JANUARY
1967
"~Ii;G1UiC~ : U· Bi~UOl\'E~1\
EFFECT OF GROUND BOARD BOUNDARY LAYER ON AIR CUSHION VEHICLE WIND TUNNE~ TESTS
by
E. K. Garay
,
~~'., ~~.,
L'
cIfi
EFFECT OF GROUND BOARD BOUNDARY LAYER
ON AIR CUSHION VEHICLE WIND TUNNEL TESTS
by E. K. Garay
Manuscript received November
1965
AC KNOWLEDGEMENTS
The author wishes to express his gratitude to Dr. G. N. Patterson for the opportunity to carry out this research at the U.T.l.A.S.
lt was part of the research carried out under U.S.A.F. Contract AF-33(657)-845l of the Control Criteria Branch, Air Force Flight Dynamics Laboratory •
The project was suggested and initiated during a number of dis-cussions with Prof. B. Etkin and
Mr.
G. Kurylowich,to whom the author is particularly indebted for their valuable advice and guidance throughout the investigation. Special thanks go to George Kurylowich for his invaluable assis-tance during the experimental work.Thanks are also due to Messrs. D. Surry, N. Umland and E. Baker for their help with this project.
SUMMARY
Forward speed tests were performed on an Air Cushion Vehicle in the UTIAS subsonic wind tunnel using a fixed ground board. The effect of the ground board boundary layer on the reactions of the vehicle was determined by comparing the test results with previous tethered flight tests conducted in the UTIAS circular track facility. Comparisons with similar reported tests were made to extend these test results.
The nesults did not indicate any noticeable differences between the wind tunnel and circular track results. Analysis showed that differences begin to appear as the forward speed increases and curves the leading edge jet back-wards. The magnitude of the boundary layer effect on the vehicle reactions is shown to depend on the testing technique used.
The suitability of low priced model aircraft engines for low budget powered model testing is demonstrated.
...
i.
I I . lIL IV.v.
VI.
TABLE OF CONrENTS NOTATION INTRODUCTION DESCRIPTION OF APPARATUS 2.1 Wind Tunnel2.2 Force and Pressure Measurements 2.3 Model
2.4 Measurement of Model Fan R.P.M. 2.5 Boundary Layer Measurements EXPERIMENTAL PROCEDURE
REDUCTION OF DATA
4.1 Computational Procedure 4.2 Jet Momentum Flux
4.3 Coefficients EXPERIMENTAL ERROR BISCUSSION OF RESULTS
6.1 Method of Analysing the Force Coefficients 6.2 Lift
6.2.1 6.2.2 6.2.3 6.2.4
Variation of Lift With Forward Speed Intake Effects
Augmentation Curve
Ground Board Boundary Layer Effect on Lift 6.3 Pitching Moment
6.4 Thrust and Drag
6.4.1 Effect of Forward Speed on CD 6.4.2 Drag Components
6.4.3 Ground Board Boundary Layer Effect on Drag 6.5 Ground Board Boundary Lyaer
6.6 Leading Edge Flow Description VII. C ONC LUS IONS
.
'
REFERENCES FIGURES APPENDIX I APPENBIX 11 iv v 1 1 1 2 23
3
3
4
4 4 5 56
6
7
7
7
8
9 10 10 10 11 I I 12 12 13l4
The positive A Aj CD CDo CL CLo CM Cx Cz C IJ. D eq D. l
:f
h H J K 1 L .. J L s m mj MFt
q qj Sb NOTATIONsenses of forces, moments, and angles are indicated in Figure 1.
augmentation factor for lift, LIJ
.,.
area of jets
total drag coefficient, - Cx
profile drag coefficient (i.e. no jet flow)
total lift coefficient, LIJ
component lift coefficient
pitching moment coefficient, M/J Deq
horizontal force coefficient, X/J
vertical force coefficient, - L/J
jet momentum coefficient, J/qSb
equiv~lent diameter of circular model momentum drag, m.V
J
engine frequency, in cps.
height above ground, as defined in Fig. 1
=
h/D eqtotal jet momentum thrust with ambient exit pressure experimental constant
balance lift model lift
statie model lift balance moment inlet mas s flow
model pitching moment base pressure
tunnel dynamie pressure
jet dynamie pressure ,
~pV~
J
V.
