Instrumented and optically-accessible rocket chambers are being developed to be used for diagnostics of small rocket (less than 440 N thrust level) flowfields. These chambers are being tested to gather local fluid dynamic and thermodynamic flowfield data over a range of test conditions. This flowfield database is being used to better understand mixing and heat transfer phenomena in small rockets, influence the numerical modeling of small rocket flowfields, and characterize small rocket components. The diagnostic chamber designs include: a chamber design for gathering wall temperature profiles to be used as boundary conditions in a finite element heat flux model; a chamber design for gathering inner wall temperature and static pressure profiles; and optically-accessible chamber designs, to be used with a suite of laser-based diagnostics for gathering local species concentration, temperature, density, and velocity profiles. These chambers were run with gaseous hydrogen/gaseous oxygen (GH2/GO2) propellants, while subsequent versions will be run on liquid oxygen/hydrocarbon (LOX/HC) propellants. The purpose, design, and initial test results of these small rocket flowfield diagnostic chambers are summarized

Morren, Sybil

Reed, Brian

English

NASA Technical Reports Server

tN94-25030
SMALL ROCKET FLOWFIELD DIAGNOSTIC CHAMBERS
Sybii Morren and Brian Reed
NASA Lewis Research Center
SUMMARY
Instrumented and optically-accessible rocket chambers are being developed to be used for diagnostics of
small rocket (< 440 N thrust level) flowfields. These chambers are being tested to gather local fluid dynamic and
thermodynamic flowfield data over a range of test conditions. This flowfield database is being used to better
understand mixing and heat transfer phenomena in small rockets, influence the numerical modeling of small rocket
flowfields, and characterize small rocket components. The diagnostic chamber designs include: 1) a chamber design
for gathering wall temperature profiles to be used as boundary conditions in a finite element heat flux model; 2) a
chamber design for gathering inner wall temperature and static pressure profiles; and 3) optically-accessible chamber
designs, to be used with a suite of laser-based diagnostics for gathering local species concentration, temperature,
density, and velocity profiles. These chambers were run with gaseous hydrogen/gaseous oxygen (GH2/GO2)
propellants, while subsequent versions will be run on liquid oxygen/hydrocarbon (LOX/HC) propellants. This
presentation summarizes the purpose, design, and initial test results of these small rocket flowfield diagnostic
chambers.
HEAT FLUX DIAGNOSTIC CHAMBER
A chamber was designed and fabricated to determine steady-state heat flux to the inner wall to be used as a
diagnostic of the flowfield behavior, l The heat flux chamber design is an instrumented, water-cooled chamber used
to gather steady-state wall temperature profiles from measured and interpolated thermocouple data. These temperature
profiles are used as boundary conditions in a finite element analysis program, MSC/NASTRAN, to calculate the
local radial and axial heat fluxes in the chamber. Normal heat flux down the length of the chamber is then calculated
as the dot product of the total heat flux and normal vectors. It is important to emphasize that heat flux is used here
as a flowfield diagnostic - the objective is not to find absolute heat flux values (which would change for differing
wall conditions), but rather to use heat flux as an indicator of fundamental characteristics of the flowfield over a range
of conditons.
A schematic of a heat flux chamber used with a GH2/GO2 injector is shown in figure 1. The chamber was
fabricated from Oxygen-Free, High-Conductivity (OFHC) copper, with a 1.27 cm thick wall. An OFHC copper
housing was welded onto the chamber to provide an annulus for water cooling along the outer wall. Chromel-
alumei, grounded-junction thermocouples were embedded in the inner wall, nominally 0.159 cm from the inner wall
and attached to the outer wall and exit face. It was found that boundary condition definition was critical to the
accuracy and resolution of the heat flux profiles generated by the finite element program. Details of the chamber
fabrication and the heat flux methodology are given in reference 1.
This heat flux chamber was used to gather temperature data over a range of mixture ratios at 61 percent fuel
film cooling (FFC) and 414 kPa chamber pressure. The normal heat flux profiles (as a function of axial position)
