Cooling characteristics of a transpiration-cooled afterburner with a porous wall of brazed and rolled wire cloth

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Material Information

Title:
Cooling characteristics of a transpiration-cooled afterburner with a porous wall of brazed and rolled wire cloth
Series Title:
NACA RM
Physical Description:
68 p. : ill. ; 28 cm.
Language:
English
Creator:
Koffel, William K
Lewis Research Center
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Afterburners -- Research   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: Cooling data from this experimental afterburner are successfully correlated over a range of Reynolds number based on the distance downstream of the leading edge of the porous wall. Calculations based on the correlation indicate that the combustion-chamber cooling-air requirements for an exhaust-gas temperature of 3700° R are only 16 percent of the air required for conventional forced-convection cooling. The cooling-air requirements were nearly independent of flight condition, but close control of cooling-air pressure is required at high flight speeds.
Bibliography:
Includes bibliographic references (p. 21-22).
Statement of Responsibility:
by William K. Koffel.
General Note:
"Report date June 11, 1954."
General Note:
"Classification changed to unclassified Authority: NACA Research Abstract No. 127 Effective date: May 16, 1958."--stamped on cover

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808495
oclc - 129591745
sobekcm - AA00006155_00001
System ID:
AA00006155:00001

Full Text

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RM E54EZ5


RESEARCH


MEMORANDUM


COOLING CHARACTERISTICS OF A TRANSPIRATION-COOLED
r*I
AFTERBURNER WITH A POROUS WALL OF BRAZE )

AND ROLLED WIRE CLOTH

By William K. Koffel
Lewis Flight Propulsion Laboratory
Cleveland, Ohio
UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE UBRARY .
P.O. BOX 117011 fi
GAINESVILLE, FL 32611-7011 USA
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I NATIONAL ADVISORY
FOR AERONAL
WASHINGTON
:.. August 19, 1954

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NACA EM E54E25


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


RESEARCH MEMORANDUM


COOLING CHARACTERISTICS OF A TRANSPIRATION-COOLED AFTERBURNER WITH A

POROUS WALL OF BRAZED AND ROLLED WIRE CLOTH

By William K. Koffel


SUMMARY

Cooling data were obtained for a transpiration-cooled afterburner
having a porous combustion-chamber wall of brazed and rolled wire cloth
for a range of exhaust-gas temperatures from 12000 to 33400 R, total
flow ratio of cooling air to combustion gas of 0.025 to 0.106, and
pressure altitudes of 15,000 to 45,000 feet. The data are successfully
correlated over a range of Reynolds number from 75,000 to 1,500,000,
based on the distance downstream of the leading edge of the porous wall.

Maximum wall temperatures based on the cooling correlation were
determined for a porous wall of uniform permeability at sea-level take-
off and for flight Mach numbers of 0.8, 1.5, and 2.0 at an altitude of
35,000 feet. The cooling-air requirements were nearly independent of
the flight conditions. A total flow ratio of cooling air to combustion
gas of about 0.032 can maintain a maximum wall temperature of 12100 R
with an exhaust-gas temperature of 37000 R. The total flow ratios with
a uniform permeability distribution and air flows sufficient to limit
the maximum wall temperatures to 12100 R are about 15 percent higher than
the minimum total flow ratios corresponding to a variable-permeability
wall with a constant wall temperature of 12100 R. The total flow ratios
for a maximum wall temperature of 12100 R with a wire-cloth afterburner
are about 16 percent of the total flow ratios required to cool a
stainless-steel afterburner wall convectively to a maximum wall tempera-
ture of 17600 R with exhaust-gas temperatures of 32000 and 37000 R.
The analysis of cooling-air requirements for the previously mentioned
flight conditions indicated that cooling-air static pressures must be
closely controlled, especially at a flight Mach number of 2.0.


INTRODUCTION

The satisfactory cooling of high-thrust-augmentation afterburners
for supersonic aircraft requires large amounts of cooling air. These
large quantities of cooling air, whether supplied from the ambient-air


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NACA RM E54E25


stream or from compressor bleed, involve large losses in net thrust.
These losses, however, can be minimized through the use of more effective
cooling methods. Transpiration cooling has been shown theoretically
(ref. 1) to be more effective than film cooling or the conventional
method of afterburner cooling by forced convection. The high effective-
ness of transpiration cooling was experimentally verified in a preliminary
investigation of a transpiration-cooled afterburner having a wall of
sintered porous stainless steel (ref. 2). The commercial sintered porous
stainless steel used in the preliminary investigation of transpiration
cooling for afterburners was unsuitable for use in flight afterburners
because of low strength-to-weight ratio, poor control over the uniformity
of permeability, and fabrication difficulties.

In the search for more suitable porous materials, wire cloth was
investigated (ref. 3) because of its high strength-to-weight ratio,
availability in large sheets, ease of fabrication, and low cost compared
with sintered porous metals. The NACA Lewis laboratory designed and
built a transpiration-cooled afterburner with a porous wall of the
brazed and rolled-wire cloth, described in reference 3. This burner was
operated on a turbojet engine to determine the feasibility of a wire-cloth
transpiration-cooled afterburner and to evaluate its cooling performance.

The cooling data from the wire-cloth afterburner are reported herein
for pressure altitudes of 15,000 to 45,000 feet, exhaust-gas temperatures
from 12000 to 33400 R, and total flow ratios of cooling air to combustion
gas of 0.025 to 0.106. A cooling correlation is derived from these data.
This correlation is then used to predict the cooling-air requirements of
this afterburner with exhaust-gas temperatures of 32000 and 37000 R for
several typical flight conditions of a supersonic airplane. A comparison
is also made between the cooling-air requirements of this afterburner
with transpiration cooling and with conventional forced-convection cooling.
The experimental cooling correlation is briefly compared with the
approximate theory of Rannie and Friedman refss. 4 and 5) for transpi-
ration cooling with turbulent boundary-layer flow.


APPARATUS AND IilJTRL TEllT''T'TON

Test Installation

The wire-cloth afterburner was mounted on a conventional axial-flow
turbojet engine installed in an altitude test chamber. The test installa-
tion aft of the turbine flange is shown in figure 1. The cr,_ine and
afterburn-tr assembly were mounted on a bed plate suspended in the altitude
chamber in such a manner as to permit the measurement of jet thrust. A
sectional view of the afterburner is shown in figure 2 together with
details of the flame holder and lu..-1-spray bars. The configuration between


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the turbine flange and the beginning of the porous wall, including the
fuel-spray bars and flame holder, was identical with the high-performance
afterburner configuration of reference 6, which was evolved from con-
figuration C of reference 7.

Cooling air from an independent source was filtered through 100
square feet of 1-inch-thick Fiberglas filter media that retained particles
greater than 1 micron. The flow was measured by means of an A.S.M.E.
standard thin-plate orifice. The plenum chamber and the high-temperature
impermeable wall upstream of the porous-wall section were Inconel, and
the cooling-air side of the mild steel shroud was painted with zinc
chromate primer in order to eliminate the possibility of clogging the
wire cloth with fine scale.

The engine fuel was clear unleaded gasoline of 62-octane rating,
and the afterburner fuel was MIL-F-5624A, grade JP-4.


Wire-Cloth Porous Wall

The porous combustion-chamber wall was made from wire cloth that
was brazed and rolled, for permeability reasons, as described in reference
3. The particular cloth used was monel 21x70 twilled Dutch weave
(designated cloth B in ref. 3) sprayed with three coats of silver solder
per side and brazed and rolled to a 35-percent reduction in original
thickness. The average final thickness was 0.0274 inch, and the average
permeability coefficient K was about x10-8 square inch. The tapered
combustion chamber was lined with a porous wall made from 20 pieces of
wire cloth formed into shallow channels. Eacht channel was 1/2 inch deep
by about 41 inches long and tapered in width from about 4 inches at the
upstream end to about 5.5 inches at the downstream end. The channels
were spot-welded to angles (fig. 3) that were fastened to the structural
shroud by blind rivets. The permeability of each channel differed
slightly from the average permeability for all the channels. It was
expected that the cooling-air static pressure would vary somewhat
circumferentially because of the tangential inlet to the plenum chamber.
Therefore, the channels were arraii:ed in an order that would tend to
produce a circumferentially uniform distribution of coolin i air.

The newly fabricated porous combustion-chamber wall is shown in
figure 4(a). The channels of wire cloth were permanently bulged toward
the center line of the combustion chamber after the initial operation of
the engine. Figure 4(b) shows the bul-ed channels at the end of the
c-olinr-, investigation with c-olir:-p tas:--: hPiji:ts of about 3/4 to 7/8
inch in midchannel. The method of suspending the wire cloth caused a
minimum of disturbance to the cooling-air film on the gas side of the
porous wall, and the bulirn,_ of the channels re-du;e tensile stresses
in the wire cloth caused by pressure forces. The impermeable wall


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upstream of the porous wall was forced-convection cooled by the cooling
air before it flowed into the 20 dead-ended cooling-air passages.


Instrumentation

The research instrumentation for measurement of temperatures and
pressures at several stations thro:.u.h the turbojet engine was the same
as in reference 6. The location of each temperature and pressure
measurement on the afterburner is shown in figure 5. The five principal
stations of instrumentation on the porous wall were 24, 32.6, 41, 49.6,
and 57.5 inches downstream of the quick-disconnect fl'iiige at the
combustion-chamber inlet (see fig. 2).

