PROPULSION SYSTEM REVIEW

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP89B00739R000400080003-8
• Release Decision: RIPPUB
• Original Classification: T
• Document Page Count: 18
• Document Creation Date: December 16, 2016
• Document Release Date: March 10, 2005
• Sequence Number: 3
• Publication Date: January 1, 1966
• Content Type: REPORT

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A P P E N D I X
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PROPULSION SYSTEM REVIEW
The final U-2R inlet system design will incorporate the
opened up area inlet and a recontoured duct. This design
should provide improved inlet recovery, reduced distortion,
and improved stall margin for the installed engine over the
U-2C inlet system.
The values of inlet recovery Pt2/Pto, used in checking
the installed engine performance estimated by Lockheed, vary
from a minimum of .966 to a maximum of .976 depending on EGT
and altitude. As indicated in the main text of this report,
these values have already been measured in actual flight test
with the opened up area inlet on the U-2C aircraft, and show
nearly a 1% improvement in inlet recovery over and above the
smaller U-2C inlet. These tests did not include recontouring
of the duct which is expected to reduce distortion(hence
improve engine stall margin).
Engine
Performance of the U-2R is based on the improved
performance of the Pratt & Whitney J75-P-13B engine over the
older P-13 version. This improved engine performance is
obtained primarily by an increase in turbine inlet temperature
and engine airflow.
A listing of all the detailed changes to the'J75-P-13
which make up the P-13B version of the engine are included
in Attachment II.
Estimated installed engine performance presented in
Reference 1 was calculated from preliminary engine performance
information released by Pratt & Whitney in December of 1965.
Since that time, a detailed specification has been released
by Pratt & Whitney for the J75-P-13B engine. From the engine
performance presented in this specification, values of esti-
mated installed engine performance were calculated by D/TECH
using the calculation procedure outlined in the specification.
These values were plotted and are presented in Figure A-1,
for comparison with the earlier Lockheed estimates.
The installed performance losses applied to the specifica-
tion engine performance to derive the estimated installed
engine performance' are as follows:
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Inlet Pressure Loss (% of ideal ram
total pressure) 2.
5 to
3.5%
Tailpipe Pressure Loss (% of turbine
discharge total pressure)
1.5%
Bleed Air Extraction (High pressure
compressor discharge) 6.
5 lb
/min.
Horsepower Extraction
spool)
(Low Pressure
10 HP
Horsepower Extraction
spool)
(High pressure
15 HP
These losses_represent a current best estimate and are
probably slightly in excess of those which will-actually occur
in the U-2R. Installation losses will be closely monitored
as the program progresses in order to identify any increase
which might be detrimental to installed engine performance.
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It is strongly recommended that a contract study be
undertaken by Pra Whitney Aircraft to evaluate various
means of reducing from the J75-P-13B engine, 25X1
so that an optimum s stem can be developed which will provide
sufficient I Ireduction, but at the same time minimize
inevitable penalties in weight and currently estimated vehicle
performance. Such a study should involve an evaluation of
various means of cooling the tailpipe and the use of various
types of exhaust ejectors and convectively cooled and transpira25X1
tion cooled plugs. Typical polar plots of the
intensity should be developed for the various svsrems studied
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Performance and weight penalties for each system
should be determined accurately.
A series of four reports summarizing the experience in this
area acquired by Pratt & Whitney to date is listed in Attachment
I. From a cursory review of these reports, it would appear that
either,--an ejector or a convectively cooled plug bullet exhausti2V1
to a low-pressure field nozzle may show the most promise for
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the J75 engine. Transpiration cooled plugs, while providing
a very cool of hot engine hardware, require a very
substantial supply o cooling air at a supply pressure higher.
than engine tailpipe pressure. For a turbojet engine this would
require compressor bleed air. In the case of the J75 engine,,
this would probably require bleeding air from the second or
third stage of.the compressor. A fairly extensive modification
of the engine would be required since a bleed air system would
have to be added, component rematching would be required, and
compressor high speed stall margin would probably be adversly
affected at altitude.
