Document
Restriction/Classification Cancelled NACA RM No. E7AO8 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MEMORANDUM for the Bureau of Aeronautics, Navy Department ALTITUDE WIND-TUNNEL INVESTIGATION OF THE AND 19X8-1 JET-PROPULSION ENGINES 19B-2 ) 19B-8 ) OF II - ANALYSIS TURBINE PERFORMANCE I 19B-8 ENGINE OF THE By Richard P. Krebs and Frank L. Suozzi SUMMARY Performance characteristics of the turbine in the 19B-8 jet- propulsion engine were determined from an investigation of the complete engine in the Cleveland altitude wind tunnel. The investigation to feet and covered a range of simulated altitudes from 5000 30,000 to 0.46 for various tail-cone positions flight Mach numbers from 0.05 over the entire operable range of engine speeds.
The characteristics of the turbine are presented as functions of the total-pressure ratio across the turbine and the turbine speed and the gas flow corrected to NACA standard atmospheric conditions at sea level. The effect of changes in altitude, flight Mach number, and tail-cone position on turbine performance is discussed.
The turbine efficiency with the tail cone in varied from a ma^cimum of 80.5 percent to minimum of 75 percent over a range of engine speeds from 7500 to 17,500 rpm at a flight Mach number of 0.055. Tur- bine efficiency was unaffected by changes in altitude up to 15,000 feet but was a function of tail-cone position and flight Mach number.
Decreasing the tail-pipe-nozzle outlet area 21 percent reduced the turbine efficiency between 2 and 4.5 percent. The turbine efficiency increased between 1.5 and 3 percent as the flight Mach number changed from 0.055 to 0.297.
Restriction/Classification Cancelled 2 NACA RM No. E7AO8 INTRODUCTION An investigation of the altitude performance and the operational characteristics of the 19B-2 19Ba-8, and 19XB-1 jet-propulsion engines has been conducted in the Cleveland altitude wind tunnel at the request of the Bureau of Aeronautics, Navy Department. A mummary of the operational characteristics of the three engines is given in reference 1.
The effects of altitude, tail-cone position, and flight Mach number on the performance of the turbine in the 19B-8 engine are given. Data for computing turbine performance were taken from tests on the complete engine. These tests were conducted over a range of simulated altitudes from 5000 to 30,000 feet and flight Mach numbers from 0.05 to 0.46 for various tail-cone positions.
DESCRIPTION OF TURBINE AND EXHAUST NOZZLE The 19B-8 engine has a single-stage turbine (fig. 1), which delivers approximately 2000 horsepower at rated engine conditions (17 500 rpm at NACA standard atmospheric conditions at sea level).
The turbine rotor has an outside diameter of 16$ inches. The rotor is overhung on the shaft, which is supported by two bearings 6 inches in front of the rotor: a thrust journal bearing located 22- from the center plane of the rotor, and a plain journal bearing inches from the same plane. Both bearings are lubricated by a
5 16
pressurized oil system. The rotor--hub diameter is 10 inches and the thickness decreases from approximately 32 inches at the shaft to 21 inches at the periphery.
The ends of the bulb-type roots of the 32 solid blades are peened over to lock them into the hub. The blade length is inches; the 31 6 chord decreases from 21 6 inches at the root to 132 inches at the tip.
The radial tip clearance for the wheel is 0.078 inch.
The turbine nozzle (fig. 2) consists of 44 solid, equally spaced blades that fit into slots in two shroud rings. The outer shroud ring
is 161 6 inches in diameter and the inner shroud ring, 916 inches, The
blades are of constant width and thickness and are so set that the chord line makes an angle of approximately 41 0 with the plane of rotation of the turbine,, Restriction/Classification Cancelled CONFIDENTIAL 3 NACA RM No. E7AO8 The exhaust nozzle has a movable inner cone actuated by a hydraulic cylinder mounted on the outside of the combustion chamber.
Moving the inner cone from the "in" position to the "4-inches-out" position decreases the tail-pipe-nozzle outlet area, measured in a plane perpendicular to the center line of the engine, from al_.Troxi- mately 135 to 106 square inches-.
