Fuel-Air Mixing and Combustion in Scramjets

18 Pages·2002·1.1 MB·English

by  Drummond, J. P.

AIAA 2002-3878
Fuel-Air Mixing and Combustion in
Scramjets
J. P. Drummond and Glenn S. Diskin
NASA Langley Research Center, Hampton, VA
A. D. Cutler
The George Washington University, Hampton, VA
38th AIAA/ASME/SAE/ASEE Joint
Propulsion Conference and Exhibit
7-10 July 2002
Indianapolis, IN
For permission to copy or to republish, contact the copyright owner named on the first page.
For AIAA-held copyright, write to AIAA Permissions Department,
1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344.

AIAA-2002-3878
FUEL-AIR MIXING AND COMBUSTION IN SCRAM JETS
J. Philip Drunnnond* and Glenn S. Diskin t
NASA Langley Research Center. Hampton, Virginia, USA
Andrew D. Cutler _
The George Washington University
Joint Institute for Advancement of Flight Sciences, Hampton, Virginia, USA
Abstract order as the fluid scale. Both nfixing and combustion
time scales must be considered in a detailed study of
Activities in the area of scramjet fllel-air mixing and mixing and reaction in a scram jet to understand the
combustion associated with the Research and Tech- flow processes and to ultinmtely achieve a successflfl
nology Organization Working Group on Technologies design.
for Propelled Hypersonic Flight. are described. Work Although the geometric configuration of a scram-
discussed in this paper has centered on the design of
jet is relatively simple compared to a turbonmchinery
two basic experiments tot studying the nfixing and
design, the flow physics associated with the simulta-
combustion of fuel and air in a scramjet. Simula,
neous injection of fuel from nmltiple injector config-
tions were conducted to aid in the design of these ex- urations, and the mixing and combustion of that fuel
periments. The experimental models were then con- downstream of the injectors is still quite complex.
structed, and data were collected in the lal)oratory. For this reason, many researchers have considered
Comparison of the data from a comxial jet mixing the more tractable problem of a spatially develol)-
experiment and a supersonic combustor experiment ing, primarily supersonic, chemically reacting mixing
with a combustor code were then made and described.
layer or jet. that relaxes only the complexities intro-
This work was conducted by NATO to validate com-
duced by engine geometry. All of the difficulties in-
bustion codes currently employed in scramjet design troduced by the fluid mechanics, combustion chem-
and to aid in the development of improved turbulence
istry, and interactions between these phenomena can
and combustion models employed by the codes. be retained in the reacting mixing layer, making it.
Introduction an ideal problem for the detailed study of supersonic
reacting flow in a scramjet. With a good understand-
At.flight speeds, the residence time for atmospheric ing of the physics of the scramjet internal flowfield.
air ingested into a scramjet inlet and exiting from the designer can then return to the actual scramjet
the engine nozzle is on the order of a millisecond. geometry with this knowledge and apply engineering
Theretbre, fuel injected into the air must efficiently design tools that more properly account for the con>
mix within tens of microseconds and react, to release plex physics. This approach will guide the discussion
its energy in the combustor. The overall combustion in the rentainder of this section.
process should be mixing controlled to provide a sta-
ble operating environlnent: in reality, however, com-
bustion in the upstream portion of the Colnbustor, Reacting Mixing Layers and Jets
particularly at. higher Mach numbers, is kinetically As described earlier, compressible shear/nfixing
controlled where ignition delay times are on the same
layers and jets provide good model problems for
*Hypersonic Airbreathiug Propulsion Branch, Associate studying the physical processes occurring in high-
Fellow, AIAA speed mixing and reacting flow in a scramjet. Nlix-
tLaser & Electro-Optics Branch, Senior Member, AIAA
ing layers are characterized by large-scale eddies that
*Senior Member. AIAA
form due to the high shear that is t)resent between
°Copyright @2002 by the American Institute Institute of
Aeronautics and Astronautics, Inc. No copyright is asserted in the fuel and air stremns. These eddies entrain fuel
the United States under "fitle 17, U.S. Code. The U.S. Gov- and air into the mixing region. Stretching occurs
ernment has a royahy-free license to exercise all rights under
in the interracial region between the fluids leading
the copyright claimed herein for Governmental purposes. All
to increased surface area and locally steep concen-
other rights are reserved by the copyright owner.