Details of a typical group of instrumentation are shown in figure 6.
The cc li'.*-air temperature probes had butt-welded iron-constantan
thermocouple junctions shaved smooth to 0.010-inch diameter. The
thermocouples on the wire cloth were 28-gage (0.013-in.) chromel-alumel
wire individually wrapped with silicone-treated asbestos insulation and
overbraided with Fiberglas insulation. The orientation and application
of the spot-welded junctions on both sides of the wire cloth can be seen
in fi -ure 6. There were only four thermocouples on the -a, side of the
wire cloth because of difficulties in installation. Trie shroud tempera-
tures were measured with 22- _-L._ (0.025-in.) iron-constantan thermo-
couples (not shown in fig. 6) spot-welded to the outside of the shroud.
The longitudinal profile of static pressure in the combustion chamber
was determined by means of 0.080-inch outside diameter by 0.010-inch
wall stainless-steel tubes sweat-brazed into the wire cloth and ground
flush. The flexibility of these tubes and the method of 1:.adi.. them
out of the p.::.- permitted the wire cloth to assume a natural curve
with minimum distortion of the cloth at the point of pressure measurement.

Strain-gage pressure pickups were mounted on the co~11-i -air plenum
chamber and on the combustion-chamber wall a:. the flame-holder station
to obtain the static-pressure pl.l'.:iit', in the cooling air and combustion
chamber during ignition and normal afterburning. The pressure pulsations
were recorded on a two-channel oscillograph.


TE:'T F:R'CEDlURE

Air-flow calibrations were obtained for each piece of wire cloth
before fabrication into the porous wall. These calibrations were taken
at points where temperature instrumentation was later applied. Additional
cold-air-flow calibrations were made of the assembled porous wall, as a
whole, at several ambient pressure levels before and duri the cooling
i vesti,_' ion to detect any over-all -1.-. i, in the air-flow calibrations
from plu ginE by fuel or combustion residues or from oxidation of the


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NACA EM E54E25


brazing alloy. Nonafterburning check points were also made during the
cooling investigation as a check on instrumentation and changes in the
wire cloth.


The data are presented in tables
of some of the variables are given in
cooling-air temperature of 5350 R:


I to V, and the approximate range
the following table for an inlet


Pressure Compressor Cooling-air Exhaust- Type of test
altitude, air flow, flow, Wa, gas total
ft lb/sec lb/sec temperature,
T
Tg,2,
OR

Sea level 0 0.970-5.810 a535 Check on air-
15,000 0 1.255-4.344 a535 flow calibration
35,000 0 .903-2.849 a535 and plu Ofiilg of
45,000 0 1.49-1.79 a535 wire cloth

15,000 45.8 1.2-3.8 1200 Cooling, non-
35,000 25.5 .83-2.7 1250 afterburning

15,000 45.8 1.37-3.25 2580-3010 Cooling,
35,000 25.5 1.26-3.16 1800-3340 afterburning
45,000 15.6 1.49-1.79 3185
a
Approximate ambient-air temperature.

The cooling-air flow was varied at a given pressure altitude and
fuel-air ratio between the following limits:

(1) Minimum static-pressure difference Ap across upstream edge of
wire cloth of 0.5 lb/sq in. to prevent reverse flow of
combustion gas tIu-L:i,_h wire cloth

(2) Maximum Ap across downstream ed e of wire cloth of 6 lb/sq in.
to avoid bursting channels of wire cloth

(3) Maximum wire-cloth temperature of 14600 R to avoid excessive
rate of plugging from oxidation of silver solder

(4) Maximum turbine-outlet gas temperature of 16250 R to protect
turbine from over-temperature


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METHODS OF CALCULATION AND ANALYSIS

Cooling-Air Flow

Local values of the cooling-air flow normal to a unit of wire-cloth
surface area (pV)a were computed for two channels, channel 1 having the
hottest and'channel 9 having the coldest porous-wall temperatures. This
computation was based on the air-flow calibrations of figure 7, the
thickness of each channel, the respective longitudinal profiles of
cooling-air static pressure, and the assumption that the l.'n._:itudinal
profile of combustion-gas static pressure was circumferentially uniform.
2
(Symbols are defined in appendix A.) The ordinate (2) of
figure 7 is in effect a measure of the pressure drop per unit thickness

across the porous wall, and the abscissa (pV)a is the reduced weight
flow of cooling air per unit of wire-cloth surface area. The viscosity
and temperature ratios, po/i and (o /[)2 To/T, reduce the air-flow data
to NACA standard temperature. A theoretical analysis in reference 8
indicates that the profile of air temperature across a porous wall
should practically coincide with profile of metal temperature except for
a very short distance on the side the air enters the wall. Consequently,
the viscosity- and temperature-ratio factors should be evaluated at the
average local wall temperature. Measured wall temperatures on either
side of the wire cloth in this afterburner were within 200 R of each
other, so for convenience the factors were evaluated at the temperature
measured by the air-side thermocouples, which were installed at each
station of channels 1 and 9.

The calculated Lotal cooling-air flows in a later section, Coc.li:,-
Air Requirements Calculated fr'-m C-..li i Correlation, were obtained from
the step-by-step summation of AD(pV)aAx.


Exhaust-Gas 'i._,j_. rature

7,.- total temperature of the exhaust :' was calculated from the
exhaust-nozzle total pressure P6S, measured jet thrust, velocity
c f .:icient, and jas flow by the fll equation (ref. 7):


.. -) ( L (1)
i, 2


L .. AL


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NACA RM E54E25


The values of R were computed according to the method of reference 9,
and the jet velocity coefficient was determined to be 0.97 from data
presented in reference 10.


Local Combustion-Gas Temperatures

Local values of the combustion-gas total temperature T,x with
afterburning were computed from the empirical equation (ref. 11)

T 1 T 7 X
Tgx gl xsi x
T -T 2 L (2)
g,2 g,l

where T gl is assumed equal to the turbine-discharge gas temperature,

and Tg,2 is the exhaust-gas total temperature. Equation (2) makes
no allowance for p..:.sible hot streaks and implies a uniform transverse
profile of combusti'._n- -. temperature. Hot streaks undoubtedly were
present, although observation of the combustion color and pattern by
means of a periscope directed up the burner axis showed good circumfer-
ential uniformity. Tl.-- transverse temperature profile was made as
uniform as possible in a previous investigation of the radial fuel-air
distribution when this same confi -ir'::ion had a convection-cooled
impermeable wall.

For nonafterburning cooling data, local values of bulk total
temperature Tx were assumed equal to the turbine-discharge gas
mp:-r-.-i.re Tg,l.


Correlation of Cooling Data

Cooling data from five stations along the two typical channels of
wire cloth were correlated by plotting the temperature-difference ratio
(Tw T)/(Tg Ta) against the coolant-flow ratio (pV)a/(pU)g for
constant values of bulk Reynolds number. The flow over the porous wall
was believed to be similar to the boundary-layer flow over a flat plate,
so that the length used in the Reynolds number was taken as the distance
from the leading edge of the porous wall. Consequently, the temperature-
difference ratios and the coolant-flow ratios are local values corre-
sponding to the local Reynolds numbers. The viscosity of the combustion
,3 was assumed to be the same as for air given in table 2 of reference
12 and was evaluated at the film t:rr~:i-.rut--re Tf. The quantity (pU)
was assumed identical to the local one-dimensional value of total weight
flow per unit of combustion-chamber flow area.


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The wall temperature T. in the temperature-difference ratio
(Tw Ta)/(Tg Ta) was measured on the cooling-air side of the wire
cloth but can be considered practically to represent the wall tempera-
ture on the hot-gas side because of the small difference in measured
temperatures across the wire cloth in this afterburner.


Cooling-Air Requirements Calculated from Cooling Correlation

The experimental cooling correlation of this investigation was used
to calculate the longitudinal profile of porous-wall temperature in this
afterburner. The conditions analyzed are given in the following table:

Flight Flight Altitude, Exhaust-gas
condition Mach ft total
number temperature, T 2
g,2,
oR

A 0 Sea level 3700
B .8 35,000 3700
B .8 35,000 3200
C 1.5 35,000 3700
D 2.0 35,000 3700

The turbine-dischar;e temperature was 17100 R for all conditions. The
cooling-air static pressure and the permeability of the porous wall were
assumed to be absolutely uniform, and the air-flow calibration of the
porous wall was assumed to be identical with the mean calibration
curve of figure 7. With a uniform permeability, most of the porous wall
will be overcooled except for some peak temperature r- -:ion, which must
not exceed the maximum allowable operating temperature of the porous
material. The calculated maximum wall temperatures were plotted against
total flow ratio for each flight condition. It should be kept in mind
that wall temperatures calculated from the cooling correlation are
assumed to be uniform around the circumference at each station. Usually,
a safety factor must be allowed for random hot spots in the porous wall
caused by hot streaks in the combustion gas or by local areas that have
less than the mean permeability.

Minimum cooling-air flow. The minimum cooling -air flow for a
given temperature limit results when the wall temperature is m:in-tained
constant at the maximum allowable operating temperature of the porous
material by use of an infinitely variable permeability distribution
along the porous wall. The .:. :tion of a constant wall temperature
makes it possible to compute the minimum i _,, l -air distribution '
the minimum total 'low ratio directly from the c( i correlation
without cousiderji he di sributions of permeability that would be


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NACA RM E54E25


required to provide the minimum cooling-air distribution. Although it
is not practical to vary the ierrme-.bility distribution for each flight
and operating condition in an actual afterburner, the minimum flow
ratios are useful for reference and for examination of trends. Minimum
total flow ratios were therefore also computed for flight conditions A to
D with a constant porous-wall temperature of 12100 R.

Longitudinal profile of combustion-gas static pressure. The static-
pressure profile along the length of the combustion chamber was
calculated step-b'-step from well-known one-dimensional flow relations
for momentum pressure drop in a constant-area duct and the isentropic
variation of Mach number with area ratio. The friction pressure drop,
which was assumed to be small relative to the momentum pressure loss,
was neglected.