J x o "f e p free-stream velocity jet velocitybalance horizontal force model horizontal force angle of attack
jet angle to vertical pitch angle
standard air density, .002378 slugs/cU.ft. true air density
denotes zero forward speed.
....
I.INTRODUCTION
This report presents a study of the feasibility of measuring the forward speed characteristics of an Air Cushion Vehicle in, a wind tunnel by simulating the ground plane with a fixed ground board. The validity of this method of ~imulating Air Cushion Vehicle flight has been questioned because of the presence of a boundary layer on the ground board during forward speed tests. Objections have been raised on the grounds that the jet sheet separates the boundary layer from the ground board ahead of the model altering the flow
pattern over the model ánd that the shape, roughness, and height of the 'effective' ground from the peripheral jet is dependent on the boun4ary layer and is largely unknown. Past investigations into the effect of the jet curtain-ground boundary layer interaction on the reactions of the vehicle have not lead to any definite agreement as to either its qualitative or quantitative nature. Results have
ranged from serious discrepencies between wind tunnel and full scale flight results (Ref. 8) to no differences being noted at all (Ref. 11). To avoid this problem various testing methods have been resorted to, such as the moving belt
technique, circular flight tracks, and free flight models. However, sLnce manytests
.. lhave been and are still being conducted using the conventional fixed
ground-plate, a need exists to check on the adequacy of this method of ground simulation. Possessing the static and dynamic test results of B. Gowans and
G. Kurylowich (Refs. 1 and 2) for a self-powered Air Cushion Vehicle model, Fig.
3,
flownaround the circular track facility at UTIAS, we decided that itwould be of considerable interest to test the same model in the UTIAS subsonic wind tunnel with as few modifications as possible and to compare the two sets of results. In this way the validity of the wind tunnel fixed plate technique can be assessed as a method of determining the forward speed characteristics of the model.
Due to budget and time limitations the objectives ;of these tests were very modest and were not meant to provide an exhaustive examination of the ground simulation problem. Sufficient wind tunnel data was obtained to provide a useful comparison with the tethered flight results of Ref. 2 on the same model, and with similar tests reported in the current literature.
These tests also demonstrate the suitability of low priced model aircraft engines for powered wind tunnel model testing when budget limitations will not allow the instal~ation of expensive electric model motors and their associated electronics.
11. DESCRIPTION OF APPARATUS 2.1 Wind Tunnel
The ACV (Air-Cushion Vehicle) model was tested in the UTIAS sub-sonic wind tunnel outlined in Fig. 2.
test section measuring 48 by 32 inches electric motor drives the fan giving a zero to about 200 fps.
The wind tunnel has a closed octagonal in the principal directions. A 60 hp continuous range of wind speeds from
The ground is simulated by a 46 in. x 40 in. plywood ground board suspended from the ceiling of the test section by four steel rods. It can be rais.ed and lowered by a screw mechanisme The leading edge of the ground
board is shaped like the front portion of a NACA 0020 airfoil and has two symmetrically placed pressure taps to indicate the stagnation point location. The taps enable the boundary layer thickness on the ground board to be mini-mized by locating the board parallel to the wind stream.
2.2 Force and Pressure Measurements
The aerodynamic forces acting on the model were measured using three of the six components available on the existing tunnel balance. This balance uses flat steel springs, whose deflections are converted by differential transformers into voltage signals. The excitation for the differential
trans-formers and the necessary demodulation and amplification of the error signals
is provided by transistorized demodulators whose filtered outputs are trans-mitted directly to a Pace 23lR analogue computer. The computer reduced the signals to give total lift, drag and pitching moment outputs which were re-corded simultaneously using three channels of a Honeywell light spot galvano-meter recorder.