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AIRCRAFT PERFORMANCE
DRAG
The drag polar for the U-2R (Fig. A-2) is a direct
buildup from the drag polar for the U-2C (Fig. A-3), which,
for this analysis, is assumed to be correct as presented by
Lockheed.
Accepting the U-2C drag polar, and taking into consideration
the configuration changes between the U-2C and the U-2R, the
estimated drag increments appear reasonable. It should be
=
understood that the profile drag coefficient change ACD
-0.001 for the U-2R (when related to the U-2C) does not
necessarily signify a lower aerodynamic drag than the aero-
dynamic drag of the U-2C, since any drag coefficient, CD, is
related to a particular reference area, S. (For an airplane,
S is referred to as the "wing reference area".) A measure
of drag is the parameter f=CDS, which is referred to as the
"equivalent flat plate area". Comparing the U-2C and U-2R
flat plate areas at CL=O:
f = 0.0197 x 600 = 11.82 sq. ft. for the U-2C
f = 0.0187 x 1000 = 18.70 sq. ft. for the U-2R
From above comparison it can be said that the U-2R has
58% more equivalent flat, plate area than the U-2C at CL = 0.
The large increase in profile drag, i.e:, Drag at CL = 0
of the U-2R (58%) has been offset somewhat by the effect of,
the larger wing area at the cruise lift coefficient (and
therefore the drag due to lift). The U-2R, because of a
larger wing area, cruises at a lower lift coefficient than
the U-2C. A lower lift coefficient, CL, means a proportionately
lower drag due to lift (induced drag) to total drag, as
compared to the U-2C.
For example: The U-2C maximum power cruise is at CL = 1.0,
and the CD = 0.056. At this cruise condition the total
equivalent fiat plate area is: f=0.056 x 600 = 33.6 sq. ft.
The U-2R maximum power cruise is at CL = 0.75, where the
CD = 0.0334 and the total f=0.0334 x 1000 = 33.4 sq. ft.
It can be said from the above example, that the maximum
power cruise equivalent flat plate area of the U-2R is slightly
lower than the one for the U-2C. However, since the absolute
drag in pounds is a function of the local density of air,
equivalent flat plate area, and of the velocity squared, and 25X1
since the U-2R cruises at a.higher speed than the U-2C, it
consequently has a higher maximum power cruise drag.
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The increase in wing area, for the U-2R versus the U-2C,
has roportionately reduced the induced drag (drag due to
lift), and increased proportionately the profile drag.
The U-2R cruises at a higher speed than the U-2C.
U-2R BUFFET BOUNDARIES
Lockheed has investigated various simple methods to
increase the cruise lifting capability for the U-2 while not
exceeding the buffet boundaries. This investigation was made
to determine which of the following three approaches should
be taken:
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Increasing lift coefficient, increasing wing area, or
increasing Mach number.
Increasing wing area was found to'be the most successful,
since the present airfoil characteristics of the U-2C are very
satisfactory. The U-2C buffet boundary (Fig. A-4) is used as
the basis for defining the U-2R cruise Mach number. Figure
A-4 is based on recent U-2 flight test data.
The maximum altitude and maximum range cruise lift
coefficients, CL, for the U-2R are 0.75 and 0.'5 correspondingly.
At these values of CL, the Mach numbers are 0.72 and 0.73
correspondingly. From Figure..A-4 it can be seen that these
lift coefficients and Mach numbers allow sufficient margin
from stall and Mach buffet.
U-2R CLIMB SCHEDULE
The climb schedule) or climb speed versus altitude, for
the U-2R is similar to that for the U-2C.
--The climb to altitude-is realized at a constant indicated
airspeed (IAS) of 160 knots, up to the altitude at which the
true airspeed is that for which the Mach number is 0.72
(0.70, for---the U-2C). After M = 0.72 has been realized, the
Mach number is held constant for higher altitudes.
