NACA instrumentation has been installed ii the X-J4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling characteristics of the airplane in the clean and gear- down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally U accelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 001-614 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern

Sadoff, Melvin

Sisk, Thomas R

English

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NACA 
RESEARCH MEMORANDUM 
STALL CHARACTERISTICS OBTAINED FROM FLIGHT 10 OF

NORTHROP X-4 NO. 2 AIRPLANE (USAF NO. 46-677)
By Melvin Sadoff and Thomas R. Sisk 
Ames Aeronautical Laboratory

Moffett Field, Calif.
 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
February 2•7, 1050
https://ntrs.nasa.gov/search.jsp?R=19930090543 2020-06-17T04:18:24+00:00Z
NACA RM A50A04	 Il 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
STALL CHARACTERISTICS OBTAINED FROM 
NORTHROP X-1 NO. 2 AIRPLANE (USAF 
	
By Melvin Sadoff and Thomas
	 Sisk
çII 
SUMMARY
	
ItA,	 AO 
NACA instrumentation has been installed ii the X-J4 airplanes to 
obtain stability and control data during the acceptance tests conducted 
by the Northrop Aircraft Corporation. This report presents data obtained 
on the stalling characteristics of the airplane in the clean and gear—
down configurations. The center of gravity was located at approximately 
18 percent of the mean aerodynamic chord during the tests. 
The results indicated that the airplane was not completely stalled 
when stall was gradually approached during nominally U accelerated flight 
but that it was completely stalled during a more abruptly approached stall 
in accelerated flight. The stall in accelerated flight was relatively 
mild, and this was attributed to the nature of the variation of lift 
with angle of attack for the 0010614
 airfoil section, the plan form of 
the wing, and to the fact that the initial sideslip at the stall pro-
duced (as shown by wind—tunnel tests of a model of the airplane) a more 
symmetrical stall pattern.
INTRODUCTION 
The X-4 airplane was constructed by the Northrop Company to provide 
the Air. Force, the Navy, and the NACA with a research vehicle for obtain-
ing stability and control information at high subsonic Mach numbers on 
an airplane having no horizontal tail. 
Several reports have been published presenting limited stability and 
control information on this airplane. Reference 1 presents some 
longitudinal—stability data, stick fixed, with the. center of gravity 
located at about 22 percent of the mean aerodynamic chord. References 2 
and 3 include information on a poorly damped directional oscillation and 
on the lateral— and directional—stability characteristics, respectively. 
Reference 14 presents stability data with the center of gravity located 
at about 19.7 percent of the mean aerodynamic chord; in addition, lateral 
oscillation characteristics and dive—brake effectiveness data are also 
included.
N 
2	 NACA RM A7OA&4
To provide data on the stalling characteristics of this airplane prior 
to accelerated stability tests, flight 10 was made on the X-4 No. 2 airplane 
on September 30, 1949.
SYMBOLS 
Vi	 indicated airspeed, miles per hour 
Az	 normal acceleration factor (the ratio of the net aerodynamic 
force along the airplane Z axis to the weight of the 
airplane) 
M	 Mach number 
q	 dynamic pressure, pounds per square foot 
F	 stick force, pounds 
S	 wing area, square feet 
W	 airplane weight, pounds 
a.	 angle of attack, degrees 
13	 sideslip angle, positive to the right, degrees 
be	 elevon angle, positive downward, degrees 
eLeR
effective longitudinal control angle, degrees 
2 
CN	 airplane normal—force coefficient (wA/qS) 
F/q	 stick—force factor, feet squared 
Subscripts 
L	 left elevon 
B	 right elevon
AIRPLANE 
The Northrop	 research airplane has a vertical tail but no 
horizontal tail. It is powered by two modified Westinghouse 
J-30--WE-7-9 engines and is designed for flight research in the high 
subsonic speed range. Photographs of the x-4 No. 1 airplane, which 
is identical to the test airplane, are presented in figure 1 and a
NACA BM A7QAO4
	
