetic energy into thermal energy, thereby increasing the boundary layer temperature. This effect is known as viscous dissipation. Owing to the temperature increase toward the surface the density must decrease to satisfy mass conservation yielding a thicker boundary layer.

It is this viscous dissipation effect that causes hypersonic boundary layers to

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grow so rapidly. This thick boundary layer will displace the inviscid portion of the flow which will effectively presents a larger than actual vehicle body.

These afore mentioned phenomena characteristic to hypersonic flows will greatly effect skin friction, heat transfer, and pressure distribution over a body.

High-Temperature Effects

Hypersonic flow can also invoke high temperature or real-gas effects. Kinetic energy transformed to thermal energy will first excite vibrational modes of the molecules. This effect is actually caloric imperfection where specific heats now depended on temperature since energy storage depends on temperature. At higher velocities dissociation occurs with a composition dependence on temperature and pressure which can in some portions of the entry contain ionized species. This ionization is the cause of the communications blackout portion of the Shuttle entry. At this point specific heats must be treated for a mixture of thermally and calorically imperfect gases. Further details of chemical non- equilibrium and real gas effects and as they apply to ground testing will be discussed later.

1.1.2 Gas Classifications Owing to High Temperature and High Pressure Effects

These special hypersonic phenomenon lead to very unique gas modeling conditions. For clarification, some of these classifications are described here.

Real Gas

A real gas is a gas where all of the following items can be a factor:

• Compressibility effects

• Variable heat capacity

• Close interaction forces

• Non-equilibrium thermodynamic effects

• Molecular dissociation and elementary reaction with variable composition

This is the most complete model but it is not always necessary needed to describe the gas. The following entries are some common classifications that will be mentioned upon throughout this thesis.

Thermally Perfect Gas

For a thermally perfect gas, the enthalpy and entropy are functions of temperature alone.

e=e(T)

h=h(T) (1.2)

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If the gas is far from condensation and the intermolecular force details are not important then the equation of state takes the following familiar form:

P =nikT (1.3)

Calorically Perfect Gas

A calorically perfect gas is defined as a gas in which the specific heats do not depend on temperature, such that enthalpy and eternal energy relations can be described as:

dh=CpdT de=C_vdT e=CvT

h=C_pT (1.4)

For this gas classification the perfect gas equation of state is still valid even at moderate supersonic speeds.

Thermally Perfect but Calorically Imperfect Gas

As the gas is heated further, vibrational energy modes of gas molecules are significant and molecular specific heats are no longer constant but now depend on temperature. As long as the particular molecular specie is far from condensation it can be classified as calorically imperfect but thermally perfect. The enthalpy,h, must now take into account the thermal energy stored in vibration.

dh=cp(T)dT de=c_v(T)dT e=cv(T)T

h=c_p(T)T (1.5)

Equation 1.6 is a common model of the vibrational energy storage for the specific heat for linear triatomic molecules such as CO2. Temperature dependent specific heats need to be considered when the temperature approaches the characteristic vibrational temperature,θv.

c_p R =7

2 +

θ_v/2T sinh(θv/2T)

2

(1.6)

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Chemically Reacting Gas

Increasing the temperature further creates even stronger vibrational excitation effects to where the gas constituent molecules begin to break their molecular bonds causing dissociation and eventually ionization. Here, individual species need to be considered such that:

h=h(T, N1, N2, N3, . . . , Ni) e=e(T, N₁, N₂, N₃, . . . , N_i) cp=f1(T, N1, N2, N3, . . . , Ni)

c_v=f₂(T, N₁, N₂, N₃, . . . , N_i) (1.7) whereNiis the number density of specie “i”. For the equilibrium case where the chemical composition remains constant over time the number density values are a function of the temperature and pressure which leads to:

h=h(T, ρ) e=e(T, ρ) cp=f1(T, ρ)

cv=f2(T, ρ) (1.8)

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1.2 Hypervelocity Facilities for Ground Testing Hypersonic Aerodynamics