Figure A-5 presents the calculated climb speed versus
altitude. The true airspeed values for this graph were
obtained by correcting the 160 KIAS of the climb schedule
for "position error", and "compressibility", in accordance
with the U-2C Flight Handbook.
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U-2R RATE OF CLIMB CORRECTION
From the graph of true climb airspeed versus altitude,
Figure A-5, a rate of climb correction factor (to account
for the acceleration term VdV/dh) was computed for the
corresponding values of altitude.
The rate of climb correction factor QC/C, to account for
acceleration, is presented'in Figure A-6.
The rate of climb correction factor is obtained from
the following equation:
QC/C 0.0886 V dV/dh
where V and dV are in knots and dh in feet.
The corrected rate of climb, R/Cc, is obtained from the
following equation: R /C
R/Cc = 1 + AC /c
where R/C is the uncorrected rate of climb.
It will be noticed from Figure A-6, that the correction
to rate of climb is equal to zero above 52,750 feet of
altitude. This is caused by the climb schedule, which calls
for a constant speed above this altitude (dV/dh=0).
gallons recommended in the U-2C Flight Handbook.)
U-2R SAMPLE CLIMB COMPUTATION
The climb performance in terms of rate of climb, time
to climb, distance gained during climb and fuel used, were
computed by the detailed classical methods. Two climb conditions
from thrust available considerations were calculated; without
and with "Badlands". "Badlands" is defined as the region
between 40,000 and 60,000 feet of altitude, where the Exhaust
Gas Temperature (EGT) of the engine has to be limited to 485?C
to prevent encountering difficulties such as engine roughness,
compressor stall, or flame-out. It is not at all clear that
this EGT limit is necessary for the U-2R.
The sample calculations Figures A-7 and A-8 are without
and with "Badlands" respectively. The sample configuration
presented is for the "Overload" or maximum fuel condition
(T.O.G.W. = 36,750 lbs.)
All climb computations were made by considering the
initial climb weight equal to the take-off gross weight less 25X1
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The thrust values for maximum power climb and respective
fuel flows used in the computations, were those obtained from
Lockheed (Ref. 1).
Rates of climb were computed for. the corresponding
altitudes and climb speeds, and corrected for acceleration.
Time to climb increments were calculated based on the corrected
rate of climb. Fuel used, and distance gained, were calculated
based on the average fuel flows and climb speeds. Graphical
results of these computations are presented in the section
titled ""U-2R Climb Performance With and Without "Badlands".
U-2R CLIMB PERFORMANCE WITH AND WITHOUT "BADLANDS"
The climb performance for the U-2R was computed by the
methods shown in the ""U-2R SAMPLE CLIMB COMPUTATION" section
of this report. Three take-off gross weights were investigated
with and without the effect of the so-called "Badlands".
Overload Take-Off Weight - The overload take-off gross
weight assumes full fuel capacity for the U-2R. This
configuration would give a T.O.G.W. = 36,750 lbs. for the
airplane being 17,400 lbs. heavy in the empty condition.
The climb performance for the U-2R in the "Overload"
condition without "Badlands" has been independently calculated
by Lockheed and D/TECH/OSA. The two separate calculations are
in perfect agreement, and a plot of these results are
presented in Figure A-9.. The rate of climb variation with
altitude is presented only for the D/TECH calculations since
Lockheed did not present these results in Reference 1.
The "Badlands" effects (485?C EGT limitation between
40,000 and 60,000 ft.) on climb are presented also in
"Figure A-9 for comparison. It should be noted that due to
the heavy configuration, and consequently high value of lift
coefficient required during the climb, the drag is also higher.
A high drag means that there is little excess thrust (T-D)
left for climb. The excess thrust available for climb, between
40,000 and 60,000 feet, is further decreased by the 485?C EGT
limitation. For this configuration, the excess thrust becomes
equal to zero at approximately 59,200 feet of altitude. This
means that the airplane will have to "burn-off" weight (fuel)
by cruise-climbing at 485?C EGT until it reaches 60,000 feet
of altitude, where the 485?C EGT restriction is removed. The
weight at 60,000 feet will be approximately 32,455 lbs. At
this altitude, the engine has a "thrust recovery" or, in other
words, the airplane has now an excess thrust available to climb
,again until the drag is equal to the thrust available. The
climb after reaching 60,000 feet of altitude due to "Thrust 25X1
Recovery" has been termed a "Secondary Climb."
