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Thèse de doctorat/ PhD Thesis Citation APA:

Meyers, J. (2009). Tunable diode laser absorption spectroscopy characterization of impulse hypervelocity CO2 flows (Unpublished doctoral dissertation).

Université libre de Bruxelles, Faculté des sciences appliquées – Mécanique, Bruxelles.

Disponible à / Available at permalink : https://dipot.ulb.ac.be/dspace/bitstream/2013/210279/4/0f7a7d95-0dc3-46a4-8585-986a87450233.txt

(English version below)

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D 03673

Tunable Diode Laser Absorption Spectroscopy Characterization of

impuise Hyperveiocity CO ² Flows

Jason Matthew Meyers

Université Libre de Bruxelles

Faculty of Applied Sciences ULB

Von Karman Institute for Fluid Dynamics Aeronautics and Aerospace Department

Supervisor: D. G. Fletcher (VKI) Promoter: G. Degrez (ULB)

Committee:

F. Dubois (ULB) M. Herman (ULB) A. K. Mohamed (ONERA)

U. Koch (DLR)

Université Libre de Bruxelles

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Nomenclature

Latin Letters

K A

<T Of a a B c Cp a D E e E' E"

F F f f Fv

9

Gv h h I I lo Enc

Erans

Force constant Wavelength

Ro-vibrational term value Half-cone angle of focus Voigt "a" parameter

Wavelength modulation amplitude Rotational constant

Speed of light

Spécifie heat capacity at constant pres

sure

Spécifie heat capacity at constant vol

ume

Correction term for non-rigid rotation Energy

Mass spécifie internai energy Upper-state State energy Lower-state energy Etalon finesse Rotational term Focal length

Force from Hooke's law model Rotational term value

Gravitaional force

Lower State degeneracy of transition i Statistical weight factor for population distribution of rotational levels J Vibrational term value

Mass spécifie enthalpy Planck's constant

Harmonies of the measurement Moment of inertia

Transmitted intensity Reference intensity Etalon incident intensity Etalon transmitted intensity

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J k L

Lres

M rUe n n n Nj riL NA NIR

P

Qelect

Qelect Qint

Qrot

Qvib

' eq

S s S*

T To

Petalon

Tref

U

Ugas V vuv

w Wo

Wf

X y

Rotational quantum number Boltzman constant

Pathlength

Length of resonator Mach number électron mass

Collisional broadening exponent Index of refraction

Number density

Population distribution of rotational levels J

Loschmidt number Numerical aperture Near-infrared Pressure

Electronic internai partition fonction (partition sum)

Electron charge

Total internai partition fonction (parti

tion sum)

Rotational internai partition fonction (partition sum)

Vibrational partition frmction (partition sum)

Displacement from Hooke's law model Equilibrium displacement from Hooke's law model

Linestrength [cm~^/ofm]

Entropy

Linestrength \cm~^/molécule ■ cm~^]

température

Reference température Transmitted light fraction

Reference température from HITRAN database

Velocity Gas velocity Voigt fimction Vacuum ultraviolet Voigt "w" parameter Beam Diameter (l/e) Focal waist diameter Mole fraction

Intégral variable of Voigt fonction

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Nomencla ture FWHM

HWHM

Greek Letters

a 0 Au

ô Auc Aud

Aufsr

7 A A

U

Vo

^08C ÜJ UJm

<t>

0c 0D 4>v P

T

6 e' 6v ù d tau

Subscripts

0 00

1 i j k S

Full-width half maximum Half-width half maximum

Absorption coefficient Ballistic coefficient

Spacing behveen resonator modes frequency shift due to Doppler shift Phase shift per traversai

Collisional broadening FWHM value Doppler broadening FWHM value Etalon free spectral range

Spécifie beat ratio Mean free path Wavelength Reduced mass Frequency

Line center frequency

Oscilla ting frequency of a spring model Angular velocity

Modulation frequency Broadening coefficient

Collisional broadening coefficient Doppler broadening coefficient Voigt broadening coefficient Density

Spectral transmittance Angle of beam w.r.t. flow Internai angle within étalon

Characteristic vibrational température Wavenumber

Etalon spacing or thickness Transmission fraction

Stagnation conditions Free stream conditions Line transition index Lower State

Specie index Upper State Shock

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Nomencla ture

PREFACE

M

^arti AN atmosphère entry (illustrated in figure 0.1) is a complicated lo- gistics problem fraught with engineering complexifies. A vehicle is laimched 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.