INSTRUMENTATION The 19B-8 jet-propulsion engine was instrumented at the stations
the instrumentation at those stations
shown in figure 3. Details of from which data for this report were taken were as follows (all instrumentation was unshielded unless otherwise noted): , station 1, was instrumented with sixteen total- The cowl inlet .
pressure tubes, nine static-pressure tubes ., and four iron-constantan thermocouples. The arrangement of the pressure tubes and thermo- couples is shown in figure 4.
The instrumentation at station 2, the compressor inlet, was carried by four rakes set at 45 0 to the center line of the engine.
These rakes contained twelve total-pressure tubes and eight iron- constantan thermocou-,)les. Four wall static-pressure orifices were also provided at this station.
At station 3, the compressor outlet ., the instrumentation was mounted on four rakes set at right angles to one another. The rakes contained six total-pressure tubes, two static-pressure tubes, and four iron-constantan thermocouples. Four wall static-pressure orifices were also provided.
At stat i on 4 2 the turbine inlet ., total pressure was measured by a Westinghouse integrating rake. Viewed from the front of the engine, the location of the rake was in the lower right quadrant 17.5 0 from the horizontal diameter and 3 5 inches ahead of the turbine. The rake is shown and its position indicated in figure 5.
The instrumentation at the turbine outlet, station 5, consisted of twelve shielded total-pressure tubes and eight chromel-alumel thermocouples mounted as shown in figure 6. Four wa l l static-pressure orifices were also installed at this station.
CONFIDENTIAL 4 +CONFIDENTIAL NACA RM No. E7AO8 RANGE OF INVESTIGATION The engine performance was investigated over a range of simulated altitudes from 5000 to 30,000 feet and flight Mach numbers from 0.05 to 0.46 with the engine tail cone in . over the entire operable range of engine speeds. The range with the tail cone 4 inches out was restr.cted by the necessity of keeping the turbine-inlet temperature below 1880 0 R.
The tunnel temperatures were maintained at approximately NACA standard atmospheric conditions for each altitude.
SYMBOLS The following symbols and necessary values are used in the com- putations; A cross-sectional area, square feet specific heat at constant pressure, Btu per pound OR g ratio of absolute to gravitational units of mass, 32.17 J mechanical equivalent of heat, 778 foot-pounds per Btu N engine speed, rpm P total pressure, pounds per square foot p static pressure, pounds per square foot T total temperature, OR T i indicated temperature, °R t static temperature, OR V velocity, feet per second Wa air f low, pounds per second Wf fuel flow, pounds per second gas flow, pounds per second W M thermocouple impact--recovery factor, 0.85 y ratio of specific heats CONFIDENTIAL CONFIDENTIAL 5 NACA RM No. E7AO8 pressure-correction factor, P 4 /2116; turbine-inlet total Ares- sure divided by NACA standard atmospheric pressure at sea level t turbine efficiency e4 temperature-correction factor, y4T4 /(184 X 519); product of y and turbaned-inlet total temperature divided by product of y and total temperature at NACA standard atmospheric conditions at sea level density, slugs per cubic foot P Subscripts; C compressor t turbine 0 free-stream or ambient 1 cowl inlet 4 turbine inlet 5 turbine outlet METHODS OF COMPUTATION Gas F low The gas flow was found by adding the fuel flow to the air flow; W g =Wf +Wa Air flow was determined from pressure and temperature measure- ments at the cowl inlet, station 1, by F_ y1_1 p yl Wa = p 1A171 = p lAl 2Jg c p' 1 ti 1
-
1p1
CONFIDENTIAL 6 CONFIDENTIAL NACA RM No. E7AO8 Temperatures Static temperature was calculated from the indicated temperature by T i 7 -1 t V Y
l+a ^ -1
The thermocouple impact-recovery factor a was determined from cali- bration tests and was found to be 0.85.
Total temperature was determined by the adiabatic relation 74 1 Y
= (
P
p)
E
Efficiency
Turbine efficiency, based on the total-pressure ratio across the turbine with bearing friction losses, power required to drive the accessories, and compressor and turbine thermal losses neglected, was
defined by
ATt nt Yt P5
i
T4 1 -k P4
L
The enthalpy rise across the compressor was assumed equal to the enthalpy drop across the turbine and the total-temperature drop across
the turbine was obtained from
dTc cp'c
^ Wf
OT t
+ W cp , t
al
The total temperature at the turbine inlet was computed from T 4 = AT + T5 CONFIDENTIAL No. 7 NACA RM E7AO8 CONFIDENTIAL The value of 7t was determined from a curve using values of fuel-air ratio and an average of the turbine-inlet and turbine-outlet temperatures. The curve was based on a hydrogen-carbon ratio of 0.170 and a combustion efficiency of 98 percent.
y-1 P 7 S Turbine efficiency was determined from a plot of` 1 -) against ATt 4 . The values of efficiency presented were obtained /T from a curve faired through the data points of this plot.