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The climb performance for the "Secondary Climb" is
presented in Figure A-10. The "Secondary Climb" is required
for any airplane T.O.G.W. which is required to cruise-climb
below 60,000 feet of altitude (due to the 485?C EGT limitation
between 40,000 and 60,000 feet).
Normal Take-off Weight - The U-2R "Normal" take-off weight
is defined as a T.O.G.W. ^ 30,130 lbs. This configuration
calls for a normal wing fuel weight of 11,930 lbs.
The Normal Take-off Weight configuration climb performance
with and without "Badlands" is presented in Figure A-11.
It will be noticed from Figure A-11, that for this
T.O.G.W. the airplane can climb directly to maximum altitude
(without having to cruise-climb below 60 000 feet) even
with the 485?C EGT restriction between 46,000 and k,000
feet of altitude. This is possible because the weight is
sufficiently low (consequently the drag), and the airplane
has excess thrust (T-D) throughout the "Badlands" region.
Reduced Fuel Take-off Weight - The U-2R "Reduced" fuel take-
off weight is defined as a T.O.G.W. = 26,790 lbs. This
configuration calls for a fuel loading of 9,390 lbs. This
fuel weight was chosen by Lockheed to make a maximum altitude
mission performance comparison between the U-2C and U-2R
for a range of 3,000 n.m. (including credit for descent).
The "Reduced" fuel take-off configuration climb performance
with and without "Badlands" is presented in Figure A-12.
Figure A-12 also shows (as in the other configurations),
that for the airplane with 485?C EGT limitation between
40,000 and 60,000 feet, the rate of climb increases upon
reaching 60,000 feet of altitude. This is caused by the
removal of the EGT limitation for altitudes above 60,000
feet (thrust is recovered since EGT is not restricted to
485?C).
U-2R MAXIMUM ALTITUDE CRUISE
The maximum altitude (maximum thrust) cruise was
estimated for two conditions:
a. For maximum thrust which would allow maximum
altitude cruise above 60,000 feet.
b. For maximum thrusts corresponding to an EGT
limit of 485?C between 40,000 and 60,000 feet of
altitude (or so called "Badlands Cruise").
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Cruise Above 60,000 Feet - The maximum altitude cruise calcu-
lations for altitudes above 60,000 feet (EGT not limited to
485?C), were made by first generalizing the engine data, i.e.,
data to vary proportionally with atmospheric pressure ratio,
so that the cruise could be most efficiently performed at a
constant CL. This generalization was made for 60,000 and
feet of altitude. Table A-1 presents the computed data.
s it can be seen from Table A-1, the data "generalizes"
within approximately 1% or less from the results which are
obtained if individual thrust, fuel flows and altitudes are
used in the calculation of specific range (nautical miles
per pound of fuel). Once the-engine data has been found to
"generalize" the maximum altitude (maximum thrust) cruise
"Range Factor" can be computed as shown in Table A-2. A
maximum altitude cruise speed of I Iwas chosen to
allow adequate buffet margin (see section on "'U-2R Buffet
Boundaries"). Choosing an arbitrary airplane weight of
26,000 lbs., the lift and drag coefficients were calculated
for altitudes between 60,000 and 70,000 feet. A plot of
maximum thrust available and drag versus altitude is shown
in "i ure A-13. From Figure A-13 it can be seen that at
the thrust available equals the drag at 69,500 feet
o altitude. At this altitude, the lift coefficient is
computed for the weight of 26,000 lbs., and found to be
0.742. This value for maximum altitude cruise lift coefficient
compares well with the Lockheed- quoted value of CL = 0.75
(Ref. 1). The Range Factor may now be obtained by multiplying
the specific range (nautical miles per pound of fuel), at
that altitude and speed,'by the airplane weight. The specific
range was obtained from Figure A-14, which is the plotted data
of specific range shown in Table A-l. The Range Factor for
maximum altitude cruise for altitudes above 60,000 feet was
found to be which is in excellent agreement with
the data presented in Reference 1.