Cooliu-,;-air prr.sur:.e aid temp-ratlire. The cooling air was assumed
to have been bled from an interstate of the compressor (at a compressor
efficiency of 0.85) at a pressure high er-,..lh to allow for a pressure
drop of 1 pound per square inch through valves and ductin to the
upstream edge of the porous wall. The pressure drop across the porous
wall was limited to a minimum of 1 pound per square inch for practical
control, and to a maximum of 6 pounds per square inch to prevent
bursting the channels of wire cloth. The cooling-air pre.ssur- and
temperature were presumed to be uniform in all cooling-air passages.
Convective heat transfer to the cooling air by the imperm.raLle wall
upstream of the porous wall was assumed to raise the air temperature
50 above that corresponding to the comiaprcssc.r bleed temperature.


Cooling-Air Requirements with F.-'rced-Convection Cooling

Cooling-air requirements for forced-convection c:ollir: of this
afterburner with ram air were calculated for exhaust-gas temperatures
of 32000 and 37000 R for fli -t: condition B from data in reference 13
and the cooling correlation of reference 11.


F:REiJLTS AND rI.CULS5IC!i

T, pical Data

Figure 8 shows typical measured circumferential r.: files of the
cocli --.--air static pressure and wire-cloth temperature. The static-
pressure profiles shown result from the t i_: I inlet on the plenum
chamber and the previously mentioned order of a.I-r i:: l the channels
i' wire cloth. As would be expected, the temperature profiles are the
reverse of the pressure rrifiles. Figure 8 shows that channel 1 was
-._rally the hottest and channel 9 the coldest, with or without after-
burning.


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Typical longitudinal profiles for the wire-cloth afterburner are
shown in figure 9. The importance of eliminating circumferential gradients
in cooling-air static pressure and in perme'bility is obvious from the
5300 R variation in wall temperature at station 62 in figure 9(a). Part of
the circumferential temperature spread in this afterburner would have been
averted if there had been no circumferential pressure gradient in the
cooling air. The pressure gradient could have been reduced by a better
plenum-chamber inlet or by cross-flow holes in the angles that supported
the channels of wire cloth. Wire cloth more closely meeting the specified
permeability could have practically eliminated circumferential variations
in wall temperature, except for those caused by any hot streaks in the
combustion gas. The slight drop in cooling-air temperature (figs. 9(a)
and (c)) along the length of the cooling-air passage is due to heat
losses through the uninsulated shroud.

The longitudinal profiles of static pressures on both sides of the
porous wall are shown at the top of figures 9(b) and (d). For the
relatively high flow ratios shown (Wa/Wg of 0.0716 to 0.0951), the
continuous bleeding of air through the porous wall caused a deceleration
of the cooling-air flow and a corresponding rise in static pressure of
the cooling air in the flow direction. With the low flow ratios of
practical interest, the static pressure of the cooling air was almost
independent of passage length.

The variation of pressure drop across the porous wall with distance
along the wall is represented by the vertical distance between the static
pressures of the cooling air and of the combustion gas. The resulting
longitudinal profiles of coolant-flow ratio (PV)a/(pU)g are shown in
figures 9(b) and (d). The coolant-flow ratios for channel 9 were
considerably higher than for channel 1. The differences in coolant-flow
ratios are caused by the differences in permeability and the absolute
pressure level in each channel. It is recalled that the air-flow cali-
brations of figure 7 indicate (pV)a is a function of the difference of
the squares of the absolute static pressure on both sides of the porous
wall instead of the pressure drop alone. The curves of the fraction of
total cooling-air flow in each channel are almost linear and nearly
identical for channels 1 and 9 (figs. 9(b) and (d)). It can be inferred,
therefore, that the longitudinal addition of cooling air was almost linear
'ILroul. all the channels of wire cloth.

Before di .:u:in the coolir,. correlation, it is of interest to
determine whether there was any progressive change in the air-flow
calibration of the wire cloth duiri the cooling investigation. A dull
greasy inul. formed on the combustion-gas side of the wire cloth in
spite of the cooling-air film flowing away from the wire cloth during all
periods of en .;ii.- operation. The smudge is visible in figure 10 to the
left of the white lines. Th.- areas to the right of the white lines have


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NACA RM E54-E'-',







NACA RM ESIE'15


been wiped clean with cleansing tissue to show the condition of the
wire cloth after 4 hours 10 minutes of afterburning. The smudge caused
no discernible effect on air-flow calibration.

Figure 11 shows the circumferential average static-pressure drop
across the porous wall against orifice cooling-air flow for several
levels of combustion-chamber pressure at stations 24 and 57.5. The
agreement between data taken before and during the cooling investigation
was excellent at all stations. On the basis of figure 11 and the close
agreement between temperature-difference ratios for nonafterburning
check points, it is concluded that there was no significant change in
the air-flow calibration of the wire cloth during the cooling
investigation.


Cooling Correlation

The experimental c-olin.j data with afterburning are correlated over
a range of Reynolds numbers in fiui-res 12(a) to (f). Cooling data
obtained from channels 1 (hottest) and 9 (coldest) met, or overlapped,
to define a single curve when the Reynolds number was held approximately
constant. The mean curves drawn throuu)i groups of data points having
approximately the same Reynolds number are summarized in figure 12(g).
Within the ranges investigated, neither Reynolds number nor radiation
had any marked effect on the correlation of afterburning cooling data
for coolant-flow ratios less than about 0.007. Above a coolant-flow
ratio of 0.007, the temperature-difference ratio, and hence the porous-
wall temperature for fixed cooling-air and gas temperatures, decreased
as Reynolds number increased from about 75,000 to 800,000. The curves
for Reynolds numbers of 1,000,000 and 1,500,000 lie, respectively, above
and below the curve for 800,000. It is probable that in this Reynolds
number range, correspundiingr to stations 49.6 and 57.5, increases in
radiant heat transfer tend to counterbalance increases in the Reynolds
number (see ref. 1), although the radiant heat transfer was calculated
to be less than one-tenth of the convective heat transfer.

A direct comparison between the curves of the experimental cooling
correlation and those predicted by th!e apipru:-:imst. theory of references
4 and 5 is not entirely realistic because of the different assumptions
used with these data and those used for developing the approximate theory.
However, a partial comparison is made in appendix B.

Although nonafterburning conditions were not of primary importance
to the present investigation, the nonaft-rburrning conditions are of
somewhat general interest in connection with the heat-transfer process
of transpiration cooling. The combustion-gas temperature was almost
constant along the length of the combustion chamber d iri.n nonafter-
burning, and radiation from the gas was negligible because of the low


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NACA RM E54E25


gas temperature and pressures. The nonafterburning cooling data are
presented in tables I to IV. Nonafterburning cooling data are shown
in figure 13 by plotting (T, Ta)/(Tg Ta) against Reynolds number
for constant values of (PV)a/(pU)g. Cooling correlation curves for
Reynolds numbers of 105 and 106 (fig. 14) were obtained from a cross
plot of figure 13. The curves of figure 14 are similar in shape and
magnitude to the afterburning data for comparable Reynolds numbers,
as can be seen in figure 15. Although somewhat better cooling is in-
dicated by the rnoncf .erburning than by the afterburning data at coolant-
flow ratios less than about 0.007, the agreement between afterburning
and nonafterburning data is considered satisfactory. The agreement
probably would have been closer if measured values had been available
of the local combustion-:s temperatures and of (pU)g near the wall,
instead of assuming (pU)g = Wg/A. The profiles of (pU) are known to
differ between nonafterburning and afterburning conditions.


Transpiration-Cooling Performance

Figure 16 is a plot of the calculated maximum wall teniperature
against total flow ratio Wa/W for the assumed 'li ht conditions A
to D when operating with an exhaust-gas temperature of 37000 R. It
should be emphasized that the air flow thrcuji the porous wall is a
function of the difference of the squares of the absolute pressures on
either side of the wall (see fig. 7). Hence the magnitude and the
longitudinal position of the maximum wall temperature are functions of
the profiles of the absolute static pressures and of the combustion-gas
temperature. The profiles of absolute static pressure are dependent on
the flight condition and the pumping characteristics of the en iie used,
in addition to pressure losses in the afterburner. For the flight
conditions investigated, the maximum wall temperature occurred at the
leading edge of the porous wall, except when the cooli. -air static
pressure was reduced to 10.85 pounds per square inch absolute. At
this pressure, the maximum wall temperature moved to about station 49.6.

Separate curves of maximum wall temperature a._-iist total flow ratio
were obtained for each fl .]it condition (fig. 16). The curves closely
overlap so that the cooling-air requirements were nearly irLnepel- .rt of
flight conditions; however, the curves would be separated more if the
gas temperature had been widely varied. The operable rung- of total flow
ratio varied with flight condition because of the related minimum and max-
imum pr re drops across the porous wall. The upper symbol on each curve
corresponds to the assumed minimum pressure drop across the porous wall of
1 pound per square inch at the leading edge. The lower symbol corr:cprionds.
to the assumed maximum pressure drop of 6 pounds per sl.q-re inch at the
'lliit' t ;'. The resultl:j.; minimum and maximum cooling-air pressures are
given in figure 16. With a uniform distribution of coolili;s-air pressure,


",i.T TLEITIAL


CONFIDENTIAL







NACA RM E54E25


the range in static pressure for control of cooling-air flow at a given
flight condition is then 5 pounds per s quare inch minus the drop in
comtbustion-chamber static pressure along the pcr::'is wall. Consequently,
large drops in combustion-chamber pressure along the porous wall tend to
decrease the range of cooling-air static pressure for control of
cooling-air flow. The control range for cooling-air static pressure
varied from 3.70 to 0.66 pounds per square inch, respectively, for
flight conditions B and D. Cooling-air pressure must be accurately
controlled within this range. For example, at fli h:. condition D, a
decrease in cooling-air static pressure of 0.66 pound per square inch
(only 1.85 percent of the absolute coc.lirg,-air pressure) causes a 2560 R
increase in maximum wall temperature.