As this report is primarily a ground test investigation, the physics and tools of ground testing are discussed in more detail here. The single goal of a hypersonic ground test facility can be summarized in one simple sentence: hypervelocity facilities require the efficient production of high enthalpy flow [8]. This is much easier said than done. In true hypersonic flight the vehicle is traveling at high velocity in a rested flow which is in a low temperature equilibrium state. The afore-mentioned hypersonic effects only become relevant in flight after the body heats and pressurizes the flow in a strong shock compression process. However, in a non-ballistic-type ground test facility, the test body or test article is at rest while the flow is energized to match conditions such that the tests can be scaled to the flight environment. There are many types of facilities to simulate the afore-mentioned characteristic phenomena associated with hypersonics, each with some variation in their method of generating the necessary energy and conditions for “hypervelocity” testing. The term “hypervelocity” is commonly used over hypersonic as the key feature is more the velocity rather than the Mach number as this gives a firmer indication of the kinetic energy involved. Each method of creating simulation conditions has its own problems. This will be illustrated in the following subsections with corresponding descriptions of various hypersonic/hypervelocity test facilities.

1.2.1 Hypersonic Generalities of Ground Test Facilities

A good starting point to get a general grasp of hypervelocity facility operation and where many problems of their operation arise, is the energy equation for inviscid compressible high velocity flow:

h0=e0+P0

ρ₀ =

e+P ρ +u²

2

∞

≈u²_∞

2 (1.9)

High enthalpy hypervelocity facilities (continuous tunnels, arc-heated plasma tunnels, impulse tunnels) require energy addition to an initial stagnation condition to achieve high-enthalpy. One could focus on increasing the stagnation internal energy,e₀, as in arc-heated facilities, detonation facilities and combus- tion tunnels. Another approach could be to increase the stagnation pressure to create the high-enthalpy flow. This compression process must take care that the state of the gas is well away from saturation/condensation. Normally this problem can be circumvented by pre-heating the test gas. Once the stagnation conditions are met, then the high-enthalpy state can be accelerated to high kinetic energy as illustrated in figure 1.2.

These high enthalpy “stagnation” conditions require energy addition which can be significant enough to produce vibrational excitation and dissociation.

The free-stream conditions are less prone to have vibrational modes of energy stored or to have chemical dissociation in an equilibrium state. However, the relaxation of these stagnation condtions do not happen instantaneously as the flow progresses to the test article. If the characteristic flow time,τf low, is much larger than the vibrational relaxation time,τvib, and the chemical reaction characteristic time,τchem,(τf low τvib, τchem)then the flow can be said to behave

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“Stagnation” ?

Conditions

Freestream Hypervelocity Flow

2

2 0 0 0 0

V e +P +

ρ

2 2

∞

∞+P∞+V ≈V e ρ

≈0

Fig. 1.2: Typical hypervelocity acceleration process through a nozzle. A high-energy

”‘stagnant”’ condition (state 1) is accelerated by a nozzle (state 2) into a hypersonic free-stream flow (state 3)

in localequilibriummanner. On the other extreme side, if the characterisitc flow time is much smaller than the vibrational relaxation time or the chemical reaction time(τ_{f low}τ_vib, τ_chem)the flow is said to be in afrozenstate. A frozen state does not imply equilibrium, rather it is just an indication that the chemical reaction time is too slow to react during the characteristic flow duration. Any- thing in between these two extremities is considered to be innon-equilibrium.

This grey area between the two is normally the case for high-enthalpy facilities.

A favorable condition of experimental flight simulation of hypervelocity flows is to have the gas arriving at the test section be in local thermodynamic and chemical equilibrium as in flight but this becomes more difficult as tunnel enthalpy requirements increase. Sometimes, though, these non-equilibrium effects can be studied and if measured properly provide value information for the validation of numerical tools.

Because of these effects, accurate determination of flow conditions in either the free stream or near a model is quite difficult. Tunnel free stream conditions in perfect gas hypervelocity facilities can be estimated using isentropic nozzle flow relations. However, owing to real gas effects, these calculations made with plenum conditions and isentropic nozzle equations can be erro- neous when surpassing Mach numbers greater than 10 especially for carbon dioxide [19].