3 
three-view drawing is included as figure 2. The physical characteristics 
of the test airplane are listed in table I. 
INSTRUMENTATION AND TEST PROCEDURE 
Standard NACA instruments were used to record the altitude; airspeed; 
normal, transverse,and longitudinal acceleration; sideslip angle; right 
and left elevon positions; rudder position; yawing and rolling velocities; 
stick force; and elevon and rudder hinge moments as a function of time. 
In addition, the normal acceleration, altitude, airspeed, right and left 
elevon positions,and the rudder position were telemetered to a ground 
station. All the internal records were correlated by a common timer. 
Because of the uncertainty regarding the validityof the absolute magni-
tudes of the hinge-moment data, these quantities are not included in this 
report. 
The airspeed and altitude recorders are connected to the airspeed 
head on the vertical fin. This installation has not been calibrated. 
The test procedure was the same for all the stalls and consisted of 
a straight and level approach using no corrective rudder or lateral con-
trol until the stall occurred. Four gradual stall maneuvers were made in 
nominally unaccelerated flight at a pressure altitude of about 11,000 feet 
with the engine speed set at about 11,000 rpm. Two of these stall maneuvers 
were in the clean configuration and two were in the gear-down configuration. 
In addition, one abrupt stall was made in accelerated flight in the clean 
configuration at a pressure altitude of about 17,000 feet with the engine 
speed set at 13,000 rpm. The Reynolds numbers for these stalls based on 
the mean aerodynamic chord varied from 8.69 X 106 to 9.84 x 106. 
RESULTS AND DISCUSSION 
Typical time histories of the motion of the airplane and controls 
during the stall maneuvers are presented in figure 3. 
The time histories of the gradual stall maneuvers in nominally unac-
celerated flight (figs. 3(a) and 3(b)) show that as stall was approached 
the right wing dropped mildly and slight buffeting occurred. However, 
the pilot commented that the right wing dropped sharply with no warning. 
Recovery was easily and rapidly effected by the use of small down-elevon 
angles. 
The time history of the abrupt stall in accelerated flight (fig. 3(c)) 
indicates fairly rapid right roll at the stall and moderate buffeting, 
although the stall is regarded as being relatively mild. The buffeting 
persisted throughout the recovery. The pilot commented that the charac-
teristics in the abrupt stall were simil ar to those in the gradual stall,
NACA EM A5QA.0 
with the exception that noticeable buffeting, which was considered as a 
positive warning, occurred just prior to the actual stall. Recovery of 
the abrupt stall was rapid and complete following the use of down-elevon 
deflection. 
From consideration of the time-history records and other data, it 
appears that the airplane was not completely stalled in the gradual stall 
maneuvers. The buffeting was milder than would be expected at complete 
stall. The velocity of roll-off at the stall was so mild as to fall, as 
indicated in reference 5, in the category of a stall warning rather than 
a stall. A maximum value of CN of but about 0 . 71 was obtained in the 
gradual stall maneuvers; whereas a value of CN of about 0. 77 was 
obtained in tests of the X-J
 No. 1 airplane reported upon in reference 4. 
A comparison of the maximum values of normal-force coefficient 
obtained in flight with the values of Ci,
	 obtained from three-max 
dimensional and two-dimensional wind-tunnel tests (references 6 and 7, 
respectively) is presented in figure Ij• The gradual stall values are 
considerably lower than the values obtained from wind-tunnel tests even 
though the wind-tunnel values of Reynolds number were considerably lover 
in general. Howver, since the wind-tunnel values were for zero elevon 
angle and the flight values were for elevons deflected about -15 0 , the 
discrepancy is not as large as it would appear. The _150 elevon deflec-
tion corresponds to a ICN of about -.0.08. Thus for zero elevon deflec-
tion the flight values of CN	 would be about 0.79. This value is 
max 
still considerably below the wind-tunnel values that are indicated in 
figure 4 for flight Reynolds numbers, although it is above the value of 
about 0.60 at which, as indicated in figure 5 obtained from reference 8, 
separated flow occurred on the outboard portion of the right wing of the 
model during the wind-tunnel tests. It is believed, however, that the 
effect of separation on lateral trim and stability would not be apparent 
in actual flight until CN was somewhat higher than 0.60, because of the 
higher Reynolds numbers-of the flight tests. The value of CNmax 
which could be obtained in flight would no doubt depend upon the pilot's 
ability to hold wings level by use of the controls after initial separa-
tion occurred over part of the wing. 
The more abrupt stall (fig. 3(c)) was believed to be complete since 
the roll-off was pronounced, and increasing up-elevon at the.stall had 
no effect in increasing C
	 above about o.8. As shown in figure 14., 
'max 
the value of C'max obtained in the abrupt stall is in better agreement 