Greater range could be obtained in cooling-air static pressure by
using several layers of wire cloth. This type of construction would
permit higher pressure drops across the porous wall and would make pos-
sible the control of low flows corresp. ,diiing to high maximum wall temper-
atures without danger of reverse flow through the wall. The peak tempera-
ture limit for the porous wall was about 14100 R because of oxidation of
the brazing alloy on the wire cloth in this afterburner. Therefore, a
temperature of 12100 R is assumed to be safe for the maximum wall temper-
atures computed from the cooling correlation. This temperature provides
a 2000 safety factor for local peak temperatures caused by hot streaks in
the combustion gas and from small random areas of the cloth that have less
than the mean permeability. A total flow ratio Wa/Wg of about 0.032 can
maintain a maximum wall temperature of 12100 R with an exhaust-gas temper-
ature of 37000 R at conditions A to C. Calculations indicated that a
maximum wall temperature of 12100 R could also be maintained at condi-
tion D by increasing the assumed maximum allowable pressure drop across
the wire cloth by about 20 percent (point indicated by the end of the
dotted extension). A visual extrapolation of the curves of figure 16
(disregarding pressure limits) to a wire-cloth tem-peratur- of 17600 R,
which is a representative temperature for a wall of stainless steel, re-
sults in a total flow ratio of about 0.018. This value compares favor-
ably with the total flow ratio of 0.016 cp.-..piuted in reference 2 for the
same maximum wall -emperatu-e with a porous wall of sintered stainless
steel and an exhaust-gas temperature of about 38000 R. The somewhat more
effective cooling indicated for a porous wall of sintered stainless steel
in reference 2 may be caused by a more uniform film of cooling air on the
sintered wall than that produced by the fewer number of larger pores in
the surface of the wire cloth.

A total flow ratio of 0.032 for f'li jt conditions A to D when
using a uniform permeability distribution is about 15 percent higher
than the minimum total flow ratio for a constant wall -.emperature of
12100 R which would generally require a different permeability distri-
bution for each flight condition. Therefore, the practicability of


COI WIDEITIAL


COIFI DEFTIAL







NACA RM E54E25


using wire cloth having a nonuniform longitudinal distribution of
permeability, compromising the variable permeability distributions
required to maintain a constant wall temperature for each of the flight
conditions investigated, is questionable.

Some reductions should be expected in the total flow ratios of
transpiration-cooled afterburners if, through quality control, the toler-
ance on the uniformity of permeability can be decreased from that of the
brazed and rolled wire cloth used. The use of porous materials having
sufficient strength and resistance to oxidation at higher operating
temperatures will also permit a reduction in cooling-air flow. For
example, the minimum total flow ratio would decrease about 49 percent
if the maximum wall temperature could be increased from 12100 to 17600 R,
which may be possible for cloth woven from stainless steel or Inconel
wires. It is expected that problems due to oxidation of a braze alloy
can be solved by the use of high-temperature brazing alloys under
development, or by the substitution of sintering for brazitig.


Comparison of Transpiration and Forced-Convection Cooling

The cooling-air requirements for transpiration and forced-convection
cooling of this afterburner are compared in figure 17 for exhaust-gas
temperatures of 32000 and 37000 R at flight condition B. A typical
maximum wall temperature for a forced-convection-cooled wall of stainless
steel is 17600 R. With a maximum wall temperature of 12100 R, or limited
by minimum practical pressure drop, for a transpiration-cooled wall of
brazed and rolled wire cloth, the total flow ratios with transpiration
cooling are about 16 percent of the convective requirements for exhaust-
gas temperatures or 32000 and 37000 R. Ph-. corresponding inlet cooling-
air temperatures are shown at the top of the figure to be about 6350
an' 4440 R, respectively, for transpiration and forced-convection cooling.
The hi_1'r temperature with transpiration cooling results from compressor
bleed and from heat absorbed ia convectively cooling the impermeable wall
upstream of the porous wall.


Pressure Environment of Wire Cloth

The wire cloth successfully withstood the pressure _ur.j:-' of six
afterburner starts and the usual pulsations in pressure during normal
steady-state afterburiLill,.. Ignition of the afterburner usually caused
sur es in combustion-chamber static pressure to peak values of 2 to 5
inches of mercury. The surges damped out in less than 1 second to tht
steady-state values. 7T,.- surge in combustion-chamber static pressure
caused similar surges in cooling--air static pressure that varied anywhere
from 0.2 to 1.0 or 1.5 inches of mercury. These _1.ir dissipated to
steady-state values in several seconds. Dur-i n d..,-state operation


''. T. E1 'TL i,.


CONFIDENTIAL







NACA RM E54E25


both the cooling-air and combustion-chamber pressure traces showed
background pressure pulsations of 100 cps with a total amplitude of
about 0.10 to 0.15 inch of mercury superimposed with peaks up to about
1 or 2 inches of mercury total amplitude at roughly 3 cps. The duration
of these peaks was about 0.01 second. In one or two instances beats were
observed in the trace of combustion-chamber static pressure at a beat
frequency of 7 cps with a total amplitude of about 1 to 2 inches of
mercury.


CONCLUDING REMARKS

Cooling data were obtained for a transpiration-cooled afterburner
having a porous combustion-chamber wall of brazed and rolled wire cloth.
The data cover a range of e:.chust-gas temperature from 12000 to 3340 R,
total flow ratio of cooling air to combustion gas of 0.025 to 0.106,
and pressure altitudes of 15,000 to 45,000 feet. The data are success-
fully correlated over a range of Reynolds numbers from 75,000 to 1,500,000
based on the distance downstream of the leading edge of the porous wall.

Maximum wall tneper.at.res, based on the cooling correlation, were
determined for a porous wall of uniform permeability at sea-level take-off
and for flight Mach numbers of 0.8, 1.5, and 2.0 at an altitude of 35,000
feet. The cooling-air requirements were nearly independent of the flight
conditions. A total flow ratio of cooling air to combustion gas of about
0.032 can maintain a maximum wall temperature of 12100 R with an exhaust-
gas temperature of 37000 R. Savings in cooling air would, of course, be
possible with porous material having a higher allowable maximum wall
temperature.

The total flow ratios with a uniform permeability distribution and
air flows sufficient to limit the maximum wall temperature to 12100 R
are about 15 percent higher than the minimum total flow ratios corre-
sponding to a variable-permeability wall with a constant wall tempera-
ture of 12100 R.

The total flow ratios of cooling air to combustion gas for a
maximum wall tneper.-t..ire of 12100 R with the wire-cloth afterburner are
about 16 percent of the total flow ratios required to convectively cool
a stainless-steel afterburner wall to a maximum temperature of 17600 R
with exhaus::-E:a; temperatures of 32000 and 37000 R.

The analysis of cooling-air requirements for the previously
mentioned fli i-. conditions indicated that cooling-air static pressure
must be closely controlled, especially at a flight Mach number of 2.0.


Lewis Flight Propulsion Laboratory
Stional Advisory Committee for Aeronautics
Cleveland, Ohio, June 11, 1954


CONFIDEYTIAL


CONrFIDENTiAL.








CONFIDENTIAL


APPENDIX A


NACA RM E54E25


Ag

Cv
C



c
P
D

F,m

g

Hg, c



Hg, r


K


L


Nu

P

Pr

p

A( 2)




logir


CCOi nK l T I.A1J,


SYMBOLS

The following symbols are used in this report:

combustion-chamber flow area, sq in.

jet velocity coefficient, measured jet thrust
divided by isentropic jet thrust for measured
mass flow

specific heat at constant pressure

inside diameter of combustion chamber, in.

measured jet thrust, lb

acceleration due to gravity, ft/sec2

convective heat-transfer coefficient that would
apply to a solid surface under identical outside
flow conditions, Btu/(sec)(sq in.)(oR)

coefficient of nonluminous heat transfer,
Btu/(sec)(sq in.)(OR)

permeability coefficient, sq in. (defined in
eq. (6) of ref. 3)

distance from flame holder to exhaust-nozzle exit,
66 in.

Nusselt number

total pressure, lb/sq in. abs

Prandtl number

static pressure, lb/sq in. abs

difference between squares of absolute static
pressures on both sides of porous wall, p 2 p;.
I'b"/in.4
I i/
A(o -oj pr sressure-drop parameter, li8/i!.
fT /-" T."







NACA!M E54E25 /


R gas constant, ft-lb/(lb)(OR)

Re Reynolds number, (pU)g(x 21)/4f

T total temperature of a gas or surface temperature of a
solid, OR

Tg + Tw oR
Tf film temperature, g + T

Tg, 1 total temperature at combustion-chamber inlet, OR

Tg,2 total temperature of exhaust gas at nozzle exit, OR

To NACA standard sea-level temperature, 518.40 R

Tw Ta
T T temperature-difference ratio
Tg Ta

U velocity in axial-flow direction, in./sec

V velocity normal to porous-wall surface, in./sec

W weight flow, lb/sec

WJWg total flow ratio

x distance downstream of quick-disconnect coupling at
combustion-chamber inlet, in.

x' distance downstream of flame holder, x 2, in.

Ax incremental length of combustion chamber, in.

ag absorptivity of combustion gas

Tg ratio of specific heats of exhaust gas

a emissivity

4 absolute viscosity, lb/(in.)(sec)

p weight density, lb/cu in.