RESULTS AND DISCUSSION The method of presentation of the turbine data used in this report follows the same pattern as that used in reference. 2. Turbine performance is shown by plots of turbine pressure ratio P4/P5 against corrected engine speed N1 ^/8 gas flow Wg "-"'04 , and 4 and corrected s474 1.4 turbine efficiency rat plotted against turbine pressure ratio. The _plot of turbine pressure ratio against corrected gas flow is designated the turbine operating line.
Each run was made at constant flight Mach number. Because a screen had been inserted at the cowl inlet to protect the engine, the ram-pressure ratio P2 /pO decreased as the air flow and the engine speed increased. Curves of the compressor-inlet ram- pressure ratio are given on most of the performance curves.
Effect of Altitude Two plots relating turbine pressure ratio with corrected turbine speed and corrected gas flow for three different simulated altitudes are shown in figure 7. These data were taken with the engine tail cone in at flight Mach numbers between 0.142 and 0.153. The small changes in flight Mach number and ram-pressure ratio were considered insufficient to impair the conclusions drawn on the effect of altitude on turbine roerf ormance. Figure 7 shows that a change in simulated altitude from 5000 to 15,000 feet had little effect on the two rela- tions. At a simulated altitude of 25,000 feet, the corrected turbine+ speed decreases and the corrected gas flow increases from those values obtained at 5000 or 15,000 feet. These differences are probably caused by erratic combustion rather than by any changes in the turbine performance at the high altitude.
CONFIDENTIAL 6 CONFIDENTIAL 1TIAL NACA RM No. E7A08 7t-1 yt
The quantity 1 - 5 is plotted against ATt /T4 for three
C4 simulated altitudes in figure 8. The turbine efficiency is calculated by dividing ATt /T4 by the adiabatic temperature-drop factor. The data ?p oints for simulated altitudes of 5000 and 15,000 feet show no change in turbine efficiency with change in altitude.
Effect of Tail-Cone Position The effect of changes in the tail-cone position on turbine per- formance is shown in figures 9 and 10. The data presented were taken at a simulated altitude of 15,000 feet and a flight Mach number of 0.147. A decrease in tail-pipe-nozzle outlet area decreases the air flow, raises the compressor pressure ratio, increases the temperature differences across both the compressor and turbine, and raises the turbine-inlet temperature. Although the turbine work per pound of gas increases, the turbine--inlet temperature increases sufficiently to cause a drop in the turbine pressure ratio. Typical values of turbine-inlet temperature, turbine pressure ratio, and temperature drop across the turbine for several engine speeds are shown in the following table; Tail-cone Engine .Turbine- Turbine Turbine pressure position speed inlet tempera- (rpm) ratio temper- ture drop ature (OR) (OR) In 141400 1427 1.890 165 4 in. out 1642 1.832 168 In 15,517 1478 1.965 188 4 in. out 1816 1.890 198 In 16,514 1578 2.066 210 4 in. out 2025 1.945 221 When the tail cone is moved out, the change in the corrected turbine speed is greater than the change in the pressure ratio; thus the curve of the pressure ratio plotted against the corrected turbine speed shifts to the left (fig. 9(a)).
CONFIDENTIAL NACA RM No. E7AO8 CONFIDENTIAL Reference 2 established that, when the velocity in the turbine nozzle was sonic (that is, the pressure ratio across the nozzle was y4)1 greater than critical, 1.84 to 1.89, depending on the value of the turbine operating line was affected. only by changes in y4.