with the wind-tunnel values, although this comparison also is not strictly 
valid because of the difference in elevon deflection and because of the 
effect on the flight value of CNmax of the relatively rapid change in 
angle of, attack as the stall was approached. The effect of this latter 
factor in the gradual stalls is probably very small. 
The relative mildness of the abrupt stall is probably due to the 
nature of the variation of lift with angle of attack for the 0010-64
NACA RM A5QAO14	 5 
airfoil section, the plan form of the wing, and to the effect of sideslip 
on the stalling characteristics. The lift curves for the 0010-6 airfoil 
section and for a 1/I4 scale model of the 
-4 airplane are shown in figure 
6 for several values of Mach number. Even though breakdown of flow first 
occurs at the tip sections of a swept-back wing, the use of airfoil sec-
tions and plan forms which maintain their lift beyond initial separation 
tends to reduce the magnitude of the rolling and pitching moments that 
are applied. Also, during roll-off at the stall sideslip occurs which 
reduces the asymmetry causing the roll. The effect of sideslip on the 
separated flow conditions over the wing is shown in figure 5. 
The stick-fixed and stick-free longitudinal-stability characteris-
tics near the stall are presented in figures 7 and 8, respectively. The 
data were obtained in steady straight flight with the exception of the 
point at CNmax for the abrupt stall. The stick-fixed data indicate an 
apparent increase in stability as the normal-force coefficient is increased. 
This Increase In stability which persists to values of CN approaching 
the stall is a desirable characteristic, since it reduces the danger of 
inadvertent stalling. The corresponding stick-free characteristics (fig. 8) 
show positive and nearly constant stability up to values of CN approach-
ing those at the stall. 
Although elevon hinge-moment data are not presented because of the 
uncertainty regarding the absolute magnitudes, the data showed that the 
right-elevon hinge-moment variation with CN was similar to that of the 
longitudinal control. 
In summary, it may be noted that the stalling characteristics, as 
measured in flight, were in general agreement with the stalling charac-
teristics predicted from wind-tunnel tests. 
CONCLUSIONS 
The results of stall tests, as determined from. flight 10 on the x-14 
No. 2 airplane, and a comparison of these results with wind-tunnel meas-
urements led to the following conclusions: 
1. The gradual stall maneuvers, made at indicated airspeeds of 
about 1140 miles per hour at approximately a pressure altitude of 17,000 
feet, corresponding to maximum values of normal-force coefficient of 0.71, 
were incomplete stalls. The abrupt stall to about an AZ of 1.60, made 
at an indicated airspeed of about 165 miles per hour and a pressure alti-
tude of about 17,000 feet, was believed to be complete. The maximum 
normal-force coefficient in this case was 0.85. 
2. The maximum normal-force coefficient obtained in the abrupt 
stall compared favorably with the wind-tunnel values. Those, values 
obtained in the gradual stall maneuvers were considerably lower than
6	 oft	 NACA RM A5oAo4 
comparable wind-tunnel values. This discrepancy was believed due to the 
fact that these stall maneuvers were not complete stalls. Conceivably, 
higher values of normal-force coefficient could have been obtained by 
using corrective control to maintain wings-level flight. 
3. The relative mildness of the abrupt stall was traced to the 
flat-top type of lift-curve characteristic of the 001064 airfoil section, 
the plan form used, and to the effect of sideslip on the stalling charac-
teristics. 
1 . The stick-fixed longitudinal-stability data near the stall indi-
cated an apparent increase in stability over that existing at low normal-
force coefficients. This increase was considered desirable, since it 
reduced the danger of inadvertent stalling. The stick-free stability was 
positive and nearly constant over the CN range. 
Ames Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Moffett Field, Calif. 
REFERENCES 
1. Drake, Hubert M.: Stability and Control Data Obtained From First 
Flight of X-4 Airplane. NACA RM L9A31 , 1949. 
2. Williams,'Walter C.: Results Obtained From Second Flight of 
Airplane (A.F. No. 6-676). NACA RM L9F21, 199. 
3. Williams, Walter C.: Results Obtained From Third Flight of Northrop 
X-4 Airplane (A.F. No. 1 6_676). NACA RM L9G20a, 199. 
4 • Valentine, George M.: Stability and Control Data Obtained From 
Fourth and Fifth Flights of the Northrop -4 Airplane (A.F. No. 
1 6_676). NACA RM L9G25a; 199. 
5. Anderson, Seth B.: Correlation of Pilot Opinion of Stall Warning 
with Flight Measurements of Various Factors Which Produce the 
Warning. NACA TN 1868, 199. 
6. Preliminary Report on High-Speed Wind-Tunnel Tests of a 1/14Scaie 
Reflection Plane Model of the Northrop XS-4 Airplane. Rep. No. 
69, Southern Calif. Cooperative Wind Tunnel, Calif. Inst. Tech., 
Pasadena, Calif., Oct. 28, 1948. 
7. Polentz, Perry P.: Comparison of the Aerodynamic Characteristics 
of the NACA 0010 and 0010-6 Airfoil Sections at High Subsonic. 
Mach Numbers. NACA RM A9G19, 1949.
NACA RM A5OAO4	 --
	