Figure 0.1: Artist's rendering of Mars entry.

The EDL complexifies begin even before a significant atmosphère is présent.

It is desired that the entry vehicle high température TPS be oriented properly such that the heat shield will be in the région of the high température bow shock so as not to compromise the after body TPS at or near continmun con

ditions before the high température effects begin. Initial maneuvering takes place in the rarefied free molecular régime where the mean free path of molecular collisions is large compared to that of the characteristic body length leading to the familiar Navier-Stokes équations becoming incapable of representing flow physics. Because of this significantly rarefied atmosphère, entry vehicles are statically unstable for the free-molecular and much of the transitional flow régime before the continuum assumptions can be made. It is therefore difficult to control a vehicle, especially one without adéquate 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 g5rroscopic spin. The aerodynamics and atmosphère 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 régime bas been done by Moss et al. [1]. As the atmospheric density increases, the transitional phase begins imtil 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 at

mosphère 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 désirable decent velocity.

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

The extreme heating rates involved call for considérable 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 région, plays a huge rôle in vehicle stability. These non-equilibrium effects can significantly alter pressure distribution and heating rates. Thus, considération from stagnation région 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 pivotai for the safe landing of such payloads as the Mars Science Laboratory Rover and the Mars Exploration Rovers illustrated in figure 0.2

Figure 0.2: Artist rendering of the Mars Science Laboratory Rover and one of the Mars Exploration Rovers illustrating the size comparison of fu

ture génération vehicles (taken from solarsystem.nasa.gov

Consider figure 0.3 as an example which illustrâtes the final EDL key phases of the spécifie case representing the Mars Science Laboratory (MSL) mission where a tethered "sky crâne" approach will be used to land the most massive vehicle ever to the Martian surface. This case is presented as it is, currently, the

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Nomencla ture next significant Mars naission scheduled to launch in the 2011 window.

■S

s

’-Æ

<

® Entry Interface

^ Peak Heating

^ 99s. Peak Décélération

225s lOkm

HYPERSONIC ENTRY

Deploy Supersonic

Décélération

Heat Shield Séparation

Backshell Séparation 800m

SUBSONIC POWERED DECENT

Figure 0.3: EDL illustration of the MSL mission using a téthered a novel "sky crâne" lander technique

The first phase, hypersonic entry, begins when the entry vehicle first cornes into contact with a sensible atmosphère carrying significant kinetic energy. The tiemendous heating during this phase can be an)rwhere between 25W/cm^

(Viking missions) to lOOW/cm^ (Mars pathfinder). The next mission with the MSL is expected to expérience a peak heating of 155W/cm^. The peak heating 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 décélération 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 régime. At 225 seconds a supersonic parachute will be deployed, marking the start of the supersonic decent phase.

Specially design parachutes are used to slow the vehicle down in the thin CO2 atmosphère. The vehicle will decelerate for 22 more seconds. Once the

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vehicle is at the appropriate altitude, the lower heat shield is removed exposing the lander/rover vehicle to the Martian atmosphère for the first time. This is expected to occur at about 8km and 247 seconds into the EDL process.

Further décélération occurs until the appropriate altitude-velocity combina

tion is met to sépara te 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 rétro rockets for décélération and crushable legs to absorb most of the energy upon landing. This technique is not favorable for mobile rovers. Mars missions hâve 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 générations of landers having larger mass will not be able to utilize this method. The tethered

"sky crâne" approach for the upcommg 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 crâne" 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.