Having established the maximum altitude cruise lift
coefficient and an airplane weight associated with an altitude
(26,000 lbs. and 69,500 feet), the entire maximum altitude
cruise above 60,000 feet is now defined, since the airplane
will cruise at a constant W/d , and at a constant Range
Factor.
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The variation of altitude. versus weigght, for maximum
altitude cruise above 60,000 feet (no "Badlands") ed
in Figure A-15. The Range Factor is constant 25X1
irrespective of weight up to a weight of approximately 32,455
lbs. Heavier weights will force the airplane to cruise inside
the "Badlands" (below 60,000 feet). 25X1
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Cruise in the "Badlands"
If a 485?C EGT limitation exists between 40,000 and
60,000 feet, the maximum thrust between these altitudes will
be decreased considerably. The reduction in thrust available
has significant effects upon the performance of the U-2R
when the airplane has a high take-off gross weight. An
extreme example of this condition is the "Overload" take-off
configuration. For this configuration, the maximum thrust
available equals the drag at 59,200 feet of altitude, and
the airplane has to cruise-climb at 485?C EGT limited maximum
altitudes until it has used enough fuel to be at a weight of
approximately 32,455 lbs., at which time it will have reached
60,000 feet of altitude. At 60,000 feet of altitude the
airplane is out of the so-called "Badlands", and the EGT
(and consequently the thrust) can be increased, which enables
the airplane to go through a "Secondary-Climb" (See section
on U-2R Climb Performance With and Without "Badlands" -
Secondary Climb from 60,000 Feet).
To determine the cruise performance in the "Badlands",
the maximum allowable weight for level flight at various
altitudes, between 40,000 and 60,000, was determined. This
was done by matchin the 485?C EGT limited thrust to the
corresponding drag (T-D), and computing the Range Factors
by dividing the fuel flows (in pounds per hour) by the
cruise speed (in knots), and multiplying the results by the
appropriate weights. The cruise performance for maximum
altitude in the "Badlands" is presented in Figure A-15.
As shown in Figure A-15, the Range Factor is not constant
and is much higher than the Range Factor for maximum altitude
cruise above the "Badlands". This is caused by the 485?C EGT
limitation itself which restricts the altitude and fuel flow
to that approaching "maximum range" (lift coefficient for
maximum lift to drag ratio). The variation of altitude and
Range Factor versus weight, for maximum altitude cruise in
the "Badlands" (485?C EGT limitation), is also presented in
Figure A-15. Also shown in Figure A-15, is the "Secondary
Climb" required once the airplane has used sufficient fuel
(therefore reduced weight) to reach 60,000 feet of altitude
(See section on climb).
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U-2R MAXIMUM RANGE CRUISE
The maximum range cruise of the U-2R can be of two
categories (as was also shown for the maximum altitude cruise),
depending on the T.O.G.W.: Cruise above "Badlands", and 25X1
cruise in the "Badlands".*
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For T.O.G.W. configurations that permit the airplane to
reach 60,000 feet of altitude without cruise-climbing in the
"Badlands", the maximum range cruise was estimated at lift
coefficient for maximum lift to drag ratio ((L/D) Max.)
approximately equal to CL = 0.5. At this . t coefficient,
the cruise speed was estimated to be to allow an 25X1
adequate buffet margin (See section on U-2R BUFFET BOUNDARIES).
The range factor for optimum cruise (or maximum range)
was computed as shown in Table A-3. For an arbitrary air-
lane weight of 26,000 lbs., the drag was calculated at
and at altitudes from 60,000 to 70,000 feet. From
the drag values obtained in this computation, the specific
fuel consumption (SFC) were obtained from the Lockheed
furnished installed engine performance (for thrust values
equal to the calculated drag, at apl corresponding 25X1
altitudes).