The tunnel dynamic pressure was monitored manually during the tests using a Betz water manometer because of the greater accuracy needed to measure the small air velocities used (0 to 15 fps.). The pressure transducers
feeding tunnel q into the analogue computer were insensitive in this range.
2.3 Model
The model (Acv-4) was constructed and calibrated by B. Gowans (Fig. 3 and Ref. 1). It is made entirely of balsa wood except for the plywood motor mount and is powered by a Cox "Special 15" model airplane engine with a displacement of .15 cubic inches capable of developing .46 horsepower at 18,000 REM. The principal dimensions of the model are summarized in Fig. 4 and Appendix I. The main calibration results by B. Gowans are also included for reference in Appendix I.
The strut-attachment fittings on the model were installed so that the internal flow in the ACV was unaltered from its tethered flight condition. The forward pivot points were placed well forward on the model sa that the tail pivot point was located at the rear of the cowl as shown in Fig. 5. These positions were chosen for strength reasons necessitated by the balsa wood construction of the model and also for convenience of installation. To have properly located the forward pivot points at the model center of gravity would have meant that extensive modifications at the tail would be required to attach the tail strut around the ventral fin.
In Fig.6 the model is shown mounted upside down in the test section. Front and side views of the installation are given in Fig.
7.
A 3/16 in. O.D. neoprene fuel line was run from the engine, out the bottom center of the model, through the ground board and out the side of the tunnel to the externally mounted fuel tank. A pressure line was run parallel to the fuel line from the engine crankcase to the fuel tank. With pressurization to the 16 oz. capacity fuel tank, test runs of over an hour were carried out with maximum engine fluctuations of ± 2cps.204 Measurement of Model Fan R.P.M.
The engine RPM was obtained during the tests using the sound equipment shown in Fig. 8 which also has a block diagram of the electronic equipment. The distinct exhaust "pop" made each cyle by the engine is picked up from the engine noise by the microphone and sound level meter. This signal or series of pulses then passes through a
1/3
octave filter which produces a good periodic signal at the engine frequency. The electronic counter records the number of pulses per second and displays the engine frequency in cps. Thissystem is the same as that used by B. Gowans (Ref. 1) and G. Kurylowich (Ref.2)0
2.5 Boundary Layer Measurements
The boundary layer thickness was measured using a total head boundary layer mouse with nine pitot tubes. Measurements were taken at the model leading edge, mid-station, and trailing edge. The transition point was located by tufting the ground board.
111. EXPERIMENTAL PROCEDURE
Total lift, drag and pitching moment on the vehicle were mea-sured for various angles of attack, forward speeds, and heights above ground. Each test run consisted of a tare, wind off tests, wind on tests, and a final check tare. The following quantities were recorded:
1, the lift measured by the balance
x, the horizontal force measured by the balance m, the moment about the main strut hinges
q, the tunnel dynamie pressure
ex,
the angle of attackh, the height above ground, as defined in
f,
the engine frequencySix test runs were performed as listed below:
1 test for
h/D
= .050 atex
= 0 1 test for h/Deq = .065 atex
= 0 3 "tests for h/nq = .080 atex
= 0 1 test forh/D
eq= .100 atex
= 0eq
Fig. 1
When angle of attack was varied in teach test the forward location of the main strut pivot points caused a change in h about the
ex
= 0 value. For this rea-son each test is referred to by itsh/
D
value atex
= O.eq
The maximum experimental range covered was:
-4
<
ex
<
-4
0.72 in.<
h<
2.71 in. .033o
<
h/D
<
.123 eq-<
q::::. 336 lb/ft2Tests were also performed with the model engine turned off and with the struts alone to measure the magnitude of the resulting forces at various wind speeds and angles of attack.
In each run the engine was started manually, tuned to maximum rpm. byear, and left in this position for the duration of the test unless erratic running developed in which case further adjustments were made to the engine needle valve.