Ail three of these phases hâve 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 subséquent 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 thirmess of the Mars atmosphère is readily apparent. At nearly 1/100‘^ the density of the Earth atmosphère vehicles tend to go through the hypersonic décélération phase at much lower altitudes. Even more, this décélération 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 groimd 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, ail of the vehicle landing sites hâve been targeted for less than -1km Mars Orbiter Laser Altime- ter (MOLA) determined altitude in order to hâve sufficient density for EDL operations. The Martian élévation reference for these altitudes is taken to be 6.105mbar which happens correspond to the triple point of H2O.

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

Idu _ 1 pv?

~Y (0.1)

where /? is the ballistic coefficient WI{CdS) which is a very important entry parameter. The lower this value the more the entry vehicle will be slowed as

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Nomencla ture

Figure 0.4: Martian atmosphère density compared to Earth atmosphère.

it passes through the atmosphère. An increase in the drag coefficient, Cd, or surface area, S (both body shape variables), as well a decrease in vehicle mass through W will lower the ballistic coefficient value. To what velocity and at what rate the vehicle must be slowed to dépends 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 bum up in the entry phase.

Figure 0.5 compares two non-lifting ballistic cases for Earth and Mars en

try. This figure illustrâtes a Mars ballishc entry vehicle with an entry velocity of about 6km/ sec and a ballistic coefficient of lOOkg/m^ compared to that of an Earth ballistic entry vehicle with an entry velocity of about 8km/sec and a ballistic coefficient of 300kg/m^. 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/m^ can 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.

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 Labo- ratory Lander (MSL) which is scheduled for a 2011 launch. The first successful probe/robotic missions were the Viking sériés 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 signihcant 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 atmosphère 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, hâve been able to utilize this technique (Mars Global Surveyor and

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All.(Km) CompanngMars and Earth EDL

Figure 0.5: Typical ballistic entry for Earth entry and Mars entry, illustrating the dramatic différences in entry trajectories owing to the différ

ences in atmospheric density

Mars Odyssey). The lighter weight of these vehicles provides the opportunity to utilize the small drag and small control surface forces inhérent in the rarefied atmosphère. If a pre-entry orbit can be achieved, which will significantly 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 cornes at an expensive launch mass penalty. More recent missions hâve ail 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 impérative [2].

This implies that nearly ail of the initial llMJ/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.

The drastic increase of mass is not the only dramatically changing parameter vexing scientists for future large scale missions. Fligher 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) envirorunents 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

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Nomencla ture

Figure 0.6: Past successful and future Mars lander entry trajectories

facility, numerical and flight data as there are many more real flight missions to compare to than for Mars entry. Moreover, Earth atmosphère 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 en

try Systems. Aero-capture and précisé control during the hypersonic phase of entry and descent will dépend on knowledge atmosphère 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 précision entry/descent [4]. More accu

rate grormd 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 System via robotic and human means [5]. Aurora seeks to develop a sériés of robotic mis

sions with a strong technology development content to act as building blocks that will support human space exploration. The program plans long term tech

nology 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

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Table 0.2: Pertanent data of past successful and future Mars lander missions Lander

EDL

Date ^ entry [km/s] P

[kg/m2] [kgl Cd

[-1

^aeroshell

[m] Qpeafc [W/cm2]

L/D [-1

VUking 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 Landers

N/A N/A N/A 36000 to 72000

N/A N/A N/A N/A

landed, the surface rover science platform, comparable in size to the two NASA exploration rovers, Spirit and Opportrmity, will help to search for past and présent evidence of life, to understand the near surface geochemical environ

ment and to better characterize the atmosphère. 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].

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

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 préparation and testing with the on-board scientific instrumentation. Characterization of the water/geochemical envirorunent 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

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Nomenclature been adequately surveyed for landing opportunities [5].