(pU)g,x total weight flow per unit of combustion-chamber flow area,

W+ Wa /Agx, lb/(sec)(sq in.)
\Wg 0


CONFIDENTIAL


CONFIDENTIAL








NACA RM E54E25


(pV)a iJ. reduced weight flow per unit-area, lb/(sec)(sq in.)


(pV)J(pU)g coolant-flow ratio

I thickness of wire cloth, in.

Subscripts:

a cooling air

f refers to property evaluated at film temperature Tf

g combustion gas

w wall

x at distance x

0 free-stream conditions

Numbers greater than 2 represent stations along combustion chamber in
inches downstream of quick-disconnect coupling.


CONFIDENTIAL


CO1FIDEIJTIMAL







NACA RM E54E25 CONFIDENTIAL 19


APPENDIX B


COMr.PAEISON OF :>_EFI-TMEI[iAL COOLING CORRELATION WITH THEORY

A brief comparison is made between the experimental cooling cor-
relation and the atlpro_::.i.iate theory of Rannie and Friedman refss. 4 and
5) for turbulent boundary-layer flow. The comparison is made between
the temperature-difference ratio for individual data points and the
theoretical temperature-difference ratio corresponding to the same ex-
preriment al c: itions.

The theoretical equation is given by Friedman (ref. 5) as


T Ta r
T T rp (Bl)
g a e +r 1

where r is the ratio of the velocity parallel to the surface at the
border between the laminar sublayer and the turbulent part of the bound-
ary layer to the stream velocity outside the boundary layer. Eckert
(ref. 14) -ives


r = 2.11 (B2)
(Re)0.
g

(pV)acpa
I = (B3)
g,c

For turbulent boundary-layer flow over a flat plate, Colburn gives

Nu = O.0296(Re)O.8(Pr)l/3 (B4)

where Re is >':-->,' on the distance from the 1 -l:'n- -.--. e-r
equation (B4) gives

Hg c 0. 0 ,7
RePr (.-') 02 (Pr2 (B7)
s .- p (Re) (Pr)'

from which the local value of convective heat-transfer i'-:- 'ficient at
any station is







C '.- i'iL,"-,itL







NACA RM E54Z13


0.0296(pUcp)gx
Hg,c,x = 0 2 (B6)
(Re)0 (Pr)2/3
g,x g,x

The effect of temperature level on the Prandtl number is so small that
a mean value was used for all cases.

The effects of nonluminous radiation were accounted for by substi-
tuting the sum of the.equations (B6) and (B7) for Hg, in equation
(B3). rr-r the data of reference 15,

0.173 Pe -Tg 7w4 4 (w7)
g,r,x : (144)(3600)(Tg,xTwx) x 100 x \100 /


where Hg,r,x is the heat-transfer coefficient for nonluminous radia-
tion at station x. A value of 0.52 was used for the pseudoemissivity
e'. The local values of combustion-gas emissivity & and absorptiv-
s g,x
ity ag,x were based on the total fuel-air ratio and l.c-: one-
dimensional values of temperature and static pressure. No distinction
was made between total and static combustion-gas temperature, because
the Mach numbers in the combustion chamber were low.

The density and viscosity in the Reynolds number in equation (B6)
are usually based on the gas temperature just outside the boundary layer.
However, in order to obtain a better correlation of heat-tr nr.fer data
over a lr -' : of Tw/Tg, the combination of equations (B3) and (B6)
(i --.'ing radiation) was modified in reference 16 to

S (Re) 0.2 ()2/3 (p) T
0.0296 (PU)g T
0 0

The temperature ratio Tw/Tg has the effect of evaluating the density
ond viscosity in the Reynolds number at the wall temperature.

In view of equation (B8), the followi".- equation was derived for (x:

Tx (PVcp)
'/wx -a.x.
g,x 0.0296(pUc p) 0.173 e 4 kT \4 ,T \4
gx s I g J-.t
(Re)0.(Pr)2/3 (144)(3600)(Tx -T ) gx \100 / gx j00
g,x g,x
(B9)


CONFIDENTIAL


CONFIDENTIAL







NACA RM E54E25


Equation (B9) was used in equation (Bl) and found to overcorrect
the theoretical temperature-difference ratio, so that the intermediate
temperature ratio Tx/Tf,x was used in the comparison figures 18
and 19. With afterburning, the inclusion of Tw,x/Tfx in equation
(B9) results in theoretical wall temperatures generally higher
(fig. 18(a)) than those measured. 'U-r! a value of 1 is assumed for
Tw,x/Tf,x, the theoretical wall temperatures are lower than those meas-
ured and the scatter is less (fig. 18(b)).

For the nonafterburning data, the inclusion of Tw,x/Tf,x in
equation (B9) results in theoretical wall temperatures that are equal
to or slightly lower than those measured (fig. 19(a)). The assumption
of Tw,x/Tf,x of 1 results in a slight lowering of the theoretical
wall temperatures over the entire range of temperature-difference
ratio (fig. 19(b)).

The determination of an empirical coefficient or exponent for
Tw,x/Tf,x to produce a better agreement does not appear warranted be-
cause of differences in the assumptions made in the theory and condi-
tions in the afterburner. Therefore, no conclusion is made as to the
validity of including the temperature ratios Twx/Tfx or Tw,x/Tg,x
in the equation for cp.


REFERENCES

1. Eckert, E. R. G., and Livingood, John N. B.: Comparison of Effec-
tiveness of Convection-, Transpiration-, and Film-Cooling Methods
with Air as Coolant. NACA TN 3010, 1953.

2. Koffel, William K.: Preliminary Experimental Investigation of
Transpiration Cooling for an Afterburner with a Sintered, Porous
Stainless-Steel Combustion-Cha'b-r Wall. NACA RM E53D08, 1953.

3. Koffel, William K.: Air-Flow Characteristics of Brazed and Rolled
Wire Filter Cloth for Transpiration-Cooled Afterburners. NACA
RM E53H24, 1953.

4. Rannie, W. D.: A Simplified Theory of Porous Wall Cooling. Prog.
Rep. 4-50, Power Plant Lab. Proj. No. MX801, Jet Prop. Lab.,
C.I.T. Nov. 24, 1947. (AMC Contract No. W-.. 7:--.--20260, Ord.
Dept. Contract No. W-04-200-ORD-455.)

5. Friedman, Joseph: A Theoretical and Experimental Investigation of
Rocket-Motor Sweat Coclint-. Jour. Am. Rocket Soc., no. 79, Dec.
1949, pp. 147-154.


CONFIDENTIAL


C ii i Fi '1.i AL







22 CONFIDENTIAL NACA RM E54E25


6. Huntley, S. C., Auble, Carmon M., and Useller, James W.: Altitude
Performance Investigation of a High-Temperature Afterburner. NACA
RM E53D22, 1953.

7. Conrad, E. William, and Campbell, Carl E.: Altitude Wind Tunnel In-
vestigation of High-Temperature Afterburners. NACA RM E51L07, 1952.

8. Weinbaum, S., and Wheeler, H. L., Jr.: Heat Transfer in Sweat-
Cooled Porous Metals. Prog. Rep. 1-58, Air Lab. Proj. No. MX121,
Jet Prop. Lab., C.I.T., Apr. 8, 1947. (AMC Contract No.
W-535-ac-20260, Ord. Dept. Contract No. W-04-200-ORD-455.)

9. Pinkel, Benjamin, and Turner, L. Richard: Thermodynamic Data for
the Computation of the Performance of Exhaust-Gas Turbines. NACA
WR E-23, 1944. (Supersedes NACA ARR 4B25.)

10. Wallner, Lewis E., and Wintler, John T.: Experimental Investiga-
tion of Typical Constant- and Variable-Area Exhaust Nozzles and
Effects on Axial-Flow Turbojet-Engine Performance. NACA RM
E51D19, 1951.

11. Koffel, William K., and Kaufman, Harold R.: Empirical Cooling
Correlation for an Experimental Afterburner with an Annular
Cooling Passage. NACA RM E2':i., 1952.

12. Keenan, Joseph H., and Kaye, Joseph: Gas Tables. John Wiley &
Sons, Inc., 1948.

15. Koffel, William K., and Kaufman, Harold R.: Cooling Characteris-
tics of an Experimental Tail-Pipe Burner with an Annular
Cooling-Air Passage. NACA RM E51K23, 1952.

14. Eckert, E. R. G.: Introduction to the Tri-i-nfer of Heat and Mass.
McGraw-Hill Book Co., Inc., 1950.

15. McAdams, William H.: Heat Transmission. Sec:.-in ed., McGraw-
Hill Book Co., Inc., 1942, pp. 64-69.

16. Esgar, Jack B.: An Analytical Method for EvaluirMtin Factors
Affecting Application of Try- spiration Cooling to Gas Turbine
Blades. NACA RM E52G01, 1952.