IV. REDUCTION OF DATA
4.1 Computational Procedure
Mean values of lift, moment and longitudinal force were obtained for each test point from the galvanometer traces. These values were put onto punched data cards along with the 0:, h, q,
-:f"
and balance tares corresponding to each case. The data was then processed by a digital program on the University of Toronto IBM7094
computer which gave all the coefficients and data required in this report. The following output parameters were recorded for each of the 118 test points taken: 0:, h, h/D ,q,yC ,
L, M, X, CL' CM' Cx' J, V., q.,~
/
I
/
eq J J~ q. ,1 C , L L •
J IJ. s
The force data was corrected only for strut tare and balance tare. No corrections were made for tunnel wall interference, both because they are unknown, and because the more remote walls are expected to have little effect compared to the adjacent ground board.
4.2
Jet Momentum FluxThe jet momentum flux of the vehicle was experimentally deter-mined under static conditions for an engine speed of 225 cps or 13,500 rpm in Ref. 1. Since the volume flux through a fixed geometry duct is directly propor-tional to the engine rpm, (neglecting effects of Mach number and Reynold's number), Jo and Vj can be determined when the engine rpm is known. By neglecting the small densîty changes, the momentum thrust is corrected to the reference speed of 225 cps by:
J
o 1.84 lbs.
Jet velocity can then be expressed in terms of the exhaust area loading as:
Vjo "
V
(A:~)
fps. AlsoAll velocities and dynamic pressures were obtained using the standard density p =
.002378
slugs/cu.ft.4.3
CoefficientsLift, moment and longitudinal force are nondimensionalized in the conventional way for hovering vehicles using the static value of J. (J was not measured with wind on).
i.e. CL L C
z
A = --
=
J 0 CM J D M 0 eq C=
X X J 0The moment reference point (see Fig. 1) was chosen at the approximate C.G. position of the vehicle. Note that the intake and thrust duct forces are in-cluded in the L, Mand X values.
For wind on conditions the jet momentum is nondimensionalized by a force proportional to the dynamic pressure to give the momentum
coefficient which is written as:
v.
EXPERIMENTAL ERRORC IJ.
=
The chief source of error in the experimental measurements was the balance. It was designed to measure large loads and when used to measure small loads it showed sticking tendencies giving erratic load changes. Zero readings wer'e impossible to get accurately due to hysteresis effects • Although the, springs used in the balance were sensi ti ve enough to measure the small loads being applied, the zero readings at the beginning and end of a test run were not repeatable within the accuracy required of the tests. This was
particularly true in the case of the pitching moment measurements.
Potentiometer drift in the analogue computer load nulling cir-cuit added to the difficulty of obtaining accurate zero or tare readings. This was especially true in the case of the moment circuit which required the greatest accuracy because of the small forces being measured.
The direct aerodynamic forces on the exposed port i ons of the model support struts and the interference effect of the strutson the model are included in the results. The direct aerodynamic forces were measured and found to be negligible at. the tunnel wind speeds used. No attempt was made to measure the strut interference effects or the wind tunnel wall effects.
The tests had to be performed on a fragile model which had been used for a number of years. As a consequence, when comparing the test results with previous tests, errors may have occurred due to bulkheads loosening, re-pairs, propeller changes and engine changes. In addition the inverted position
of the model in the wind tunnel would lead to small changes in the internal dud aerodynamics of the model due to the castor oil from the motor settling in different locations in the model than when it flew in its normal position.
as:
The random error due to the balance limitations was estimated
+ 6%, -6%
+68%, -30%
+
8%, -
%
Note that these are estimates of the maximum possible error.
Other random errors introduced into the system originated from: i) balance calibration
+
'è'/o
ii) frequency measurement
+
'è'/o
iii) dynamic pressure measure-
+ .004
-
psf. mentiV) output recorder
+ 1%
v) angle of attack
+
-
2'vi) flow misalignment
+ 0.5
vii) model changes
+
-
5%
VI. DISCUSSION OF RESULTS
6.1
Method of Analysing the Force CoefficientsSince both H
=
h/Deq and q were altered asa
was varied during the tests, all the coefficient data (CL' CM, Cx) were first plotted versus qjqj so that cross-plots could be made of the coefficients versus H ora
while holding q/qj constant. In this way variation"s of q. during the tests were also accounted for. The results of these plots will be~iscussed
in the sections to f.ollow.in mind:
Throughout the following discus sion four points should be kept
i) the experiment al results are restricted to the low forward speeds required to simulate the circular track flight results
ii) intake effects are present
iii) the stability jets used on this model reduce the augmentation effect and modify the pitching moment characteristics as compared to pure annular jet results.