Because the turbine in the 19B-8 engine is a reaction design, the static-pressure drops through the rotor and the turbine pressure ratio may be considerably higher than the critical value before sonic flow is established in the turbine nozzle. When the tail cone is moved out, the turbine-inlet temperature is increased by the addition of more fuel and y4 is reduced by both the increase in temperature and fuel-air ratio. The amount of increase in the corrected gas flow caused by this reduction in y 4 was computed to be 0.3 percent. The turbine operating line shifts toward the higher corrected gas flows by about 0.8 percent (fig. 9(b)).
Turbine efficiency is plotted against turbine pressure ratio for two different tail-cone positions and a flight Mach number of 0.147 in figure 1.0. The turbine efficiency varies from a maximum of 80.5 percent to a minimum of 76.5 percent over a range of turbine pressure ratios from 1.2 to 2.1 corresponding to engine speeds from 7500 to 17,500 rpm with the tail cone in. The efficiencies are from 2 to 4.5 percent lower when the tail cone is moved out and the tail- pipe-nozzle outlet area is thus decreased 21 percent.
The curves in figure 10 have shapes different from efficiency curves for the turbines of other ,jet-propulsion engines investigated in the altitude wind tunnel. For the other turbines, the efficiencies are low at low turbine pressure ratios, rise to a maximum near the critical turbine pressure ratio, and then decrease as the pressure ratio is further increased. For such a turbine, the operating lines lie to the left of the region of maximum turbine efficiency. (See fig. 8, reference 2.) For the turbine in the 19B-8 engine, however, the operating lines are to the right of the region of maximum turbine efficiency (fig. 11). If the range of tests were extended to higher or lower turbine pressure ratios, the efficiencies would begin to decrease. The efficiency curve for the tail cone 4 inches out exhibits this tendency at the low turbine pressure ratios (fig. 10)).
Effect of Flight Mach Number The effect of changes in flight Mach number on turbine-inlet temperature, turbine pressure ratio, and the temperature drop across the turbine for several engine speeds is shown in the following table; CONFIDENTIAL 10 NACA RM No, E7AO8 CONFIDENTIAL Turbine Flight Engine Turbine- Turbine inlet pressure tempera- Mach speed number ratio ture drop (rpm) temper- ature (OR) (OR) 0.055 14 5 400 1510 1.860 161 .1.42 1472 1.838 164 ,297 1420 1.800 160 0.055 15,517 1567 1.955 187 .142 1510 1.976 187 .297 1509 1.930 1 82 0.055 17 1 000 1619 2.108 222 .142 1635 2.082 220 f .297 1616 2 .070 217 ( Although the table contains small inconsistencies, the data indicate that probably an increase in flight Mach number decreases the turbine- inlet temperature, the turbine pressure ratio, and the temperature drop across the turbine. Because the temperature drop across the turbine decreases, the work done by the turbine per pound of gas also decreases as the flight Mach number increases. Low turbine-inlet temperature increases the corrected turbine speed and the curves in figure 12(a) are shifted to the right as the flight Mach number is increased.
The shift in the three operating lines, shown in figure 12(b), is attributable to faulty measurements, probably of the air flow.
Air flow was measured at the cowl inlet by the rakes shown in fig- ure 4. No total-pressure tubes were situated near enough the outside of the inlet to detect a boundarylayer. A change in boundary-layer thickness of about 0.2 inch would account for the shift in the oper- ating lines shown in figure 12(b).
With the tail cone in ., the turbine efficiency varies from a maximum of 80.5 percent to a minimum of 75 percent over a range of turbine pressure ratios from 1.2 to 2:1 corresponding to engine speeds from 7500 to 17,500 rpm at a flight Mach number of 0.055 and is increased by an increase in flight Mach number. This effect is illustrated in figure 13 where turbine efficiency is plotted against turbine pressure ratio for several flight Mach numbers. The data were taken at a simulated altitude of 5000 feet with the tail cone in. Efficiencies are increased between 1.5 and 3 percent when the flight Mach number is increased from 0.055 to 0.297.
CONFIDENTIAL NACA RM No. E7AO8 CONFIDENTIAL SUMMARY OF RESULTS An investigation of the turbine performance of the 19B-8 jet- propulsion engine over a range of simulated altitudes from 5000 to 30,000 feet and flight Mach numbers from 0.05 to 0.46 gave the following results; 1. The turbine efficiency (uncorrected for bearing-friction losses, compressor and turbine thermal losses, and the power required to drive the accessories) varied from a maximum of 80.5 percent to a minimum of 75 percent as the turbine pressure ratio changed from 1.2 to 2.1 corresponding to a change of engine speed from 7500 to 17,500 rpm at a flight Mach number of 0.055 with the tail cone in.