CONFIDENTIAL
















NACA RM E54E25


CONFIDENTIAL













d -? ,a i "
. C. ..= o









I i








































t a






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,u "






































aCt, a,
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0 -
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ral









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M r


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t-, c Z-




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umr.
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W



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-rs,


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'- ---
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CONFIDENTIAL


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i ,


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00c- CC' NS



































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NACA RM E54E25


TABLE II. COMBeUSTION-CHAMBER WALL TEMPERATURES, OF


Series Channel Solid metal Wire-cloth porous wall
and run wall

Tw,10 Tw,19 Tv,24 Tw,32.6 Tww,1 T,49.6 Tw,57.5


127-27


127-26


127-25


127-24


127-23


127-28


127-21


127-20


130-17


127-17


127-22


130-24


128-4


129-18


129-7


129-32


127-18


558
588

543
570

499
524

468
490

429
452

428
451

600
610

567
583

557
568

538
568

544
568

569
576

839
686

542
370

532
558

539
566

484
.-' ./


553
530

533
512

489
464

449
429

411
392

409
393

549
534

518
499

514
489

503
482

500
481

518
493

815
802

509
482

496
473

510
483

433
413


429
459

412
420

348
338

302
283

268
240

269
240

437
422

394
372

372
338

367
338

367
337

376
347

810
719

371
338

366
326

372
338

297
255


459
422

422
382

338
292

282
239

240
198

240
198

462
388

407
331

355
302

372
295

373
294

364
310

725
677

360
298

353
288

359
296

288
213


443
366

410
331

326
254

270
210

228
183

229
186

460
346

406
293

347
287

368
261

365
260

354
289

780
762

348
279

342
259

346
280

272
192


419
323

389
293

310
231

260
193

213
160

219
160

434
312

379
263

321
252

339
236

343
219

330
262

745
885

329
242

325
237

327
244

251
176


400
358

370
330

300
265

250
220

212
181

213
182

422
339

362
287

322
273

333
256

332
265

330
282

850
823

324
267

322
259

324
270

243
192


CONFIDENTIAL


CONFIDENTIAL








NACA RM E54E25


TABLE II. Continued. COMBUSTION-CHAMBER WALL TEMPERATURES, OF


Series Channel Solid metal Wire-cloth porous wall
and run wall

Tw,10 Tw,19 w,24 Tw,32.6 Tw,41 Tw,49.6 Tw,57.5


127-19


129-25


129-26


129-24


129-23


130-37


130-32


130-33


130-31


130-30


130-29


130-34


130-35


130-36


129-21


129-22


129-20


454 I 401


470

648
669

649
672


388

662
625

669
632


628 I 640


648

618
640

598
605

648
662

628
640

622
633

622
688

622
638

664
674

676
682

685
687

379
592

625
653

620
650


600

630
599

605
577

665
649

643
622

632
602

636
608

632
608

681
682

695
707

703
725

577
541

640
650

631
629


220

596
519

608
540

560
481

552
451

506
467

658
556

606
498

586
515

600
500

593
492

692
538

718
551

735
545

457
390

580
417

541
397


182

649
478

669
499

599
433

581
403

552
425

684
506

622
448

608
473

611
454

606
448

751
478

806
488

845
482

468
351

619
362

574
347


232 I 219


172

680
456

699
488

622
420

611
391

582
396

729
452

662
408

631
437

642
418

648
410

823
427

891
436

930
430

478
348

665
352

617
339


151

703
384

728
397

648
359

642
334

577
359

752
403

691
368

638
792

658
375

672
370

877
378

930
382

955
380

478
398

709
299

652
290


163

768
458

790
476

700
429

686
401

602
419

798
467

733
428

665
453

702
437

701
430

890
432

948
438

978
430

505
356

760
355

707
340


CONFIDENTIAL


CONFIDENTIAL








NACA RM E54E25


TABLE II. Concluded. COMBUSTION-CHAMBER WALL TEMPERATURES, OF


Series Channel Solid metal Wire-cloth porous wall
and run wall
T T T T T T T
w,10 w,19 w,24 w,32.6 w,41 w,49.6 w,57.5


128-5


128-6


130-20


130-21


130-18


130-19


129-19


129-11


129-10


129-8


129-9


130-27


130-28


130-26


130-25


130-22


130-23


475
489

493
508

690
669

726
721

652
669

670
682

650
696

685
732

662
707

637
678

634
679

709
735

730
751

741
768

753
787

723
769

749
690


435
422

456
441

678
648

710
676

649
626

662
634

672
690

712
748

689
718

660
677

652
678

720
752

741
774

758
808

787
813

741
798

777
845


300
262

321
284

564
505

617
588

537
431

544
465

589
425

647
492

600
449

570
408

546
390

611
484

647
519

662
518

679
532

530
472

560
522


I I


285
220

308
240

629
470

701
572

569
388

589
422

647
368

752
433

680
392

623
353

589
338

682
426

732
464

761
458

803
472

577
420

630
478


281
215

302
235

666
457

744
527

603
380

627
415

692
352

828
418

738
379

675
344

628
330

751
400

820
432

847
430

900
948

602
392

663
447


269
187

293
203

689
401

760
490

629
334

654
364

718
302

870
354

713
322

712
297

650
282

778
361

846
393

880
391

926
408

616
362

681
416


264
214

287
232

759
472

842
558

703
388

727
428

830
354

933
412

857
378

802
340

727
330

842
420

905
449

943
457

986
478

680
426

749
490


CONFIDENTIAL


CONFIDENTIAL








NACA EM E54E25


TABLE III. COOLING-AIR AND SHROUD TEMPERATURES, OF


Series Channel Cooling air Shroud wall
and run Ta,19 Ta,24 Ta,32.6 a,41 Ta,49.6 Ta,57.5 Ts,24 Ts,41 Ts,57.5
(a)


112 214 I 204 I 195 185


127-27


127-26


127-25


127-24


127-23


127-28


127-21


127-20


130-17


127-17


127-22


130-24


128-4


129-18


129-7


129-32


127-18


121

107
117

101
108

97
103

94
98

94
98

115
125

105
117

104
113

101
110

106
115

107
117

97
107

101
112

99
109

104
115

95
101


186

200
175

170
152

152
140

137
128

136
128

213
179

190
160

171
153

177
152

184
158

175
159

175
151

170
151

169
149

171
154

149
185


179

182
114

157
108

143
108

130
107

130
123

199
173

178
101

162
151

167
104

172
105

166
154

165
149

161
155

160
143

164
155

193
101


184


157


142


131


130


196


175


159


164


170


165


163


160


158


163


142


167

175
157

150
140

137
130

127
122

127
122

182
166

161
145

153
140

153
138

162
150

156
144

152
139

133
135

135
136

144
140

136
127


169


160


138


126


116


116


168


178
165

166
154

145
135

131
125

120
117

123
118

174
162

146
138

132
132

138
132

160
149

144
141

137
132

148
126

145
131

150
136

131


188


175


147


133


124


125


180


194


179


150


133


121


123


190


150 157 155


144


148


157


151


142 148


138


136


141


144


142


148


141


147


168


154


148


138


140


146


123 130 I 131


123 --- --- I


I_ _I. I. I __ L _I _


aValues read from circumferential temperature pr.files.

CONFIDENTIAL


139


140


147


144


1
9


CONFIDENTIAL








CONFIDENTIAL NACA RM E54E25


TABLE III. Continued. COOLING-AIR AND SHROUD TEMPERATURES, OF


Series Channel Cooling air Shroud wall
and run Ta,19 Ta,24 Ta,32.6 a,41 Ta,49.6 Ta,57.5 Ts,24 Ts,41 Ts,57.5

(a)


127-19 1 92 137 133 131


129-25


129-26


97

126
130

129
134

122
127

117
122

112
120

124
125

119
121

117
122

117
122

115
121

125
124

127
128

128
128

107
117

116
121

114


127

255
200

264
207

235
190

223
182

219
190

265
213

240
196

241
201

138
197

235
194

264
208

273
212

274
210

190
167

219
174

210


229


236


214


205


198


236


217


215


214


213


239


248


250


176


204


197


121 115 121 120


127
120

242
181

254
185

224
173

210
165

195
163

231
177

216
168

210
171

208
168

210
166

239
174

249
177

256
177

178
154

210
160

203


aValues read from circumferential temperature profiles.


195 224 257


200


181


171


167


194


179


180


177


176


192


196


198


123-23


130-37


130-32


130-33


130-31


130-30


130-29


130-34


130-35


130-36


129-21


129-22


129-20


169


164


232


208


195


191


227


210


206


205


206


230


239


246
S---


191


183


273


239


216


209


256


244


226


222


230


275


198


315


184


218


205


I __ I


CONFIDENTIAL


199
161


157
157


150 165


I I I I I I








NACA RM E54E25


TABLE III. Concluded. COOLING-AIR AND SHROUD TEPEFA.Ti.iFE., OF


Series Channel Cooling air Shroud wall
and run Tal 19Ta,24 Ta,32.6 Ta41 Ta,49.6 Ta,57.5 Ts,24 4 Ts T57.5
(a)I I


128-5


128-6


130-20


130-21


130-18


130-19


129-19


129-11


129-10


129-8


129-9


150-27


130-28


130-26


130-25


130-22


130-23


89 140 135 134 130


96

90
98

129
132

137
142

117
122

124
127

118
123

124
128

118
125

114
119

118
120

127
133

129
137

129
137

126
137

127
139

132


131

146
136

243
197

273
218

215
177

229
187

224
177

249
191

233
182

217
171

213
170

246
198

258
204

260
205

256
206

227
198

247


130

140
136

253
253

248
334

199
170

211
182

220
175

247
187

214
197

200
170

195
165

226
144

236
130

239
197

236
191

212
185

229
201


139


179


259


203


216


213


238


220


205


200


236


247


251


247


216


235


149 211


122

134
127

224
179

254
199

194
158

211
170

212
160

240
175

228
164

191
154

210
156

241
179

250
179

257
185

245
180

209
173

233


--- 1 198


121 117 125 123


191


217


169


121 129 128


219


256


186


117

124
123

239
190

274
214

186
156

244
177

206
158

236
172

214
155

185
150

197
157

275
190

279
184

296
198

255
186

219
171

258


173


191


179


166


164 184 215


192


200


201


181


195


197


224


204


186


225


236


241


199 235 283


196


219


256


303


199


220


275


238


199


293


302


325


229


178


180 203 234


values read from circumferential temperature profiles.