generated from the temperature data is shown in figure 2. The heat flux profiles showed a significant increase in
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https://ntrs.nasa.gov/search.jsp?R=19940018557 2020-06-16T15:54:37+00:00Z
shearlayermixinginthebarrelsectionofthechamberformixtureratiosabovefive,indicatedbythesharpheatflux
peaksthere.Thisincreasewasfeltnottobearesultoftheflowtransitioningtoturbulent,sinceReynoldsnumbers
werewellwithinthelaminarregime,butratherduetoincreasedchemicalreactionsintheshearlayerbetweenthe
hydrogenfilmcoolingandoxidizer-richcoreflows.Itwasuncertainwhathetriggermechanismwasforthis
increasedshearlayermixingathighermixtureratios.Thisheatfluxmethodologywillbeusedfurtherfor
investigationfflowfieldbehaviorandcharacterizationofLOX/HCsmallrocketinjectorconfigurations.
TEMPERATURE/PRESSUREDIAGNOSTIC CHAMBER
A chamber was designed to investigate small rocket flowfields via inner wall temperature and static pressure
measurements,2 as shown in figure 3. The chamber liner was fabricated from OFHC copper, with milled channels
on the back side to provide water cooling. An OFHC copper outer housing was split into two parts axially along its
length and slid over the chamber liner in a clam-shell fashion. The housing was joined to the liner at the inlet and
outlet water manifolds, and sealed along the two axial housing seams. The chamber contained 4 axial rows of
thermocouple ports which spanned from the combustion chamber to the nozzle section. The chromel-alumel
thermocouples were located 0.076 cm from the hot gas side wall. The chamber also contained one row of static
pressure ports. Because the backside of the chamber was not sealed to the outer housing, instrumentation seals were
developed to accommodate the chamber cooling system.
The diagnostic chamber was hot fired using two platelet stack injectors of two, slightly different designs,
designated SN 02 and SN 03. The chamber was operated at 262 kPa, 379 kPa, and 517 kPa chamber pressures and
60% and 75% FFC, for mixture ratios between 4.0 and 8.0. At 75% FFC, both chambers exhibited the same
performance and thermal trends. However, at 60% FFC the data trends differed between the two injectors. Figure 4
show the axial inner wall temperature profiles for both injectors at 60% FFC. The inner wall temperatures for the
SN02 injector displayed a greater sensitivity to mixture ratio than for the SN03 injector. Furthermore, the inner
wall temperature generally increased with increasing mixture ratio for the SN02 injector, while the opposite trend
was observed for the SN03 injector. The results of this testing indicated that small rocket flowfield behavior may be
very sensitive to minor changes in operating conditons and injector design. Further discussion of tests results and
the details of the chamber design and fabrication are found in Reference 2.
This diagnostic chamber design will be used further for investigation of flowfield behavior (in particular, an
unsteady fiowfield phenomenon suspected to be associated with the GH2/GO2 injectors used previously),
characterization of LOX/HC small rocket injector configurations, and investigation of low Reynolds number nozzle
design issues. Furthermore, the pressure ports of this chamber could be utilize for gas sampling and for the insertion
of fiber optics for laser-based diagnostics.
OPTICALLY-ACCESSIBLE CHAMBERS FOR LASER-BASED DIAGNOSTICS
Optical access to the combustion chamber allows the use of non-intrusive, laser based diagnostics to gather
local fluid dyamic and thermodynamic data. These diagnostics include Raman spectroscopy to measure species
concentrations, gas temperatures, and flow velocity profiles near the injector and Laser-Induced Fluorescence (LIF)
used to visualize the shear layer mixing process and to measure minor species concentrations.3 Schlieren and
6
shadowgraphtechniquescouldalsobeusedwithopticallyaccessiblechamberstovisualizetheflowfield.A square
rocketchamberwasdesigned,fabricated,andhotfiredtoprovideopticalaccesstothecombustionchamberfromthe
sides.Thesquare(2D),optically-accessiblechamberwaseenasanintermediates ptowardanaxisymmetric(3D),
optically-accessiblechamberandwasthereforedesignedtoberobust.
The2Dchamberisshownin figure5. ThechamberwasmachinedoutofOFHCcopperwithwatercooling
channelsdrilledinthewallaboveandbelowthecombustionchamberandnozzle.Ontheleftandrightsideofthe
chambertherewereslotsfortheplacementof 13.3cmby6.35cmby1.27cmwindows.Highopticalquality,fused
silicawasselectedasthewindowmaterial.Fusedsilicahasamaximumoperatingtemperatureofonly980°C,but