6.2 Lift
iv) the data is subject to considerable error in some cases due to tbe balance limitations discussed in previous section.
6.2.1 Variation of Lift With Forward Speed
The basic lift data of this investigation. is presented in Fig.
9
for.the four principal height ranges examined. The results are similar tothose found by Greif, Kelly and Tolhurst (Ref. 15) in that the effect of
for-ward speed can be either favourable or unfavourable depending on the particular
values of altitude, at.titude, forward speed, and jet momentum considered.
At positive angles of attack the aerodynamic forces increase
lift, and at negati ve angles they decrease' lift except at the low heights where
with negative angles of attack the air intake is tilted upstream so that the
ram recovery in the intake increases lift initially until the aerodynamic
forces become strong enough to overcome the increased duct efficiency due to
the ram effect.
6.2.2 Intake Effects
A discussion of the effect of the jet intake on the performance
of the ACV is necessary at this point to avoid confusion when comparing the
results of this report with other reports. Examining all the references used
in this report that deal with forward speed tests on ACV's, it was found that
all the externally fed models experienced a lossin lift with forward speed
whereas all the self-propelled models with air intakes experienced a gain in
lift with forward speed except for one test reported by R.E. Kuhn and A. W.
Carter (Ref. 11) where a slight loss in lift was noted. This effect is
dis-cussed in some detail by N.K. Walker (Ref. 12) using test results obtain~d' by
H.
Chaplin of DTMB and by B.M. Carmichael at the Aeronutronics Division of theFord Motor Co. Walker's analysis of the lift of a general ACV in forward flight
indicates that the lift is composed of four components:
i) cushion lift
ii) jet reaction lift
iii) circulation lift
iv) intake thrust recovery lift
Cushion lift and jet reaction lift are common to all ACV's •
Studies have shown conclusively that a loss in base pressure wil~ be experienced
with increasing forward speed provided no other lift components are present
be-sides these two. G.D. Boeller (Ref. 9) h~s shown this loss of cushion pressure
to depend on:
( V
~)
(~)~
or (~)
(~ )~
J J
This "Universal Augmentation Curye" of Boehiler's is shown in Fig. 10 along with
the scatter of the data points the wind tunnel results show that the lift for the UTIAS model remains almost constant with increasing forward speed with
per-haps a very slight drop with increasing velocity. The circular track results are too scattered to make any definite conclusions. However, the ratio of lift to static lift measured on the circular track seems to be generally less than the corresponding values measured in the wind tunnel. The reason the measured wind tunnel. lift remains almost constant with forward velocity is due to the aero-dynamic lift and intake thrust recovery. These effects were not measured by Boehler's model which had air supplied from an external source and was designed to provide no aerodynamic lift.
The circulation lift depends on the upper surface shape of the vehicle and on the angle of attack. Chaplin (see Ref. 12) estimates the aero-dynamic lift to be of the order of 0.2 to 0.5 and has a large effect on the overall efficiency of the vehicle. T.D. Earle (Ref. 18) points out that additional circulation lift is also obtained from the faired fan shroud.
With a jet intake on the model, ram recovery of the dynamic head due to forward speed helps minimize the duct losses and increase the base
pressure. The pressure recovery at the intake will change the flow pattern over the model, slightly altering the aerodynamic lift as well. Walker further points out that if the jet flow from the leading edge is swallowed by the in-take the resulting recirculation can have a small effect on the lift of the vehicle.