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American Institute of Astronautics and Aeronautics
A1AA-2002-387_
trationgradients.Moleculadriffusionthenoccurs
The growth of a mixing layer produces a displace-
acrossthestrainedinterfacesT.herehasbeenasig-
ment effect, on the surrounding flow field. This dis-
nificamamountofexperimentaalndnumericarle-
placement in confined flow produces pressure gradi-
searchtostudymixing
layer and jet flows. 1-,DFor the ents that can affect the later development of the nfix-
same velocity and density ratios between fuel and air, ing layer, typically retarding growth. When chemical
increased compressibility, to the levels present, in a reaction occurs in a mixing layer, resulting in heat
scramjet, results in reduced mixing layer growth rates release, the growth of the mixing layer is retarded in
and reduced nfixing. The level of compressibility in a botil subsonic and supersonic flow. i'2 The effect of
mixing layer with air stream 1and tirol stream 2 can heat release call also vary spatially as a function of
be approximately characterized by the velocity ratio, the local st.oichiometry and chemical reaction. The
1"= [-.72ii[.;1, tile density ratio, s = P2/Pl, and the retarded growth in both instances call be reversed,
convective Mach number, Jl<, = (U2 - U1)/(al + a2) however, by allowing the bounding wall to diverge rel-
where a is the speed of sound. Increased compress- ative to the initial wall angles where retarded growth
ibility reorganizes the turbulence field and nlodifies was noted.1 While the process of nfixing layer growth
the development of turbulent structures. The restllt- is affected by the combustor geometry and design,
ing suppressed transverse Reynolds normal stresses fuel injector design carried out wit.h proper consider-
appear to restilt in reduced monmntum transport. In ation for the inlet, and combustor geometry can have a
addition, the primary Reynolds shear stresses respon- strong influence on overall nfixing and combustion ef
sible fbr nfixing layer growth rate also are reduced. fieiency. Considerable effort has been expended over
The priniarv nfixing la_vr instability beconles three- the past. fifteen years to achieve efficient fuel injector
dimensional with a convective Mach number above designs. Injector design will be considered in the next
0.5, reducing the growth of the large scale eddies. section.
Finally. the turbulent eddies become skewed, fat,
and less organized as compressibility increases. All
of these effects combine to reduce the growth rate of Scramjet Fuel Injectors
the mixing layer and the overall level of nfixing that
There are severa] key issues thai nmst be consid-
is achieved.
ered in lain design of an efficient fuel injector. Of
Several phenomena result in the reduction of mix- particular importance are the total pressure losses
ing with increasing flow velocity, including the veloc- created by the injector and the injection processes.
ity differential between fuel and air. and compressibil- that nmst be nfininfized since the losses reduce the
it.v. Potentially, the existence of both high and tow thrust of the engine. The injector design also nmst
growth and mixing rates are possible, and the engine produce rapid nfixing and combustion of the fuel and
designer with an understanding of the flow physics air. Rapid mixing and combustion allow the con>
controlling these phenomena can ad_,'antageously use bustor length and weight, to be minimized, and they
these effects. The shock and expansion wave struc- provide the heat release tot conversion to thrus_ by
ture in and about the mixing layer can interact with the engine nozzle. The tirol injector distribution in
the turbulence field to affect mixing laver growth. 1 the engine also should result in as uniform a com-
Shock and expansion waves interacting with the laver bustor profle as possible entering the nozzle so as
result from the engine internal structure. Experi- to produce an efficient nozzle expansion process. At
ments have shown that the shocks that would re- moderate fight Mach numbers, up to Math 10. fuel
sult from wall and strut compressions appear to en- injection may have a nornml component into the flow
hance the growth of the two-dimensional eddy struc- from the inlet, but at higher Mach numbers, the in-
ture (rollers) of a mixing layer. This effect is most jection nmst be nearly axial since the fnel nlomentum
pronounced when the duct height in the experiment provides a significant portion of the engine thrust. In-
and the shear layer width become comparable. Waves trusive injection devices can provide good tirol disper-
nmy be produced by the mixing layer itself under ap- sal into the surrounding air. but they require active
propriate conditions. Localized shocks (ofi.en termed cooling of the injector structure. Tile injector design
shock]ets) occur within the mixing layer when the and the flow disturbances produced by injection also
accelerating flow over an eddy becomes supersonic should provide a region tot flameholding, resulting
even when the surrounding flow is subsonic. When in a stable pilot.ing source for downstream ignition
the overall flow is supersonic, the eddy shocklets will of the fuel. The injector cannot result in too severe
extend as shocks into the flow beyond the individual a local flow disturbance, that could result in locally
eddies. These shocklets can retard eddy growth due high wall static pressures and temperatures, leading
to increased localized pressure around the eddy. to increased frictional losses and severe wall cooling
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AIAA-2002-3878
requirements. of the features of the unswept ramp, but. the sweep
results in better axial vorticity generation and mix-
Anumbeorfoptionsareavailabletbrinjectingfllel
andenhancintghemixingofthefuelandair ing. A novel variation on the swept wedge injector.