7 
8. Pass, H. B., and Hayes, B. B.: Wind—Tunnel Tests of a 1/4Sca1e 
Model of the Northrop XS-4 Airplane. Rep. No. A W'T 46, Northrop 
Aviation Corp., 197.
ru 
L•.J NACA PM A5OAO4 
TABLE I .-  PHYSICAL CHARACTERISTICS OF X-l- AIRPLANE 
Engines (two) ............' . . . . Westinghouse J-30_-WE-.7-9 
Eating (each) static thrust at sea level, pounds . ......1600 
lAirplane weight, pounds 
Maximum (238 gal fuel) .......................7786 
Minimum (10 gal fuel trapped) ....... 6'-o6 
Wing loading, pounds per square foot 
Maximum..............................38.9 
Minimum............................32.0 
Center-of--gravity travel (flight 10), percent M.A.C. 
Gear down, full load .....................19.35 
Gear down, empty .......................16.65 
Gear up, full load ........................19.05 
Gearup, empty ........................16.25 
Height, over-all, feet ........................111.83 
Length, over-all, feet .......................23.25 
Ewing 
Area, square feet ........................200 
Span, feet .........................26.83 
Airfoil section .......................00l0-6-
Mean aerodynamic chord, feet ...................7.81 
Aspectratio	 ...........................3.6 
Root chord, feet ...................... 10.25 
Tip chord, feet ........................1.67 
Taperratio ..........................2.2:1 
Sweepback (leading edge), degrees ...............li-l.57 
Dihedral (chord plane), degrees .................
	 0 
ing flaps (split) 
Area, square feet .. . . . . . . . . . . . . . . .. . . . . . .
	 16.7 
Span, feet . .........................	 8.92 
Chord, percent wing chord	 ......................25 
Travel, degrees	 ........................30 
w
NACA EM A50A014-
TABLE I.— CONCLUDED 
Dive brake dimensions as flaps 
Travel, degrees ........................ ±60
 E levons 
Area (total), square feet ...................17.20 
Span (2 elevons), feet ....................l5.1-5 
Chord, percent wing chord ...................20 
Movement, degrees 
Up............................... 35 
Down.............................20 
Operation ...........Hydraulic with electrical emergency 
Vertical tail 
Area, square feet ......................... 16
 Height, feet	 ......................... 5.96
Rudder 
Area, square feet ...................	 .	 Li
Span, feet 
Travel, degrees ........................ ±30
 Operation ..........................Direct
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NACA RN A5OAO 4	 11 
(a) Front view.
A-14771 
6676	 _____ 
It 11 of 
(b) Side view.
A-14772 
Figure 1.— The Northrop X-4 Airplane.
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NACA RN A5OAC
6676
_ 
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(c) Quarter front view.	 A-14773 
(d) Three—quarter rear view. A- 14774 
Figure 1.— Continued.
13
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NACA RN A50AO4
	
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Figure 2.— Three—view drawing of X
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Stall Characteristics Obtained from Flight 10 of Northrop X-4 No. 2 Airplane (USAF No. 46-677)

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