In conclusion, understanding the physics of the initial entry phase is impér

ative 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 characteristics can significantly influence the vehicle entry performance. The thin Mars atmosphère créâtes an environment that requires a more rapid hypersonic dé

célération process than that of Earth entry compoimding the problem. This is because the supersonic décélération 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 pivotai for safe decent and landing operations for future missiorrs. This is especially necessary if the future calls for increases in landing mass, improvements in landing accuracy and increases in landing élévation 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 Régime

H

^ypersonic aerodynamics is generally classified as a flow régime where the Mach number exceeds 5. This is, of course, a rough gerrerality as the true nature of hypersonics is a flow régime where the enthalpy is significant enough that inviscid, perfect gas aerodynamics fail to describe the nature of the flow. Hj^ersonics is, perhaps, better described as the flow régime 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 définition of the hypersonic régime is a régime characterized where one or more of the effects described in the next subsection hâve an influence that can not be neglected. The imderstanding of this flight régime is based upon theory reinforced by limited testmg and computational simulation abilities.

Flight tests, which can provide useful information about these phenomena, are, unfortunately, prohibitively expensive. Data from flight tests hâve corne 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 mis

sions usually serve an altemate purpose than the study of hypersonic aero

dynamics 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 hâve been flown and others in development. NASA's Hypersonic Boimd- ary Layer Transition (HyBoLT) rocket's mission is to obtain imique high-speed flight data for fondamental boundary layer transition flow physics. This vehicle imfortunately failed on larmch. The NASA X-43 program is an air-breathing self-propelled h5q?ersonic vehicle technology demorrstrator. Tests hâve been few with the last occurring in 2004. The vehicle has proven many technolo

gies mcluding 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 hâve been doing SCRAMjet propul

sion 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 atmosphère. 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 rmtil a peak Mach num

ber 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 transi

tion, catalicity and oxidation, real gas effects on shock-wave boimdary layer in

teractions, microaerothermodynamics and blackout. It is a highly collaborahve effort between several research institutes including the Von Karman Institute.

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

Groimd testing m 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

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1.1 The Hypersonic Régime full flight environment. Thus, a ground test campaign requires many tests from varions 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 spécifie hypersonic phenomena (once the facility is calibrated). Numerical validation is another use of these 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 fa

cilities as they are very capable of providing the non-thermal equilibrium and non-chemical equilibrium effects inhérent in hypersonic flight. More spécifie details about varions ground facility types with their operation principles will be discussed in section 1.2.

Numerics are normally the cheapest and easiest option to incorporate pelding the capability to map the entire entry corridor if the physics are modeled appropriately. However, appropriately modeling the physics is quite a chal

lenge as the thermal non-equilibrium and Chemical non-equilibrium effects in

hérent of h5q>ersonic 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. Groxmd tests are the typical validahon tool as they can be implemented more easily and controlled to spécifie phenomena of interest more accurately. Since these groxmd-test and flight test results are limited in the h5q3ersonic régime, CFD tends to hâve a greater impact in the hypersonic régime 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 il

lustrâtes that these three fields are needed to complété 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 vmder- standing the hypersonic régime and the success of real flight vehicles dépends heavily on an itérative évolution 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 nu

merical tool development.

1.1.1 Characteristics of Hypersonic Flows

There are several important physical phenomenon that are associated with hy- persorûc aerodynamics. These are as foUows:

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Figure 1.1: CFD, ground tests and flight data ail rely on one another and are ail needed to complété the entire picture of hypersonic flight Th in 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 ré

gion becomes thinner as the shock angle decreases. Fligh-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 boimdary layer. This phenomenon can be taken advantage of leading to the simple and straightforward thin shock layer theory.

Substantial Entropy Layer

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

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

U X (V X u) = V/io — TVs (1.1)

Total enthalpy across a shock, curvature or not, does not vary (V/io = 0).