in high-
termed the aero-ramp injector, l:t utilizes three arrays
speed flows typical of those found in a scramjet com-
of injector nozzles at various inclination and yaw an-
bustor.t0.11 Two traditional approaches for inject-
gles to approximate the physical swept ramp design.
ing fllel include injection fl'om the combustor walls
and ila-st.ream injection fl'om slruts. The simplest ap- In-stream injection also has been utilized for fuel
proach ibr wall injection involves the transverse injec- injection in a scramjet. Traditional approaches in-
tion of the fllel from wall orifices. Transverse injectors volve fnel injection from the sides and the base of
offer relatively rapid near-field mixing and good fuel an in-stream strut.. Transverse injection results in
penetration. Penetration of the fuel strealn into the behavior similar to transverse fueling from the wall.
cross-flow is governed by the jet-to-freestream mo- although differences can occur due to much thinner
mentum flux ratio. The fllel jet interacts strongly boundary layers on the strut. Injection from the base
with the cross-flow, producing a bow shock and a of the strut results in slower mixing as compared
localized highly three-dimensional flow field. Result- to transverse injection. A combination of transverse
ing upstream and doxvnst.ream wall flow separations and st.reamwise injection, varied over the flight Mach
also provide regions tor radical production and flame- number range, often has been utilized to control re-
holding, but they can also result in locally high wall action and heat release in a scramjet combustor. As
heat transfer. Compressibility effects that were noted noted earlier, however, streamwise injection h_ the
earlier for mixing layer flows also are evident in the advantage of adding to the thrust, component of the
mixing regime downstremn of a transverse jet. Coln- engine. To increase the mixing from streamwise in-
pressibility again retards eddy growth and breakup in ject,ors, many of the approaches used to improve wall
the mixing layer and suppresses entrainmellt of fllel injection, including non-circular orifices, tabs. and
and air, resulting in a reduction in mixing and reac- ramps, have been successfully utilized. Several new
tion. concepts have emerged as well. Pulsed injection us-
ing either mechanical devices or fluidic oscillation
Improved mixing has also been achieved using al-
techniques have shown pronfise for improved mixing.
ternative wall injector designs. Wall injection using
Pulsed injection of fllel utilizing a shuttering tech-
geometrical shapes that introduce ,axial vorticity into
nique to control injection has been shown to improve
the flow field has been successful. Vorticity can be in-
mixing.r" Fuel injection schemes integrated with cav-
duced into the fuel st.ream using convoluted surfaces
ities also provide the potential for improved mixing
or small tabs at the exit of the fuel injector. Al-
and flameholding. This type of integrated fllel injec-
ternatively, vorticity can be introduced into the air
tion/flameholding device, utilizing fuel injection into
upstream of the injector using wedge shaped bodies
a cavity and from its base, integrates the fuel injec-
placed on the combustor walls. Vorticity addition to
tion with a cavity that provides flameholding, flame
the air stream provides more significant mixing en-
stabilization, and mixing enhancement if the cavity
hancement of fuel and air .12 When strong pressure
is properly tuned. Air exchange rates with the cav-
gradients are present in the flowfield, e.g. at. a shock,
ity mas" be low, however, limiting the amount of fuel
vorticity aligned with the flow can be induced at. a
that can be added. Additional scramjet fuel injec-
fllel-air interface, where a strong density gradient ex-
tor designs continue to be introduced and studied to
ists. by virtue of the baroclinic torque. Fuel injection
achieve even higher levels of mixing and combustion
ramps have proven to be an effective means |br fuel
efficiency.