To illustrate these effects the Aeronutronics results for an ACV with flush intakes is presented in Fig. 11. The lift rises immediately without any initial drop and the net rise corresponds to an effective CLo value of 0.42 which agrees reasonably with the 0.31 measured with no jet flow but over a dummy cushion. However, the integrated base lift shows a net loss of C~
=
-0.25 so that the aerodynamic lift is in fact much greater than the no flow value from which it may be inferred that the inflow at the intake generated almost as much lift as the body itself.
Failure to recognize the various lift contributions which can act on an ACV has led to many false conclusions in previous reports.
6.2.3
Augmentation CurveThe augmentation curve presented in Fig. 12 was obtained by cross-plotting from Fig.
9
at the hovering condition and at q/qj=
0.2 and 0.4. The static augmentation curve obtained for the same model in Ref. land the circular track results from Ref. 2 are also included.The augmentation curve derived from the wind tunnel data indi-cates that the lift is independent of changes in a and q/qj to within the
.accuracy of the experimental measurements. This dependence of CL on H alone is due to the compartmentation by the stability jets stabilizing the effect of variations of angle of attack and curtain movement due to forward velocities.
Kuhn
&
Carter (Ref. 11) also find the forward velocity to have no effect on the augmentation curve.The circular track lift data seems to be slightly higher than the wind tunnel results , especiallyat the higher q/q. values of 0.34. This agrees wi th the findings of Thunholm (Ref. 14) which
a~e
discussed in Section6.2.4.
The reason the wind tunnel and circular track results are below
the curve found by B. Gowans in Ref. l i s due to changes in the balsa wood
model from wear and tear since Gowans first tested the model when it was new.
The augmentation factor drops below one at negative angles and
the highest height-diameter ratios because of the downward acting aerodynamic
force vector overriding the weakened ground effect.
6.2.4
Ground Board Boundary Layer Effect on LiftP. Colin (Ref. 8) compared tests on a pipe-fed model swung on
a pendulum with wind tunnel tests and found discrepancies as high as 30%
be-tween the two sets of data. He attributed this difference to the flow changes
over the model resulting from the separation of the boundary layer on the
ground board ahead of the model by the jet curtain.
Two of Colin's curves are reproduced in Fig. 13 along with the
wind tunnel and circular track results for the UTIAS model. The agreement
between the UTIAS wind tunnel and circular track data is excellent and does
not show the large differences found by Colin. This shows that the UTIAS model
could be adequately tested in a wind tunnel throughout its entire flight range
without worrying about boundary layer effects.
The fact that the lift does not decrease with increasing forward
speed (decreasing C~) in these curves is again the result of the additional
lift provided by the intake thrust recovery process.
C. Härge G. Thunholm (Ref. 14) compared wind tunnel test on a
pipe-fed model using a fixed ground board and a moving belt ground. Two of the
curves he obtained are shown in Fig. 14 for convenience. The results agree
with the augmentation curve, Fig. 12, disucssed previously in thatthe fixed
ground values are less than the moving ground values initially while q or
l/C~ increases.
A direct comparison is not possible because our range of test
points is not the same as those examined by Thunholm. However, the boundary
layer effect on the wind tunnel results is seen to be greatest at the lower
heights and negligible at the higher height-diameter ratios. The effect also
appears to have a maximum at l/C~
=
10 and decreases from there to l/C~ ~ 17where the two results become identical again.
J.
Williams and S. Butler (Ref. 17) conducted moving belt testson a self-propelled model. The results were encouraging since no change was
found in lift or drag up to the critical speed at least, when the fixed and
moving ground results were compared. Their results at an H
=
0.05 are shownin Fig. 15. The curves show that the significance of the boundary layer effect
should be checked if model tests are to be conducted at or beyond the critical
speed where q
=
Pb or when thefront jet curtain begins to be blown under themodel.
R.E. Kuhn and A.W. Carter (Ref. 11) towed a self propelled
model over ground boards in a towing tank and found negligible differences
between similar tests performed in a wind tunnel. Their test results very
Although the results are encouraging for our particular model, they cannot be assumed to be valid for air cushion vehicles in general. Further
studies at higher forward speeds will still be necessary to clearly define the
limitations of the fixed ground board method of simulation in the wind tunnel.