Thus, if vorticity is created as the flow crosses a curved shock (ü x ( V x u) 0), then there exists an entropy gradient. Near the stagnahon streamline there is a

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1.1 The Hypersonic Régime higher curvatvire gradient creating a strong entropy zone in the stagnation ré

gion 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 température. This effect is known as viscous dissipation. Owing to the température 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 grow so rapidly. This thick boundary layer will displace the mviscid portion of the flow which will effectively présents 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 température or real-gas effects. BCinetic energy transformed to thermal energy will first excite vibrational modes of the molécules. This effect is actually calorie imperfection where spécifie beats now depended on température since energy storage dépends on température. At higher velocities dissociation occurs with a composition dependence on tem

pérature 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 spécifie beats must be treated for a mixture of thermally and calorically Lmperfect gases, hence, real gas. Further details of Chemical non-equilibrium and real gas effects and as they apply to grotmd testing will be discussed later.

1.1.2 Gas Classifications Owing to High Température and High Pressure Effects

These spécial hypersonic phenomenon lead to very unique gas modeling con

ditions. For clarification, some of these classifications are described here.

Real Gas

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

• Compressibility effects

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• Variable beat capacity

• Close interaction forces

• Non-equilibrium thermodynamic effects

• Molecular dissociation and elementary reaction with variable composi

tion

This is the most complété 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 fonctions of tempér

ature alone.

e = e{T)

h = h{T) (1.2)

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

P = mkT (1.3)

Calorically Perfect Gas

A calorically perfect gas is defined as a gas in which the spécifie beats do not dépend on température, such that enthalpy and etemal energy relations can be described as:

dh = CpdT de = CvdT e = CvP

h = CpT (1.4)

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

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1.1 The Hypersoiûc Régime

Thermally Perfect but Calorically Imperfect Gas

As the gas is heated further, vibrational energy modes of gas molécules are significant and molecular spécifie beats are no longer constant but now dépend on température. As long as the particular molecular specie is far from conden

sation it can be classified as calorically imperfect but thermally perfect. The enthalpy, h,must now take into account the thermal energy stored in vibra

tion.

dh = Cp{T)dT

de = Cv{T)dT

e = c^(T)T

h = Cp{T)T (1.5)

Equation 1.6 is a common model of the vibrational energy storage for the spé

cifie heat for linear triatomic molécules such as CO2. Température dépendent spécifie heats need to be considered when the température approaches the characteristic vibrational température, 9p.

Cp _ 7 / 9p/2T y

R

2 ^

\smh{9p/2T) J ^ ’

Chemically Reacting Gas

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

h = h{T,NuN2,N3,...,Ni) e = e{T,NuN2,N3,...,Ni) Cp^fi{T,Ni,N2,N3,...,Ni)

Cp = f2{T,Ni,N2,N3,...,Ni) (1.7)

where Ni is the number density of specie "i". For the equilibrium case where the Chemical composition remains constant over time the number density val

ues are a function of the température and pressure which leads to:

h = h{T, u) e = e{T, u) Cp = fi{T,p)

Cp = f2[T,u) (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 hyper

sonic groimd test facility can be summarized in one simple sentence: hyperve

locity 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 température equilibrium State. The afore-mentioned hypersonic effects only become relevant in flight after the body beats and pressurizes the flow in a strong shock compression process. However, in a non-ballistic-type grormd 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 fa

cilities 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 "h}q>ervelocity" testing. The term "hy

pervelocity" is commonly used over hypersonic as the key feature is more the velocity rather than the Mach number as this gives a Armer 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 varions hypersonic/hjqjervelocity 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 équation for inviscid compressible high velocity flow:

Ano — ^0 H--- Po

P U21

^ + T + T P 2

**oo

2 (1.9)

High enthalpy hypervelocity facilities (continuons timnels, arc-heated plasma tunnels, impulse tunnels) require energy addition to an initial stagnation con

dition to achieve high-enthalpy. One could focus on increasing the stagnation internai energy, eo, as in arc-heated facilities, détonation facilities and combus

tion timnels. 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 stagna

tion 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 hâve vibrational modes of energy