injection in a seramjet engine. 2 Fuel is injected from
the base of the ramp. The unswept ramp configu- Scramjet design is built upon both experimental
ration provktes nearly streamwise inject.ion of tilel to and computational research. To assure that com-
produce a thrust, component. The effects of angled putational tools properly represent the complex flow
injection on axial thrust only go as the cosine of the physics in a scramjet, careful evaluation of t.he com-
angle, so small injection angles result, in little loss putational tools is necessary. Benchmark experi-
in thrust. Flow separation at the base of the ramp ments are becoming available that provide the nec-
provides a region for flame holding and flame stabi- essary data for evaluating the accuracy of the nu-
lization through the buildup of a radical pool. The merical algorithms and the physical models that the
ramp itself produces streamwise vorticity as the air computational tools employ. In addition, these ex-
stream sheds off of its edges, improving the down- periments provide in some instances the information
stream mixing. The swept ramp design provides all necessary to improve the modeling employed by the
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codesT. woexperimenatsvailableforassessinhgigh-
speedcombustiocnodesaredescribeidnthenextsec-
tion.Resultosbtainedfromtheapplicationofacon>
bustioncodet,otheseexperimenatsrealsoshownand
discussed.
Mixing and Combustion Experiments
Two basic experiments are being conducted at
the NASA Langley Research Center to collect, de-
tailed high-speed mixing and combustion data for
use in physical model development and code valida-
tion. The first experiment concerns comxiat jet mix-
ing of a helium/oxygen center jet with a coflowing air
outer jet and was chosen to provide detailed super-
sonic mixing data. The second experiment was de-
veloped to study high-speed mixing and combustion
in a simple "scramjet like" engine environment. The
experiment utilizes a ducted flow rig containing viti-
ated supersonic air with a single tirol injector that in-
troduces supersonic g_eous hydrogen from the lower
Figure 1: Coaxial jet ,_ssembly cross-section
walt. Detailed wall and in flow surveys and nonin-
terference diagnostics are used in both experiments.
These experinmms will be described in the following
sections.
air into which the air jet exhausts. Shock-expansion
wave structure emanating outward from the center-
Coaxial Jet Mixing Experiment
body nozzle lip can also be seen. Inward propagating
A coaxial jet mixing experiment h_ been devel- waves from the inner lip. due to the finite thickness of
oped to stud,v the high-speed compressible mixing of the lip (0.25 ram), can be observed in the air jet once
helimn and air. Details of the experiment are de- they pass through the helium jet. These waves are
scribed in references a4 and. x'_ The low-density he- not visible in the helium jet due to the low refractive
lium, which serves as a sinmlant of hydrogen fuel, index within the center jet. A third wave can also
was chosen to allow detailed studies of mLxing with- be observed emanating inward from the outer noz-
out chemical reaction. Oxygen is added to the he- zle lip and traversing both the air jet and the helium
lium jet as a diagnostic aid for an oxygen flow-tagging jet. Additional results from the experiment will also
technique (RELIEF). Several methods are utilized to be considered later in this section when comparisons
characterize the flow field including Schlieren visu- of the measured data with numerical sinmlations are
alization, pitot pressure, total temperature, and g_ made.
sampling probe surveying, and RELIEF velocinmtry.
A schematic of the coaxial jet configuration is shown SCHOLAR Combustor Experiment
in Figure 1. The rig consists of a l0 mm inner nozzle
A direct-connect supersonic combustor model,
froln which helium, nfixed with 5 percent oxygen by
known by its acronym SCHOLAR, has been devel-
volume, is injected at Mach 1.8 and an outer nozzle
oped for testing in a combustion heated test facil-
60 mm in diameter from which coflowing air is imro-
ity at the NASA Langley Research Center. This ex-
duced also at Math 1.8. The velocity ratio between
periment has been designed to provide optical access
the two jets is 2.25, the convective Mach number is
to a reacting supersonic flowfield typical of the flow
0.7, and the jet exit pressures are nmtched to one
present in a scramjet engine. Details of the experi-
atmosphere.
ment are described in reference. 1_;The model shown
The resulting flow downstream of the nozzles can in Figure 3a consists of a section 546 mm in length
be seen in Figure 2, which shows a Schlieren image constructed of copper for thermal control followed by
of the flowfield. The development of the nfixing layer a 914 mm long section of carbon steel attached to
between the cemral helium jet and the air jet can be the aftr end of the copper section. The copper section
seen along with the shear layer development between contains a single fuel injector that introduces gaseous
the air jet and the surrounding quiescent laboratory hydrogen into the vitiated air stream flowing through
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American Institute of Astronautics and Aeronautics
A1AA-2002-3878
'" A- _-,2,. ! i ui i ....... e.... l!