A plot of the drag and SFQ values given in Table A-3,
was plotted versus altitude, and presented in figure A-16,
Referring back to Table A-3, the fuel flows at the-various
altitudes were obtained by multiplying the SFC with the drag
at corresponding altitudes. The specific range values
(nautical miles per pound of fuel) were obtained by. dividing.
-the airspeed (V) by the' fuel flows. The range fap~ors (R. F. )
is the product of specific range times weight.
From Table A-3, it can be seen that the "Optimum" range
factor is approximately 10,452, and that..the "Optimum Altitude"
for this weight is approximately 61,500 feet, These results
are in very good agreement with values obtained by Lockheed,
and presented,in Reference 1.
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U-2R MAXIMUM ALTITUDE MISSION PROFILES
WITH AND WITHOUT BAD1.,ANDS
Maximum altitude mission profiles were assembled for
the three configurations investigated in the Climb Section
(Overload, Normal and Reduced T.O.G.W.). The missions
were assembled by piecing together the climb and cruise
performance applicable to each configuration. The time,
airplane weight and distance traveled, at any altitude
throughout the mission, were obtained for the airplane
without and with "Badlands" between 40,000 and 60,000 feet
of altitude.
to the manner the missions were assembled for the airplane
Without Badlands - The mission profiles were assembled from
the climb performance curves shown in Figures A-9}A-11,
and A-12, and the Maximum Altitude Cruise data presented in
Figure A-15. For mission profiles for airplane, without
"Badlands", the airplane does not have to make a 1!Spcondary
Climb" from 60,000 feet for any of the three ponflgurations,
and the Range Factor for maximum alt.i.tude crul;so l.s .580
throughout the cruise portion of the mission. ',figures A-1.71
A-18, and A-19, present the mission profiles without "Ba.d-
lands".in terms of time airplane weight, and range versus
---altitude (from take-offj. Two end,alpttudes are shown for
each mission, depending on'the 'fuel reserve allowed,.
With Badlands - The mission profiles wore assembled from the
climb performance curves shown in Figures A-9, A- p, A-11 and
A-12, and the Maximum Altitude Cruise data presened in
Figure A-15. For the "Overload" configuration) the airplane
has to cruise-climb up to 60,000 feet of altitude (cruise in
the "Badlands"). The Range Factor variation with weight, for
this segment of the mission, is shown in Figure A-15. For
the "Overload" configuration, the airplane has to initiate a
"Secondary Climb" (Figure A-10) to maximum altitude for
maximum thrust (once it has reached 6Q,Q00 ft.). After the
airplane has reached the maximum altitude of the "Secondary
Climb" (See Figures A-1Q and A-15), the cruise at maximum
altitude can be accomplished. The Range Factor for this
cruise portion will be at a constant Value of 858Q, regardless
of the weight and altitude.
For the two other configurations ("Normal" and "Reduced"
T.O.G.W.), the airplane weight is qw enough that the impact
of "Badlands" is not as severe as in the "OverloadHH configura-
tion. The airplane will be able to'climb directly to maximum
altitude without having to cruise _el?$mb-? it -the--."1Ba 4l.andsIt.
The assembling of the mission profiles for the "Normal" and 25X1
"Reduced" take-off gross weight egpf figurations, is identical
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without "Badlands". The only exception is that the climb
curves for 485?C EGT limitation between 40,000 and 60,000
feet of altitude (Figures A-11 and A-12) are utilized for
the climb. portion of the mission.
Figures A-20, A-21, and A-'22 present the mission profiles
with "Badlands" in terms of time,'airplane weight, and range
versus altitude (from take-off). Two end altitudes are shown
for each mission, depending on the fuel reserves allowed.
Maximum altitude mission profiles for intermediate take-
off gross weights (between "Overload" and "Reduced") may be
estimated fairly accurately by interpolating between the
climb profiles shown in Figures A-9, A-11, and A-12, and
obtaining maximum altitude cruise performance from Figure A-15.