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1.2 Hypervelocity Facilities for Ground Testing Hypersonic Aerodynamics

Figure 1.2: Typical hypervelocity accélération process through a nozzle. A high-energy '"stagnant"' condition (state 1) is accelerated by a noz

zle (state 2) into a hypersonic free-stream flow (state 3)

stored or to hâve 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, t/iow/ is much larger than the vibrational relaxation time, Tvh, and the Chemical reaction char

acteristic time, Tchem, {Tfiow 3> T^ib, Tchem) then the flow can be said to behave in local equilibrium mariner. On the other extreme side, if the characterisitc flow time is much smaller than the vibrational relaxation time or the Chemical reac

tion time (Tfiow 'Ai6> Tchem) the flow is said to be in afrozen state. 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 in non-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 hâve the gas arriving at the test section be in local thermodynamic and Chemical equilibrium as in flight but this becomes more difficult as timnel en- thalpy requirements increase. Sometimes, though, these non-equilibrium ef- fects can be studied and if measured properly provide value information for the validation of numerical tools.

Because of these effects, accurate détermination of flow conditions in either the free stream or near a model is quite difficult. Tunnel free stream conditions in perfect gas h3q)ervelocity facilities can be estimated using isentropic nozzle flow relations. However, owing to real gas effects, these calculations made with plénum conditions and isentropic nozzle équations can be erroneous when sur- passing 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 beat flux and pressure data acquired from intrusive probing. This infor

mation is then used to rebuild the flow field with some model. This is not a

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direct straight forward method data réduction method. The non-equilibrium conditions will also complicate the data réduction process as simple models de- scribing the flow properties, while useful in low-enthalpy envirorunents, begin to lose their fidelity at higher enthalpies.

As stated earlier, many different types of facilities are needed to map a spé

cifie 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 pro

vide an estûnate of the entry environment. The following tw^o sub-sections describe some of the most common groimd test facilities used in hypersonic 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 réplication of the intense thermal environ- ments 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 understandmg of entry physics. An ex

cellent review of the two main types of arc heaters, the Huels and segemtented type, can be foimd in reference [10].

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

Huels type of heater. The sûnplidty of design and relative ease of operation allow for easy mamtenence and shor tum aroimd 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 mto kinetic energy. Huels type of arc heaters can nm at relatively high pressure but at a lower enthalpy as the current density must be relatively low to maintain integrity of électrodes. Also an issue is that the arc discharge has little no constramt on termmation 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 électrodes. This leads to only a moder

ato 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 tum increases the stability and repeatability of testing.

Higher enthalpy levels can be achieve as high current arcs can be used since

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the arc elctrode heating is spread over many électrodes. However, design com- plexities limit operating pressures to generally lower than that of Huels-type heaters. Segmented heaters also require more frequent inspection and mainte- nence than Huels-type.

Inductively Coupled Plasma (ICP) Facilities

One approach to avoid contamination from érosion of cathode and anode ma- terial is to change the method of adding energy to the flow such as an induc

tively coupled plasma (ICP). An ICP facility consists of a large current power supply cormected to a coil surrormding a quartz tube. A large current flows through the coil creating a strong electromagnetic field. This electromagnetic field ionizes the gas and the électrons impart considérable energy to the gas by collisions. ICP facilities typically opéra te in the subsonic (some times sonie!) régime but they are capable of creating the high thermal and chemistry effects inhérent 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.

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 hâve 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 ail impulse facilities are based, to some extent, on its operation. Consider équation 1.10 which is a relation for shock tube parameters as illustrated in figure 1.3.

74-1

(1.10)

^ ^ 27iM| - (7i - 1)

■Pi 7i + 1 (7i-lKV^ Ms

This équation 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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Flow Distance

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

Figure 1.3: Staitdard shock-tube

with the driver and driven gas constituents through the spécifie beat ratios, 7i, and local sound speeds, üi- However, there are implications to altering these parameters which must be considered especially if one wants to simulate chemistry properly. Typical facilities may run hélium or hydrogen as driver gases generating only modest enthalpy levels.