2. Turbine efficiency was unaffected by changes in altitude up to 15,000 feet.
3. Turbine efficiency decreased from 2 to 4.5 percent when the tail-pipe-nozzle outlet area was decreased 21 percent, 4. Turbine efficiency increased from 1®5 to 3 percent when the flight Mach number was increased from 0.055 to 0,297.
Aircraft Engine Research Laboratory, National Advisory Coirzittee for Aeronautics, Cleveland, Ohio.
R ichard P . K^-e`bs, Physicist.
Frank L. Suozzi, . i Aeronautical Engineer.
Approved; Newell D. Sanders, Mechanical En ;_neer.
A Silverstein, Aeronautical Engineer.
lrp CONFIDENTIAL 12 CONFIDENTIAL NACA RM No. E7AO8 REFERENCES 1. Fleming, William A.: Altitude-Wind-Tunnel Investigation of the Westinghouse 19B-2, 19Be8, and 19XB-1 Jet-Propulsion Engines.
I - Operational Characteristics. NACA MR No. E6EO6, Bur. Aero., 1946.
2. Krebs, Richard P., and Hensley , Reece V.: Altitude-Wind-Tunnel Tests of the General Electric TG-180 Jet Propulsion+amine.
m Analysis of Turbine Performance. NACA MR No. E6F 11, Army- V Air Forces, 1946.
CONFIDENTIAL
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F i g u r e I . - S i n g l e - s t a g e t u r b i n e r o t o r a n d s h a f t o f 195-8 j e t - p r o p u l s i o n e n g i n e .
' F|g^ 2 CONFIDENTIAL NACA RW Mo. E7kO8 NACA C- 14566 m~es-*a Figure 2. — Turbine nozzle of 1913-8 jet—propulsion engine.
CON F1 DEN T|AL n n s n m v n co c a C7 f^ - n O O e rn C -.I
m
z
A Lion 7 — n r Station 5 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 8 jet - propulsion - engine installation showing measuring Figure 30 - Side view of 19B- stations, T^ cL} NACA RM No. E7 A08 Fig CONFIDENTIAL 4 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
* Total—pressure tube
Static—pressure tube
Thermocouple
Figure $— Location of instrumentation at cowl inlet,,
inches ahead o
station l,, Plane of survey ., front flange of oil cooler.
CON F I DENT 1 AL NACA RM No. E7AO8 Fig. 5 CONFI DEN TI AL NATIONAL ADVISORY CONWITTE£ FOR AERONAUTICS
Center line of unit
CON F I DEN TI AL NACA RNI NO. E7A08 F g, 6 CONFIDENTIAL
0 Total—pressure tube (shielded)
11 Wall static —pressure orifice
Thermocouple
Figure 6,, — Location of instrumentation at turbine outlet,
station 5,, Location of rakes, 5a inches behind rear
flange of combustion chamber.
CONFIDENTIAL F i 9 , 7 I AL NAC.A RM NO. E7AOd CO N F I D EN T 1.10 NATIONAL ADVISORY 0 0 COMMITTEE FOR AERONAUTICS C11,00 ai I I I I a) Jo 2.2 JAq Simulated Flight altitude Mach (ft) numberi 0 5,000 0.142 1.4 15,000 .147 V 25 5 000 ^11 5 3 U> 1^01 — 65000 8,000 105 . 000 4,000 4 , rpm Corrected turbine speed, N/;57 9. (a) Corrected turbine speed, s.
U) W c.
1 ^ 16 18 20 Wg F64 , Corrected gas flown, lb/sec 64Y4/1 4 (b) Corrected gas flow.