CONFIDENTIAL


CONFIDENTIAL


218 I --- I --- I ---












CONFIDENTIAL


TABLE IV. COMBUSTION-GAS AND COOLING-AIR PRESSURES


Series Channel
and run-



9

127-26 1
17-26


g, 10


Combustion-chamber static pressures, Ib/sq in. abs

Pg,19 g,24 g g,32.6Pg,41 Pg,49.6 8g,57.5g,65. Pg,67.8


9.507 9.778 9.840 9.688


127-25


127-24


127-23


127-28


127-21


127-20


130-17


127-17


127-22


130-24


128-4


129-18


129-7


129-32


127-1i


1;:7-19





129-2K


129-24









130-52


130-3151


9.556 9.243 9.188 8.715 8.354


9.194


9.257


9.285


9.292


9.299


4.326


4.340


4.444


4.354


4.326


4.389


4.33


4.431


4.493


4.465


4.382


4.417


12.65


12.63


12.65


12.64


13.73


14.09


14.12


9.132


9.146


9.139


9.097


9.104


4.257


4.250


4.194


4.250


4.222


4.146


4.22





4.354


6.438


4.208


4.208


8.625


8.653


8.646


8.681


8.646


3.854


3.833


3.931


3.875


3.826


3.854


3.82


3.910


3.958


3.922


3.847


3.847


10.38


10.36


10.38


10.40





11.43


11.50


I 1 i


Cooling-air pressures, lb/sq In. abs

Pa,24 TPa,24 IPa,32.6 P,41 Ia,49.61.P,57.5


10.29
10.42


Pa,15

10.4


10.62


11.71





14.81


14.70


5.563


6.055


6.605


6.493


6.514


6.470


6.425


6.61


6.82


6.748


8.326


9.722


15.81


15.59


16.23


16.76


17.46


17.33


18.01


1 4.500 4.708 4.757


1 4.542 4.750 4.778
9

1 4.625 4.84 4.889


1 4.569 4.778 4.819




9

1 4.56 4.78 4.876
9

1 4.576 4.792 4.816
9
1 1 4.576 4.792 4.891


1 4.715 4.931 4.93E
9

L 4.625 4.833 4.854
9

1 4.701 4.903 4.931
9

1 4.B06 5.000 5.00(
9

1 4.63 14.55 14.39


1 14.58 14.51 14.36
9

1 14.65 14. 60 14.42
9

1 14.74 14.65, 14.49
9





1 16.65 14.47 16.42


1 f0) 1 1 .. )


10.21
10.30

10.37
10.51

11.19
11.49

12.14
12.65

13.43
14.25

13.31
14.13

5.39
5.48

5.780
5.928

6.195
6.471

6.121
6.322

6.136
6.349

6.079
6.345

6.031
6.239

6.21
6.47

6.39
6.68

6.313
6.600

7.549
8.017

8.66
9.33

15.17
15.58

15.02
15.38

15.08
15.82

15.41
16.28

16.67
17.18

16.70
17.10


10.22
10.30

10.39
10.51

11.23
11.52

12.20
12.70

13.53
14.33

13.42
14.22

5.40
5.48

5.795
5.93

6.227
6.498

6.153
6.33

6.158
6.35

6.105
6.366

6.053


6.23
6.48

6.41
6.69

6.334
6.175

7.613
8.045

8.74
9.38

15.23
15.60

15.06
15.40

15.16
15.87

15.51
16,36

16.74
17.24

16.76
17.14


17.2 17.26 17.25
17.74 17.77 17.77


CONFIDENTIAL


NACA RM E54E25


17.53 17.01 17.19
18.08 17.56 |l7.71


10.22
10.29

10.38
10.50

11.23
11.51

12.21
12.69

13.54
14.33

13.42
14.21

5.39
5.46

5.790
5.918

6.216
6.492

6.147
6.323

6.158
6.338

6.100
6.366

6.053
6.265

6.25
6.49

6.41
6.70

6.334
6.621

7.613
8.028

8.74
9.37

15.23
15.61

15.06
15.41

15.18
15.89

15.52
16.38

16.74
17.24

16.76
16.79











NACA RM E54E25


CONFIDENTIAL


TABLE IV. Concluded. COMBUSTION-GAS AND COOLING-AIR PRESSURES


Series Channel Combustion-chamber static pressures, lb/sq in. abs
and run
PglO Pg,19 Pg,24 Pg,32.6 Pg,41 Pg,49.6 Pg,57.5 Pg,63.8


130-31


130-30


130-29


130-34


130-35


130-36


129-21


129-22


129-20


128-5


128-6


130-20


130-21


130-18


130-19


129-19


129-11


129-10


129-8


13-9


130-27


130-28


130-26


130-25


130-22


130-23


5.6041 5.5421 5.472


15.36


15.53


15.56


16.43


16.59


16.76


11.43


12.48


10.45


6.54


6.47


7.778


7.681


7.986


7.924


8.382


8.299


8.299


8.396


8.396


8.542


8.528


9.319


8.792


5.340


5.208


13.69


12.47 11.15


12.59
^ "


4.903 4.660 I 4.250


Pg,67.8 Pa,15

8.340 16.86


8.312 17.40


8.361 17.52


8.381 18.48


8.361 18.75


8.381 19.18


7.153 14.42


7.056 16.42


5.840 14.54


3.60 11.33


3.51 10.67


3.500 9.854


3.528 8.798


3.535 10.75


3.528 10.22


3.646 12.15


3.542 10.98


3.563 11.50


3.438 12.08


3.590 12.70


3.500 11.51


3.500 11.04


3.500 11.61


3.528 11.87


2.208 8.430


2.229 7.048


Cooling-air pressures, Ib/sq in. abs

Pa,24 24 24 -.." ]a,41 P,49.6 / a,57.5


8.107
8.474

7.483
7.780


16.08 16.22
16.52 16.64

16.48
17.00

16.57 16.74
16.98 |17.25

17.60 17.77
18.11 18.26

17.80 17.96
18.30 18.44

18.16 16.35
18.70 16.85

12.75
13.79 .-

14.26 14.72
15.51 15.83

13.14 15.45
13.96 14.16

9.70 10.08
10.56 10.82

9.24 10.17
9.98 10.21

9.013 9.119
9.342 9.417

8.464 8.523
8.6611 8.703


7.687 7.841
8.128 8.229

7.175 7.291
7.515 7.589


16.27 i1.26
16.67 16.68

16.70 16.71
17.18 17.19

16.61 16.82
17.30 17.31

17.83 17.84
18.31 18.32

18.03 16.03
18.50 18.50

18.41 18.41
18.91 18.91

13.25 13.26
14.13 14.15

14.85 14.87
15.92 15.94

13.52 13.53
14.22 14.23

10.17 10.17
10. 1 10.93

9.64 9.64
10.29 10.30

9.151 9.151
9.449 9.496

8.544 8.544
8.714 8.714

10.10 10.10
10.57 10.57

'8 i 9.678


10.43


11.27 11.27

11.29 11.31
11.84 11.86 .

11.75 11.77 11.77
12.40 12.41 12.42

-" 10.86 10.87
11.30 11.30

10.49 10.51 10.51
10.85 10.87 10.86

10.98 11.00
: 11.40

11.24 11.25 11.26
11.64 11.66 11.66

- -1 7.879 7.879
S 8.261 8.266

7.307 7.323 7.323
7.600 7.610 7.610


CONFIDENTIAL













CONFIDENTIAL


NACA RM E54E25


CC "N 0CC CCC N~ C)0 -~ 'CC Cf 0CC CCC CCC CO3 Z~ s ~ ~ j
C) NC) 'CCC CCC 'CCC CCC NC) N~ C) C) C)C 0CC NC) CC)C C) I)S


C)C CCi C C C C C C ) CCC 'CCC -, C)- ij
CY C )C ) C C X C* N C C C O) C )









s i ; o a e E -Q NC) Cf



N-C -C- NC' C- NN XC= N CCC C>C N N))C C '

C -fC- C) C : CC CC CC C C
K C C.,C CC'= s~ f=9 r-C9~ f, C~ cN C C C C C~ m C.-~C"C ,:" ,,:C



C ~ ~ ~ ~ m C)S C)C CC H-CC CCC CC C)C C)))) CHC ))- C ) C C NO


8 g _.9 8 8 8 8





3 o o o 8 o 8 8 8
0 a? i 8 r 8 81 y
-- o o o o


S ^ S S ^ o ^ S


C)w R C
CC CC N


CONFIDENTIAL


~c~c~Oa~
Nnn~~r
mmmmnn

~~n~rN~
~""
~ina ~a
mmmmnn


Nnnni~a
mmmitn


~rz~=a ~
Nnnne
-rmmmio


m o
08~'
" "
=jmm


F "
,,,~


ioOd
-rvn
a~rv;~
c


r
~mammmmln
f
g?"Jr;'D:: C~~
N4Qn~n
=amOmmmmln


~ammi3mmm~n

i i ~
c;aa~m~an


t~'~ s


d=~gl


9 =
=""~


c

oicooi


Pm4mlllin



4~"RX"I
~ul~tnin~u7


""=a
'7 j: O C
C o d


1 0.










NACA RM E54E25


CONFIDENTIAL


*0 ..' 4,*- ~ifi ( 0 C-CO 00 -C 00 00i ~ j; y 0 .-.-* 0- o 0 u






T O N C O, 00 0 "" -0 0 .'
ot






.- ,
0 'i0. 0 0 0 .. ._
0i CO uCi 't O 00 0^^ '0 00 0 00 00i 0 0 00, ,; 0s' ',



.r .~ _0 .-i\ .~ .Q .. .i a .li C ^ o ( i? N> 0 Q -s f-' -0 > ^ : O ~
0 o 04, 0 o 0 0 00 0 0.. ..00 ..- .- o ., ~ o .2o 0,
C. 00 0n 0^i 0 '* cr> { 00 00~ 0- .00 0 -i '0 -. 0- 0 0 O 0i
0 0 CC'-. 0.t^ ^ ~ 3 00 00 00 00 0f- o"( K F 00 00- .- .-.I ^*W t^ *^i- ~ SC




S- s -. < -






0 0 0 0: a : 0
ina *',D















0t 0 0s 0> 0S 0" C I.'''i 0^ <0 0^ 0^ D f r 0 f o
4i 0.





