The flow behind the incident shock, région 2, is at a relatively high enthalpy State. Recall that vibrational excitation dépends on the température and begins at arotmd 600K for N2 and O2. Using the normal shock relation in équation 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 température effects in a simple facility makes the shock-tube facility attractive for experiments aimed at numerical validation of vibrational and Chemical kinetics.

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 accélération tube and is maintained initially at low pressure. When the downstream traveling shcKk ruptures the diaphragm to the accélération tube, an imsteady expansion process will convert the shocked thermal energy into kinetic energy thus increasing the flow velocity. This addition helps to overcome the Mach number

Tl

[27iMi^ - (7 - 1)] [(7-l)Mf + 2]

(7 + 1)"m2 (1.11)

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limitation of shock tubes. An accélération tube process is generally more effi

cient 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 ex

pansion tube facilities hâve very short run times. These are similar limitations shared with shock tube facilihes.

1.2 Hypervelodty Facilities for Groimd Testing Hypersonic Aerodynamics

DIaphragm DIaphragm DIaphragm

Figure 1.4: Standard expansion type shock-tube

(Reflectionless) Shock Tunnel

The Mach number limitation of the shock tube can also be overcome by in- corporating a nozzle tuming the shock or expansion tube facility into wMt is known as a shock tunnel. This shock-tube adapted facility consists of a tra- ditional shock tube with an expanding nozzle at the end. However, the large exparisions involved create a limiting Reynold's number.

The shocked gas in the driven section flows straight through the nozzle and driven tube interface without shock reflections passing upstream. This facil

ity type will hâve lower enthalpy than reflected-type tunnels pelding a lower level of vibrational excitation, Chemical non-equilibirum and real gas effects.

The test section enthalpy after the gas has been accelerated by the nozzle will be around the same level as the initially shocked gas. Modest hypersonic Mach numbers are achievable with this facility type nearing 10. However, the run times (dictated as the time between initial test gas arrivai and interface arrivai at the nozzle) are again quite short. This type of facility is generally used for model testing owing to the larger flow diameter which is needed to accommo- date such experiments.

A problem exists though. Potentially, the flow plénum conditions may al-

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Figure 1.5; Standard expansion type shock-tube

ready hâve some stored vibrational energy and dissociation and not allow to relax through the nozzle thus being partially frozen. With a model installed, further shock heating will occur addrng to already vibrationally excited and dissociated gas. This gas will go through further vibrational excitation and dissociation owing to the shock near the model which will be different than that of an assumed gas in the free-stream with no vibrational excitation or dis

sociation.

Reflected Shock Tunnel

In a reflected shock tunnel, an incident shock beats and pressurizes the test gas traveling at high velocity toward the test section. This incident shock is then re

flected off of the constricted région of the tunnel and propagates back upstream through the test gas. This process further beats and pressurizes the flow but at the same time stagnâtes the gas into a Virtual réservoir where high enthalpy fluid accumulâtes. This high enthalpy gas can then be accelerated through an expansion process via a converging diverging nozzle similar to a blow-down facihty to produce the hypervelocity test flow. This facility type can beat the test gas to very high enthalpy levels, but only for a short time. This test time is govemed by the time from the arrivai of the incident shock to the arrivai of the driver/driven gas interface or the arrivai of the leading expansion wave.

A tunnel is considered to be using "tailored" conditions when the driver gas is halted by the reflected shock leading to maximum run times for testing. This high enthalpy is quite significant and will created substantial vibrational exci

tation and non-equilibriiun conditions that freeze during the rapid accélération

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1.2 Hypervelocity Facilities for Grotmd Testing Hypersonic Aerodynanücs to the freestream conditions which significantly impacts model test data.