{al v-
U
D.,_ml,m. m,T_u
(b)
Figure 3: SCHOLAR conlbustor model: (a) nozzle.
copper and steel duct sections: (b) detail near fuel
injector and pilots.
Figure 2: Schlieren inmge of coaxial jet mixing (con- upper wall of the carbon steel section that is attached
ical extension cap removed) to the copper model.
Five measurement locations for optical access are
provided in the copper section of the model. Two
additional measurement stations are provided in the
the model. carbon steel section. The measurement stations in-
dicated in Figure 3a are slits in both model side
The injector region of the combustor model is
walls, through which planar BOXCARS measure-
shown in Figure 3b. The model consists of a con-
ment bemns enter and exit, allowing single-shot mea-
stant area channel initially 38.6 mm high and 87.9
surements to be made of static temperature (ro-
mm wide tollowed by a 4.8 rain rearward-facing step
vibrational tenlperature of N., molecules). These
and a 48.8 mm long constant, area section. Combus-
measurements are single point measurenmnts, but the
tion heated vitiated air is introduced into the channel
location of the measurement is translated during the
at Mach 2, 1184 K, and 100405 Pa. The vitiated air
tests to provide a full plane of data at each station.
contains 20.35 percent water by mass introduced by'
From these single-shot nmasurements, averages and
the facility heater. A small amount of hydrogen fuel
RMS values are derived. In additional to optical mea-
at a tirol equivalence ratio of 0.15 is introduced down-
surements, wall pressures are measured using an ar-
stream from five pilot injectors along the upper wall
ray of orifices.
at. the 74.1 mm station. Each pilot, injector is 1.27
mm in diameter. Thee central pilot injector lies on
the duct. centerline, and the remaining four injectors Simulation of Mixing and Combustion
are spaced 12.7 mm from one another on either side of Experiments
the central jet.. This fuel is used to pilot, the prinmry
Mach 2.5 hydrogen tirol injector further downstream The mixing and combustion exl)eriments described
at the 173.2 mm station by igniting and producing a earlier were numerically sinmlated betore data was
radical pool to enhance ignition of the prinmry fuel- collected to assist in the experimental design. Ad-
air mixture. The prinmry fuel injector is inclined at ditional simulations were also performed during and
ao degrees to the horizontal and has a circular cross- following t.he experimental study to compare with the
section 7.6 mm in diameter. A a degree expansion of nmasured data. Initial sinmlations were made with
the upper wall begins immediately at the primary tirol the SPARK combustion code. Additional studies
injector. This 3degree expansion continues along the with other combustion codes are being conducted. L'
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American Institute of Astronautics and Aeronautics
AIAA-2002-3878
Detailson
tile simulations of both experiments and
comparisons wit 11measured data are given ill the fol-
lowing sections.
Simulations of the Coaxial Jet Mixing
Experiment
Tile axisynnnetric version of the SPARh: code was
used to sinmlate the flowfield in the helium/oxygen
center nozzle and the outer air nozzle of the coax-
Figure 4: Mach contours in inner nozzle
ial jet experiment. Details of the code are given in
reference, lr This version of the code solves the a,,:-
isymmetric full Navier-Stokes equations and species
continuity equations describing the production, con-
vection, and mixing of each relevant species. The 01 ....
analysis of the experiment was begun by first, solv- 008
ing for the flowfield in the center and outer nozzles
using the nozzle contours specified in the last, sec-
tion. The domain of the inner nozzle was discretized
with a grid of 201 points in the streamwise direction
and 51 points in the radial direction. The domain of
the outer nozzle was discretized with a grid of 201
points in t.he streamwise direction and 101 points in
Figure 5: Mach contours in outer nozzle
the radial direction. Initial conditions in the subsonic
portions of the nozzles were specified at. the x = 76.2
mm station. For the inner nozzle at this station, the
streamwise velocity was 141.8 m/s, the static temper- for the downst.ream donmin beyond the nozzles where
ature was 297.4 K. and the static pressure was 614300 mixing of the jets occurred. The downstream domain
Pa. The helium mass fraction in the inner jet was was 150 nun long. The upper boundary of"the domain
0.6995 and the oxygen mass fraction was o.aooa. In was specified at y = 30.24 mm to coincide with the
the outer nozzle at the 76.2 mm station, the stream- end of the outer nozzle wall. Radiall,v beyond this
wise velocity, static temperature and pressure of the point., laboratory air is entrained by the outer jet.