S* 0 0 C 0 0 0 0 0 0 0 '0 0 C







0 0C S y? ef SD Q m B3 t n o 0 0e 0- 0r C~ 0N .0 000
0~ 0~ 0C 0m 0~ 0 0i 0N 0rr 0 0

e B
























0 00 0- 0 0 N0
0 C. 10 ~ 0j 0~ 0n 0N 0N 0n ~ 0n 0v 0a 0


jO 0 0 ; 0 o 0 Y 0 0 0~ -. Nn 0= 0m 0oo h ~ s~ ;
n3 ti ii r-ri, i ,,
ti EC 5 u: ,
's ri "^ eo i< io o c- e'- f-- N tN>.< r- O O 'f
Z; s^ =:DS im


*<* w ** r N tsi IM

d r f M i 'l'


10 080 01S0


CONFIDENTIAL


B1..-C
'a:
I~
-L-

,r


s sn


77NYY~
ii~irtr'











CONFIDENTIAL


NACA RM E54E25


,? I"C mn=~ oo -N -sZn Z- Ir au Cr- N n~








2 -- --C
HI


H3


H=
o




z
H
I ^-" '" bl ^ ^







U

0






o
& ?;








C)

o H E -
m, I : ? s *

u '"

| |_^^





r-i : *" __ ^ ^ _____


' jii^L- -. ~, ^

> !!tS 2 ; ; : !
w 1 ''"______


CONFIDENTIAL


r

s: s F r
R
3 :: ;1 ~: I



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CONFIDENTIAL


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NACA RM E54E25


CONFIDENTIAL


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Combustion-gas static pressure
Wire-cloth or solid metal wall air-side temperature
Wire-cloth gas-side temperature


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Station, x, in.
(a) Wire-cloth and combustion-gas instrumentation.
Figure 5. hmchematic development showing locations of temperature and pressure
instrumnlnt iaton on wire- loth purous-wall afterburner.


CONFIDENTIAL


Gas
flow


'


I


CONFIDENTIAL


PI






NACA RM E54E25


Cooling-air static pressure
Cooling-air total pressure
Cooling-air temperature
Shroud outside-skin temperature
I I


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32.6
Station x, in.


(b) Shroud and cooling-air instrumentation.
Figure 5. Concluded. Schematic development showing locations of temperature and
pressure instrumentation on wire-cloth porous-wall afterburner.

CONFIDENTIAL


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1 5 9 13 17 1
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(a) Afterburning. Exhaust-gas temperature, 29540 R; total
flow ratio, 0.0716.

Figure 8. Typical circumferential profiles of wire-cloth
temperature and cooling-air static pressure. Altitude,
35,000 feet; flight Mach number, 1.0.


CONFIDENTIAL


Station
24
32.6
41
49.6
57.5


;EiPtL- I WV.^T


10(


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NACA RM E54E25



























800 -
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CONFIDENTIAL


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1 5 9 13 17 1 5
Channel

(b) Nonafterburning. Exhaust-gas temperature, 12490 R; total flow ratio,
0.0951.

Figure 8. Concluded. Typical circumferential profiles of wire-cloth
temperature and cooling-air static pressure. Altitude, 35,000 feet;
flight Mach number, 1.0.


CONFIDENTIAL







CONFIDENTIAL


NACA RM E54E25


*17



4I0 I I I _0_i__ ___ __

-- y ._-4- -.-C- ..


180(


S160(
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140C


StatIon, In.
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i ~ T 1[ < t'dl al pr(/,'l 1 f ,: wire-cloth after urner. A1tltud e, 3S,000


C''NFIDEIITIAL







NACA RM E54E25


i I


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CONFIDENTIAL


CONFIDENTIAL


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CONFIDENTIAL


NACA RM E54E25





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NACA RM E54E25 CONFIDENTIAL 49













i0
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t-- t- 01 --- ---











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I ) Na afte i 1 1tat' c-p n 1r I hIo 7 -1 .-r I / I i _ga t FrI .1
1249 H ratio, C.9;1.
Figure 9. Concluded. Typical longitudinal profiles for wire-e lot afterburner. Altitude,
35,o00 feet; flight Mach number, 1.0.

CONFIDENTIAL









50 CONFIfDEITIAL








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NACA RM E54E25









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3 .005 .010 .015 .020 .025

Coolant-flow ratio, (pV)a/(pU)g


(a) Reynolds number, approximately 75,000.

Figure 12. Correlation of afterburning cooling data for
wire-cloth afterburner.


CONFIDENTIAL


CONFIDENTIAL









NACA RM E54E25


.005 .010 .015 .020 .025
Coolant-flow ratio, (pV)a/(pU)g

(b) Reynolds number, approximately 300,000.

Figure 12. Continred. Correlation of afterburning cooling
data for wire-cloth afterburner.


CONFIDENTIAL


0



Cd
I-I


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CONFIDENTIAL









NACA RM E54E25







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CONFIDENTIAL


0 .005 .010 .015 .020
Coolant-flow ratio, (pV)a/(pU)g


(c) Reynolds number, approximately 500,000.


Figure 12. Continued. Correlation of after-
burning cooling data for wire-cloth afterburner.


CONFIDENTIAL








CONFIDENTIAL


NACA RM E54E25


D .005 .010 .015 .020
Coolant-flow ratio, (pV)a/(pU)g

(d) Reynolds number, approximately 800,000.

Figure 12. Contirnueid. Correlation of after-
burnring cooling data for wire-cloth afterburner.


CONFIDENTIAL


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CONFIDENTIAL


0 .005 .010 .015 .020
Coolant-flow ratio, (pV)a/(pU)g

(e) Reynolds number, approximately 1,000,000.


Figure 12. Continued. Correlation of after-
burning cooling data for wire-cloth afterburner.


CONFIDENTIAL









NACA RM E54E25


Series Channel
1 9

128 0 M
129 *
130 A A


















































0 .005 .010 .015 .020
Coolant-flow ratio, (pV)a/(pU)g


(f) Reynolds number, approximately 1,500,000.


Figure 12. Continued. Correlation of after-
burning cooling data for wire-cloth afterburner.


CONFIDENTIAL


ca as

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c,
40

a)
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a)

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NACA RM E54E25



1.0







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CONFIDENTIAL


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Reynolds number,
Re


75,000







500,000
500 000



800,000



.1,000,000

1,500,000










0 .005 .010 .015 .020 .025
Coolant-flow ratio, (pV)a/(pU)g

(g) Mean curves for approximate Reynolds numbers of 75,000 to
1,500,000.

Figure 12. Concluded. Correlation of afterburning cooling
data for wire-cloth afterburner.


CONFIDENTIAL










CONFIIENTIAL


NACA RM E54E25


4 6 10
Reynolds number, Re


Figure 13. Variation of temperature-difference ratio with
for nonafterburning cooling data.


Reynolds number


CONFIDENTIAL


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E-




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Figure 14. Correlation of nonafterburning cooling data for
wire-cloth afterburner.


COaFIDENTIAL


CONFIDENTIAL









NACA RM E54E25


.005 .010
Coolant-flow ratio,


.015
(pV)a/(PU)g


.020


.025


Figure 15. Comparison of cooling correlations for afterburning and
nonafterburning for wire-cloth afterburner.


CONFIDENTIAL


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-- Nonafterburning








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number,
Re

105
0.75x105












\\ 106
____-s-i0


CONFIDENTIAL







NACA RM E54E25


Flight Flight Altitude, Cooling-air pressure,
condition Mach ft Ib/sq in. abs
number Min. Max.

0 A 0 Sea level 25.60. 27.02
O B .8 35,000 10.85 14.55
0 C 1.5 35,000 20.89 24.20
A D 2.0 35,000 35.04 35.70













Approximate peak temperature
limit of present wire cloth
due to oxidation of braze
alloy .







-------- C








0 .02 .04 .06 .08 .10
Total flow ratio, Wa/Wg

Figure 16. Cooling characteristics of transpiration-cooled wire-cloth
afterburner having uniform permeability distribution. Exhaust-gas
temperature, 3700 R.


CONFIDENTIAL


1800





1600


1400





1200





1000


CONFIDENTIAL








CONFIDENTIAL


NACA RM E54E25


0
'Oi~
,a "
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o :



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CONFIDENTIAL


I I Exhaust-gas
Transpiration- temperature,
a________cooled Tg,2, OR
.000 ~ afterburner
3700
_____ 3200



Forced-convection-
cooled afterburner













Forced-convection-
Scooled afterburner







S i Approximate peak temperature
limit for present wire cloth
due to oxidation of braze
alloy.


7^---. .--
Minimum ,
practical Maximum practical
cooling- cooling-air pressure
air pressure
Wire-cloth transpiration-
cooled afterburner




3 .04 .08 .12 .16 .20 .24
..:tit flow ratio, WaWg

Figure 17. Comparison of forced-convection and transpiration cooling
applied to same afterburner. Uniform-permeability porous wall. Flight
Mach number, 0.8; altitude 35,000 feet.


400


2400






2000


1600






1200






900









NACA RM E54E25


CONFIDENTIAL


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CONFIDENTIAL

UNIVERSITY OF FLORIDA
R ill fllllx l Tit Ii) l

3 1262 08106 532 7


UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MaRSTON SCIENCE UBRARY
P.O. BOX 117011
GAINESVILLE, FL 32611-7011 USA


CONFIDENTIAL