Test Time

Distance

DriverGas(6) Driven/TestGas(l)

Diaphragm

Figure 1.6: Reflected shock tunnel hypervelocity facility

Gun Tunnel

The earliest use of piston compression was in gun-tunnel type facilities. Fig

ure 1.7 illustrâtes this facility type, which is basically a modified reflected shock tunnel. The double diaphragm section is replaced by a light piston which sép

arâtes the driver and driven gases. When the diaphragm supporting the piston is ruptured, the piston rapidly travels down the barrel of the driven section like a bullet in a gim. A strong shock forms ahead of the piston and is reflected several times between the end of the barrel and the piston front face before the piston cornes to rest. The gas volume ahead of the piston is heated and compressed mechanically by the piston and gas-dynamically by the leading shock leading to very high pressures and températures before the accélération process. This technique is very favorable in creating large Reynold's number but the stagnation enthalpy levels are limited.

Free-Piston Tunnel

The free-piston shock tunnel (FPST) (figure 1.8) is a facility designed to make up for some of the shortcomings of the gun tunnel facility. The facility has a more complex initial section comprised of a high-pressure air réservoir that is used to move the piston and a driver section whose gas contents are compressed by the large piston. This driver gas is compressed imtil the high pres

sure ruptures the diaphragm between the driver and driven portion of the

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Test Time

Distance

Piston at Initial Position

Piston During Flight

Pis*

Fin

High Pressure Driven/Test Gas Test

Section

Figure 1.7: Free-piston gun tunnel (i.e. Longshot)

shock tube. The driven gas is then accelerated through the typical reflected shock tunnel process. The downside to this facility type is that the test times, even rurming in optimized tailored conditions, are on the order of a few milliseconds at best whereas gun tunnel facilities can run for over 20ms.

Hot shot tunnel

The hotshot tunnel is a hypervelocity facility within which high-enthalpy is generated by rapidly discharging a large amount of energy via an electric arc into a fixed volume of gas. TÏüs high enthalpy gas is then allowed to expand and accelerate through a nozzle into a test section. The high energy arc poses similar contamination problems as arc-heated plasma facilities and the arc dis

charge has the potentiel to erode many components in the arc chamber. The érosion of the nozzle throat owing to significant heating is also an issue with the high température arc chamber conditions. Test times can run into the several himdreds of milliseconds but this significantly adds to the contamination problem as components will beat up dramatically with increased lun time. This longer run duration also poses a risk to the survivability of the throat of the tunnel. Chemical effects in the arc chamber of these facilities can be quite high and with the rapid expansion process there is little time for Chemical recombi

nation yielding to frozen flow chemistry that can be far from equilibrium.

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Time

Figure 1.8: Free-piston shock tunnel

ArcChamberwtth TestGas

Figure 1.9: Hotshot impulse facility

Disponible à / Available at permalink :

D 03673

Tunable Diode Laser Absorption Spectroscopy Characterization of

impuise Hyperveiocity CO 2 Flows

Jason Matthew Meyers

Université Libre de Bruxelles

Faculty of Applied Sciences ULB

Von Karman Institute for Fluid Dynamics Aeronautics and Aerospace Department

Supervisor: D. G. Fletcher (VKI) Promoter: G. Degrez (ULB)

Committee:

F. Dubois (ULB) M. Herman (ULB) A. K. Mohamed (ONERA)

U. Koch (DLR)

Contents

Nomenclature

Latin Letters

Greek Letters

Subscripts

PREFACE

M

s

1 Introduction

1.1 The Hypersonic Régime

H

1.1.1 Characteristics of Hypersonic Flows

1.1.2 Gas Classifications Owing to High Température and High Pressure Effects

2 ^

1.2 Hypervelocity Facilities for Ground Testing Hypersonic Aerodynamics

1.2.1 Hypersonic Generalities of Ground Test Facilities

1.2,2 Common Plasma Facility Types

1.2.3 Common Impulse Hypervelocity Facility Types

impuise Hyperveiocity CO ² Flows