For some heavy configurations (approaching the "Overload"
condition) with "Badlands", there will be a cruise-climb
portion under 60,000 feet of altitude, and a "Secondary
Climb" from 60,000 feet. The data to be used for these
,portions are presented in Figure A-15 and A-10.
For all cruise portions, the time, fuel, and distance
has to be calculated in small increments (about a 1,000
pounds in change of airplane weight) and then add these
increments.
The small changes in weight represent the fuel used
during that segment. Cruise portions may be calculated in
the following fashion:
Q Fuel = W1 - W2
(NM = R.F.1 + R.F.2 !-)AVE. Wl + W2
Lbs.
R.F. from Figure A-15
Q Distance = (NdkM')AVE x (W1 - W2) N.M.
A Time - Q Distance
Cruise Speed in Knots
Hours.
The maximum cruise altitudes are obtained from Figure
A-15 at weights W1, W2, W3 and so on.
Initial Climb Weight = T.O.G.W. - Take-off Fuel
Total Fuel Used = Take-off Fuel + Climb Fuel + Cruise Fuel 25X1
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Total Time = Time to Climb + Time to Cruise
Total Distance = Distance gained in Climb + Distance
gained in Cruise
End Mission Weight = T.O.G.W. - Total Fuel Used
= Weight Empty + Fuel Reserves
Missions profiles similar to those shown in Figures A-17
through 22 may then be plotted from the calculated data.
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J75-P-13B ENGINE DESIGN CHANGES
The following differences between the current configura-
tion of the J75-P-13 engines and the advanced P-13B
proposal are herewith defined.
1. First Turbine Disk - having increased rupture strength
at elevated temperatures. (Ref. P&W Msg #1778
discussion).
2. First Tubing Blades - P/N 457001, PWA47 coated U700
material, a bill of material feature in J75-P19W
engines.
3. First Turbine Vanes - Change in area only.
4. Second Turbine Vanes - P/N 367052. Upgraded material
to W152, a JT4 commercial part. Area change.
5. Compressor Inlet Case - Revision in annulus area, OD
of case identical to present engine, smaller diameter
of inner passage. Housing covers, gearing and pump
revisions to be made to fit within the smaller area.
Aircraft duct will be affected.
6. Combustion Chambers - Hastaloy X stellite swirlers
as used on J75-P-13,.19W bill of material.
7. Fuel Control - J75-P-13, JFC25-15 control incorporating
revised bellows, linkage clearance revisions.
8.'- Bleed-Valves - P/N 348720, J75-P-17 bill of material,
..to be installed test flight to determine need for
bleeds.
Note: All of the above parts are fully qualified.
Item 1, the first stage turbine disk has already been
,
ordered by our customer as additional protection against
'-inadvertent overtemperature.
The'thrust_ increase will be accomplished largely by an
increase in airflow and turbine inlet temperature. The
airflow 'increase is. accomplished by using a revised inlet
case withTbigger annulus airflow area and vane camber.
The engine turbine inlet temperature increase is made
possible by using new materials in the combustion chambers,
first stage turbine disk and blades, and second stage
nozzle guide vane assembly. We will revise the first and
second stage turbine nozzle guide vane areas to provide
as much stall margin as possible at maximum altitude.
The EGT equivalent,to the increased operating temperature
is 6659C.
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Attachment II to-
As noted in the tabulation above, the engine will require
intercompressor bleed valves for low thrust, descent
conditions because the rematched engine will operate
much closer to stall under these conditions.
The weight increase on existing P-13 engines converted
to the P-13B is about 40 pounds. Since present engines
are.being delivered at nominal weight of 4790#, a nominal
(not guaranteed spec weight). weight of 4830# may be used.
Since new P-13B engines would include additional changes
which are currently used in other J75's, a weight increase
of 70# above present P-13 should be used. Note that
present P-13 specification weight is 4950#.
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