Development of instrumentation to make accurate and reliable measure- ments in the typical harsh nature of hypervelocity facilities has been challeng- ing at best. A typical approach uses pitot pressure and hemisphere stagnation point heat flux and pressure data acquired from intrusive probing. This information is then used to rebuild the flow field with some model. This is not a direct straight forward method data reduction method. The non-equilibrium conditions will also complicate the data reduction process as simple models de- scribing the flow properties, while useful in low-enthalpy environments, begin to lose their fidelity at higher enthalpies.

As stated earlier, many different types of facilities are needed to map a specific entry corridor as each facility can only partially simulate the large varia- tions in high-enthalpy conditions. These limited results, along with flight test data, are used for validation of numerical simulation which fills gaps to provide an estimate of the entry environment. The following two sub-sections describe some of the most common ground test facilities used in hypersonic

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testing to do just that. Their advantages and disadvantages will also be briefly touched upon.

1.2.2 Common Plasma Facility Types

Plasma facilities provide reasonable replication of the intense thermal environments experienced in hypersonic flight. There exist two types of these facilities;

arc-heated wind tunnels and inductively coupled plasma (ICP) tunnels. These facilities normally run for long periods and the test duration is often limited by the sample/test article survivability.

Arc-heated Tunnels

Arc-heated facilities were first developed in the late 1950s when increased space access activities necessitated improved understanding of entry physics. An ex- cellent review of the two main types of arc heaters, the Huels and segemtented type, can be found in reference [10].

The most simple type of the arc-heater is known as a “vortex-stabilized”

Huels type of heater. The simplicity of design and relative ease of operation allow for easy maintenence and shor turn around times during testing. This facility type consists of a gas swirl arc chamber embedded in a water cooling circuit. A convergent-divergent nozzle section is attached at the end to convert a portion of the thermal energy into kinetic energy. Huels type of arc heaters can run at relatively high pressure but at a lower enthalpy as the current density must be relatively low to maintain integrity of electrodes. Also an issue is that the arc discharge has little no constraint on termination point and can lead to lack of repeatability.

Current arc-heated facilities provide stable discharge by using a segmented electrode and constriction design. This facility is known as a segmented arc- heater. Current is distributed over a stack of electrode disks which produces a lower thermal load to the individual electrodes. This leads to only a moderate improvement in contamination than the Huels type but, more importantly, allows control of the arcing locations and thus arc lengths. This increase in arc length control in turn increases the stability and repeatability of testing.

Higher enthalpy levels can be achieve as high current arcs can be used since the arc elctrode heating is spread over many electrodes. However, design complexities limit operating pressures to generally lower than that of Huels-type heaters. Segmented heaters also require more frequent inspection and maintenence than Huels-type.

Inductively Coupled Plasma (ICP) Facilities

One approach to avoid contamination from erosion of cathode and anode ma- terial is to change the method of adding energy to the flow such as an inductively coupled plasma (ICP). An ICP facility consists of a large current power supply connected to a coil surrounding a quartz tube. A large current flows through the coil creating a strong electromagnetic field. This electromagnetic field ionizes the gas and the electrons impart considerable energy to the gas by collisions. ICP facilities typically operate in the subsonic (some times sonic!)

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regime but they are capable of creating the high thermal and chemistry effects inherent in hypersonic flight.

1.2.3 Common Impulse Hypervelocity Facility Types Shock Tube

Another common approach for increasing the enthalpy of a gas for hypersonic testing is via a shock-tube. Such devices consist of two chambers containing two gases at different pressures separated by a diaphragm. The bursting of this diaphragm is usually done in a pseudo-controlled method via a double diaphragm. The volume between the double diaphragm contains an intermediate pressure. This intermediate pressure is set to where the pressure ratio between the higher pressure driver and diaphragm volume pressure as well as the pressure ratio between the diaphragm volume pressure and the lower pressure driven portion is not enough to rupture the diaphragm.