Figure ?,— Effect of altitude on turbine pressure ratio as Tall cone in, Turbine function of two corrected parameters, speed and gas flow corrected to NACA standard atmospheric conditions at sea level, CON F I DENT I AL s Fi g , 8 E7 AO8 CON FI DEN TI AL N ACA RFC htO 1.10 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1.00 m1 I ^ a o ^ H , ro M a ^.
v v ^v
V^Zzl
X15 i, $13 i ^--/ Simulated Flight I a7 titude Peach „09 (ft) number O 5,000 08142 O 15,000 .147 V 25,000 .153 ^ X 07 ®05 X11 X13 .,15 $07 .09 X 03 X05 ATt/T4 cone used in computing turbine efficiency, Tail Figure 8,-° Plot ins CONFI DEN TI AL FI E7AO8 CONFI DENTI AL NACA RN{ NO, 9 , 9 1.10 a 0 NATIONAL AOY ISORY COMMITTEE FOR AERONAUTICS o \ c " 1 8 00 a s.
a. o = 4; „90 2,2 iE cd cd f r^ t, I o, 1 1.8 Tail—cone position 0 In q 4 in, out 1.4 i 8,000 10,000 6,000 0. 4,000 rpm Corrected turbine speed, N/ F41 Corrected turbine speed .- (a) 9:.
v
i-
a 2,2 a I ^^ P.
a^ s; i s.
^ 1„8 1.4 1^0 14 16 18 _20 Corrected gas flow,, lb,/sec b wgs94 ' 4 4 (b) Corrected gas flow, Figure 9 R — Effect of tail —cone position on turbine pressure ratio as function of two corrected parameters, Simulated altitude, 15,000 feet; flight Mach number, 0,147, Turbine speed and gas flown corrected to NACA standard atmospheric conditions at sea level, CONFI DENTI AL, E7A08 F i g , 10 NACA R a h N O . CON I AL F I DEN T a of U) 04 CL 0 s.
Arl 1"0 1,2 1.4 1^6 1.8 2.0 2,2 /P 5 Turbine pressure ratio, P4, Effect of tall —cone position on turbine efficiency, Figure 10, — Simulated altitude, 15,000 feet; flight Mach number, O^147^ T i CON F I DEN AL I RM NO, E7 AO 8 N ACA CON F1 DEN T I AL F i g . I I ADVISORY N ATIONAL AERONAUTICS TTEE FOR L COn Tail—cone position 2.2 In 4 in, out 2,0 1^8 LO a.
1.6 .H ro 1.2 E_ 14 15 16 17 18 19 20
la 4
g^
Corrected gas flow,
lb/sec
'r4y7f7'
Fi,gure ll.—
Efficiency contours plotted on two turbine operating
lines,, Simulated altitude, 15,000 feet; flight Mach number,
0,147,, Gas flow corrected to
WACA standard atmospheric
conditions at sea level, CON FI- DENT 1 AL NO. F7AO8 NACA RM CONFIDENTIAL F 12 (D 0 110 L G NAT SONAL ADD ISORY V) w 1 COMMITTEE FOR AERONAUTICS ICS L, w 1,00 c c ,90 2.2 (Z co c- Flight Mach 1„8 nu-nher 0 0,055 q 4142 O ,297 I a 4
a 1.0 1 I I I I
8,000 10,000 4,000 6,000 Corrected turbine speed, N/^ 4, rpm (a) Corrected turbine spee9.
m z.
Q, L V.
(h c 2° v L h H lab 1,4 1 1 1 1 1 1 1 1 1,0 1
1_1 1
16 18 14 20 W F® Corrected gas flow, g 4 4 , lb/sec 64y4/1 ,4 (b) Corrected gas flow, Figure 12, — Effect of flight Mach number on turbine pressure ratio as function of two corrected parameters. Simulated altitude, 5000 feet; tail cone in, Turbine speed and gas floss corrected to NACA standard atmospheric conditions at sea level,, CON F I DENT I AL Restriction/Classification Cancelled N ACA RM No, E7AO8 Fi 13 9^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Flight Mach number —0.055 — — — -- ®142 --- .297 (U 01,110 ^ na aP^ 0) w 1 00 P4 o 8 .11 ro .90 v U ^a 78 `tea U a-1 U N h 1 '70 .1 E-' 1 1 1 1 1 1 681 L_ 1— 1 1 --1 1„0 1^2 1^4 1.6 1.8 2,,0 2,2 Turbine pressure ratio, P^/P5 Figure 13.-- Effect of flight Mach number on turbine efficiency, Simulated altitude, 5000 feet; tail cone in, Restriction/Classification Cancelled R