hasarelativelylowcoefficientofthermalexpansioncomparedtosapphire.Furthermorefusedsilicacanbe
machinedwithbettersurfacequalitiescomparedtosapphire,reducingtheamountofdiffusescatteringof laser light
within the chamber. Gaseous nitrogen was injected against the windows for cooling.
The 2D chamber provides optical access down the entire length of the rocket. Because the windows are
cooled independently, the interaction of fuel film cooling flow with the oxidizer-rich core flow in the shear layer (a
critical phenomenon in small rocket flows) can be studied without the threat of a film cooling layer breakdown
melting the windows. The main drawback, however, is that the measurements - and any conclusions drawn from
them - are restricted to a two-dimensional flowfield. It will be important to distinguish between effects that are
representative of the flowfield behavior and effects that result from the two dimensionality of the chamber. Also the
nitrogen flow, though directed against the windows, will slowly penetrate into the hydrogen/oxygen flow.
The 2D chamber was checked out using copper slabs instrumented with thermocouples in place of the
windows. The nitrogen flow was sufficient to keep the inner wall temperature below 93 "C, with no more than a 27
°C temperature variation across the slab. The chamber was then run with the windows in a series of 15-second
duration tests at a chamber pressure of 296 kPa and a total mass flowrate of 0.033 kg/s. At an overall mixture ratio
6 and 55% FFC, the core mixture ratio was 13.3. After a total of 232 tests with the same set of windows, there
was no evidence of cracking or degradation of the windows. Laser Raman spectroscopy was used to measure the gas
temperature and oxygen concentration at the exit of the chamber sleeve insert used with the injector. The measured
temperature and oxygen concentration profiles were found to be significantly different from the profiles that are
usually assumed in numerical calculations. Detailed discussion of the Raman measurements with this chamber can
be found in reference 4.
The next step from the 2D chamber will be an axisymmetric, optically-accessible rocket chamber. The
preliminary design of a 3D chamber to be used with GH2/GO2 propellants is shown in figure 6. A modular design
is envisioned with a flush-mounted injector, a cylindrical, fused silica window for the barrel section, and a water-
cooled converging/diverging nozzle, all clamped together. A concern with this design are the thermal stresses that
could be induced by temperature differences between the window mating surfaces. Fuel film cooling from the
injector and water cooling of the nozzle section should keep the mating surfaces at a fairly uniform temperature.
Water cooling of the nozzle section should also minimize thermal soakback into the window, although a low
thermal conductivity interface could be used between the window and nozzle if water cooling is not sufficient.
As with the 2D chamber, the 3D chamber will allow the used of Raman spectroscopy, LIF, Schlieren, and
shadowgraph diagnostics, but optical access will only be provided in the head end, relatively low temperature region
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of thebarrelsection.It is thoughtthatmuchoftheshearlayereactionsprobablyoccursin thisregion,however.
Coolingofthewindowwillbebythehydrogenfilmcoolinglayeronly,socaremustbetakenthatthewindow
sectiondoesnotextendtoofarintothechambero thathechamberisnotrunatconditionswherethefilmlayer
willbreakdownquickly.Resultsfromthe2Dchambershouldprovideguidancein thisdesignissue.Laterversions
ofthe3DchambercouldincludeafusedsilicanozzlextensiontostudylowReynoldsnumberflowinthenozzle
andwindowslotsinthethroatregion.
REFERENCES
1. Reed, B.D., "Small Hydrogen/Oxygen Rocket Flowfield Behavior From Heat Flux Measurements", AIAA Paper
93-2162, NASA TM-106233, June 1993.
2. Morren, S., et. al., "A Laboratory Model of a Hydrogen/Oxygen Engine for Combustion and Nozzle Studies",
AIAA Paper 93-1825, NASA TM-106281, June 1993.
3. Schneider, SJ., "Low Thrust Chemical Rocket Technology", IAF Paper 92-0669, NASA TM-105927, August
1992.
4. DeGroot, W.A., "Gaseous Hydrogen/Oxygen Injector Performance Characterization", AIAA Paper 94-0220, to
be presented at 1994 AIAA Aerospace Sciences Meeting.
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9
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10
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Small rocket flowfield diagnostic chambers

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