Max Test Time

Time

Distance

Driver Gas (4) Driven/Test Gas (1)

Incident Shock Reflec

ted Shock

Driver/Driv en Int

erface Leading Expansion

Wave

Diaphragm

2

4

Flow

1

Fig. 1.3:Standard shock-tube

To start the flow, the intermediate pressure is evacuated, which increases the pressure ratio and causes the diaphragm(s) to rupture. Shock heating is a well-proven technique to add significant enthalpy to the flow. The shock tube can simulate real gas effects satisfactorily but they have limits on Mach number and model sizes allowable. Run times are also extremely small, limited by the time it takes the contact surface to arrive at the test section after the initial shock. The description of a shock tube here is needed as all impulse facilities are based, to some extent, on its operation. Consider equation 1.10 which is a relation for shock tube parameters as illustrated in figure 1.3.

P₄ P1

= 2γ₁M_S²−(γ₁−1) γ1+ 1

1−(γ₄−1)a₁ (γ1−1)a4

M_S− 1 MS

_γ^2γ₄₋₁⁴

(1.10) This equation shows that the shock Mach number, MS, can be increased by not only the driver to driven gas pressure ratio,P4/P1, but also by playing

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with the driver and driven gas constituents through the specific heat ratios, γi, and local sound speeds, ai. However, there are implications to altering these parameters which must be considered especially if one wants to simulate chemistry properly. Typical facilities may run helium or hydrogen as driver gases generating only modest enthalpy levels.

The flow behind the incident shock, region 2, is at a relatively high enthalpy state. Recall that vibrational excitation depends on the temperature and begins at around 600K for N2and O2. Using the normal shock relation in equation 1.11 one can see that this corresponds to a modest Mach number of 2.4. These phenomenon will absorb energy from the flow, thereby decreasing the shock layer thickness. The ability to control these high temperature effects in a simple facility makes the shock-tube facility attractive for experiments aimed at numerical validation of vibrational and chemical kinetics.

T2

T₁ =

2γ1M₁²−(γ−1) (γ−1)M₁²+ 2

(γ+ 1)²M₁² (1.11)

Expansion Shock Tube

This type of facility is basically an extension of the shock tube design with a third tube mounted downstream. This third tube is called the acceleration tube and is maintained initially at low pressure. When the downstream traveling shock ruptures the diaphragm to the acceleration tube, an unsteady expansion process will convert the shocked thermal energy into kinetic energy thus increasing the flow velocity. This addition helps to overcome the Mach number limitation of shock tubes. An acceleration tube process is generally more efficient than a nozzle at converting the thermal energy into kinetic energy but, the size of the tube diameter severely limits the test article size. Moreover, these expansion tube facilities have very short run times. These are similar limitations shared with shock tube facilities.

Time

Distance Driver Gas (4) Intermediate Tube (1)

Incident Shock Driv

er/Int ermediat

e

Tube Int erface LEW

Diaphragm Diaphragm

1 4

Flow

Acceleration Tube (10)

Diaphragm 10

Test Inter./ Acc. Tube Interface Time

2 LEW

LEW = Leading Expansion Wave

Fig. 1.4:Standard expansion type shock-tube

Tunable Diode Laser Absorption Spectroscopy Characterization of Impulse Hypervelocity CO

Tunable Diode Laser Absorption

Spectroscopy Characterization of Impulse Hypervelocity CO 2 Flows

A Thesis submitted for the degree of

DOCTORAT EN SCIENCES DE L’INGENIEUR Jason Matthew Meyers

UNIVERSITE LIBRE DE BRUXELLES F

´ S

A

´

VON KARMAN INSITUTE FOR FLUID DYNAMICS A

A

D

Tunable Diode Laser Absorption

Spectroscopy Characterization of Impulse Hypervelocity CO 2 Flows

A Thesis submitted for the degree of DOCTORAT EN SCIENCES DE L’INGENIEUR

Jason Matthew Meyers

UNIVERSITE LIBRE DE BRUXELLES F

´ S

A

´

VON KARMAN INSITUTE FOR FLUID DYNAMICS A

A

D

September 9, 2009

SUMMARY

ACKNOWLEDGMENTS

CONTENTS

NOMENCLATURE

Latin Letters

Greek Letters

Subscripts

PREFACE

M

1. INTRODUCTION

1.1 The Hypersonic Regime

H

1.2 Hypervelocity Facilities for Ground Testing Hypersonic Aerodynamics