air were 22.94 m/s, 299.74 K, and 578100 Pa, respec- but measurements have shown relatively low stream-
tively. No slip conditions were specified along the wise velocities of only a few meters per second in the
nozzle walls and first order extrapolation was used at entrained flow. Later measurements are planned to
the supersonic outflow station of each nozzle. Sym- confirm these conditions. It is not. appropriate to sin>
metry conditions were specified along the centerline ulate such a low speed flow with a compressible code,
of the inner nozzle. Tnrbnlence in the outer nozzle
so a slip boundary was chosen at y = 30.24 mm so as
was modeled using a Cebeci-Smith model. The flow to consider only the compressible regime. Large scale
in the inner nozzle was assumed to be laminar consis- structure certainly forms at this interface as the two
tent. with the behavior observed in the experiment at streams viscously interact. _¥aves from this st.ruc-
the outflow of the nozzle. Results from the center and ture reach the helium-air interface near the nozzle
outer nozzle sinmlations detailing Mach number dis- exit. However, the air coflow-ambient air mixing laver
tributions are given in Figures 4 and 5, respectively. does not. start to merge with the helium-air mixing
The initial station (x = 76.2 ram) of both nozzles is layer until the end of the experimental domain. The
reset, to zero in the calculations. The Mach number domain was discretized with a grid of 40t points in
ranges from a mininmnl of 0.1 in the subsonic por- the st.reamwise direction and 201 points in the ra-
tion of the nozzles to a maxinmm of 1.8 at the end
dial direction. The grid was radialh" compressed in
of each nozzle. As can be seen from the figures, both the region of the helium-air mixing layer. Synnne-
nozzles produce very, uniform exit flow fields resulting try boundary conditions were specified along the y =
in ideal initial conditions for the mixing study in the 0 boundary and slip conditions were specified along
region downstream of the nozzle exits. the upper boundary at y = 30.24 ram. The outflow
Results obtained at the end of each nozzle were boundary at. x = 150 nnn remained supersonic, and
titan used to specif',," the supersonic inflow conditions extrapolation conditions were specified at. this loca-
tion. Turbulence was modeled in the downstream do-
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American Institute of Astronautics and Aeronautics
A1AA-2002-_3b
0.05
he: 0.044 0.087 0.131 0.175 0.219 0.262 0.306 0.350 0.394 0.437 0.481 0.525 0.569 0.612 0.656 i
0.04
0.03
E
0.02
0.01
I
0 0 0.05 0.1 o.15
x, m
Figure 6: Helium mass fraction contonrs downstreanl of nozzles
main with the turbulent jet mixing model of Eggers 00t,:
and Eklund. 1_'1,_
001; x=O.Om
Results from the downstream calculation are shown x: 0.04303 m
1=0.00121 m
in Figures 6 through 8. Helium-air mixing down- - - x: 0,12074 m
stream of the nozzles is shown in Figure 6. Tile 0.0' -- x=0.15000m
1,_ L_ x=0.04293 m
helium mass fraction in the figure ranges from a
x= 0.08110 m
mininmm of zero to a maximum of approximately O.OOI
E
0.7. There is significant mixing of the helium and
air throughout the downstream region although rela- 0.006
tively high mass fractions of helium still remain near
the centerline. 0.004
A comparison of the measured helium m_s frac-
tion data with the simulation results at several sta, 0.002 .... , ....
tions downstream of the nozzles is given in Figure 7.
I
Agreement between the sinmlation and the data is 0; 0.25 0,5 075
very good at each station. The code somewhat over- Helium Mass Fraction
predicts the mixing near the cent.erline at tile x = 0.12
m station, although the prediction improves with in- Figure 7: Comparison of helium mass fraction data
creasing radial distance. A comparison of measured (symbols) with simulation results
pitot pressures with the simulation is shown in Fig-
ure 8. Agreement is good in the region of the air
coflowing jet, but. the sinmlation somewhat overpre-
dicts the pitot pressure in the helium-air mixing re- ing region between the helium and air coflowing jets
gion. The comparison with the experimental data in agreement with the pitot pressure results.
differs at large radial distances greater than 0.025 m
Simulations of the SCHOLAR, Combustor
as the code does not. consider the effects of the labora-
Experiment
tory air entrained by' the coaxial air jet. The RELIEF
stremnwise velocity data is compared with the sim- The three-dimensional version of the SPARK code
ulation in Figure 9. The prediction agrees well with was used to sinmlate the flowfield in the SCHOLAR
the data at the first three stations and slightly over- combustor model. Details of the code are given in the
predicts the data at. the remaining stations near the references. 17'2_j This version of the code solves the
centerline. The sinmlation somewhat underpredicts 3-D flfll Navier-Stokes equations and species continu-
the the velocity at the fnal three stations in the nfix- ity equations describing the production, convection,
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American Institute of Astronautics and Aeronautics
A1AA-2002-.387S
and mixing of chelnical species. Calculations have
been used in the design and reflnemem of the experi-
ment. In tim calculation the model was.rotated from
0.03 the orientation shown in Figure .3such that the in-
jector wall was aligned with the lower computational
boundary.
0.025
Calculations were begun at the x = 0 station of the
..... x: 0.04303 m 1 SCHOLAR model where vitiated air from the facil-
0.02
it3' enters the duct. Vitiated air entered the model
x,, 0.08121 m
x=0.12074 m | at Mach 2.0 yielding a velocity of 139.5.7 m/s, a
x- 0.15000 m
E>,. 0.015 lJ x= 0.04293 m static temperature of 118:1 K, and a static pressure of
A x= 0.08110 m ;_
100405 Pa. The calculated equilibrium mole fractions
x= 0.12080 m
of the species present in the vitiated air, determined
0.01
C) x: 0.1S080 m _
by a quasi-one-dimensional nozzle code. are given in
Table 1.
0.005
The initial channel cross-section is ,38.6 mm high
and 87.9 mm wide. The hydrogen fuel injectror in-
o_ 0.2 0.4 0.6 0.8--_" i trroduces hydrogen through a choked nozzle al Mach
Pitot Pressure Ratio 2.5, a static temperature of 1,34.2 I,L and a static pres-
sure of 201:300 Pa. The pilot fuel injectors described
Figure 8: Comparison of pitot pressure data with earlier were activated to improve flameholding un-
simulation results der the present test conditions. The pilot, injectors
are assumed to be choked at the wall surface, result-
ing in a static temperature and pressure of 2,51.7 I(
and 722535 Pa, respectively. No slip conditions were
specified along the upper, lower, and near- and far-
side channel walls. First order extrapolation was used
at the supersonic outflow station located at the 546
mm station for this calculation. This domain was dis-
cret.ized with a grid of 401 points in the st.reamwise
0.014 direction, 61 points in the cross-stream direction, and
121 points in the spanwise direction. The grid was
0.012 x: 0.04198 m compressed near the solid walls and the fuel injector.
x: 0.08202 m
Turbulence was modeled in the near wall region using
- x=0.12207 m
001 ii ---- xx,=00..0105200030 mm the Bauldwin-Lomax model, and in the interior field
'7 x: 0.00200 m
x=0.04200 m using the turbulent jet. mixing model .ls, l_J Chem-
_ _ x: 0.08200 m
E 0.008 _\ @ _ x=o.123oo., istry was modeled using the 9 species, 18 reaction
c_,_'_kE_2_,_ (':' X= 0.15300 m model described in reference.It This model provides
0,006 a detailed description of hydrogen-air chemistry, but
does not consider the effects of the small quantities
vv " ",_.,<) r,_ _-.- -- -v
of oxides of nitrogen and hydrocarbon species present.
_.',_.- -A ^ cdv in the vitiated air.
0.004 I
o0o;L
The results of flowtield simulations of the
SCHOLAR combustor model are shown in Figures
10o17. Figure 10 shows static pressure contours
400 600 800 1000 1200
along the stremnwise plane (:entered on the fuel jet.
Streamwise velocity, m/s
Traversing the combustor from inflow to outfow, a
weak bow shock produced by the pilot injectors can
Figure 9: Comparison of RELIEF velocity data with
simulation results be seen. This pressure rise is conmmnicated through
the wall boundary layer resulting in a weak shock at.
the inflow to the combustor. This is followed by the
expansion of the flow over the lower wall step. Just
downstream, the flow is compressed through a re(ore-
8
American Institute of Astronautics and Aeronautics