6.3 Pitching Moment
The basic moment data of this investigation is presented in Fig.
16 for the four principal height ranges examined. Due to the large error
in-volved in the moment measurements, quantitative analysis of the moment data
would serve no purpose towards determining the boundary layer effects since the
differences being sought are so smalle However, the general trends of the
effect of forward speed show up well and will be discussed.
A nose up change in trim as speed increases is seen on all of the graphs. This is due to the ram drag acting on the intake. Without an intake a general nose down tendency would be expected from flow considerations. Any theo-rectical results would be very approximate since considerable correction has to be applied to account for the intake flow.
Referring back to the results of Ref. 17 in Fig. 15 we can see that the nose up moments are reduced b~ nearly 1/3 in the neighbourhood of the
critical speed by the moving ground. According to the moment curve in Fig. 15
very slight differences, if any, should be noted in the speed range within
which the UTIAS model was tested.
6.4 Thrust and Drag
6.4.1 Effect of Forward Speed on CD
The basic longitudinal force data of this investigation is
pre-sented in Fig. 17 for the four principal height ranges tested. The four graphs all show a very steep drag ri se from q/qj
=
0 to 0.01 and then a less rapid rise in drag as the forward speed increases. Because of the amount of scatter involved in the data it could not be determined whether the humping tendency at q/qj=
0.03 in Figs. 17a and 17c was realor just due to balance sticking. Since a similar hump appeared in the drag curve for the SR-Nl reported in Ref. 22 it should not be ignored until further tests are made.In the speed range tested the drag appears to be relatively in-sensitive to height and angle of attack changes. As a re sult , only one smooth curve was put through each of the graphs except for Fig. 17b. This graph had a very large scatter band due to components loosening in the balance during the test.run.
The thrust equals the drag at approximate1
2
q/qj=
0.25 on thelongitudinal force graphs. Using a value of q.
=
5 lb/ft a forward velocityof 10.2 ft/sec results, which is identical to ihe average speed attained by the vehicle in the circular track tests .
6.4.2 Drag Components
The predominant factor in the large drag forces that were
measured is the inlet momentum drag. This is the force required to accelerate the inlet mass flow to the velocity of the vehicle (Di
=
mjV). However, Earl(Ref. 18) and Boeller (Ref. 9) point out that the amount of momentum drag acting on a vehicle depends on the intake design and is always less than the theoretical estimate since a certain amoûllt of the momentum drag is recovered as thrust. Future tests are still needed to determine the amount of thrust re-covered by different geometries.
The remaining drag contributions are the body aerodynamic drag and the suspension system drag. These elements are thoroughly covered in Ref. 18 and are not repeated here.
N.K. Walker (Ref. 12) has shown that the total drag coefficient for an ACV can be represented very well by an expression of the form:
C D
=
C Do+
K CD momentumwhere K can vary between 0.3 and 2.3 depending on the speed parameter and intake design used.
6.4.3 Ground Board Boundary Layer Effect on Drag
Since very little comparative data between fixed and moving ground tests exists in the current literature for the drag of self propelled models, and since no drag measurements were made in Ref. 2, no attemptwill be made to assess the boundary layer effect on the drag using the results of this
investigation. However, the agreement on the forward speed of 10.2 ft./sec between wind tunnel and circular track results is evidence that the drag data
is correct and that little or no differences in drag measurements would be found between the two modes of testing.
The fact that Williams and Butler could not find any appreciable differences in drag using moving and fixed ground boards is probably the re-sult of the much larger momentum drag contribution of their self prope1
1
ed model masking out the small differences found by Thunholm. Thunholm found that the aerodynamic and cushion drag is lower for large jet coefficients (C~) when a moving ground is used. The differences between moving and fixed ground results were greatest at the lowest heights tested.Since no drag derivatives were measured in Ref. 2 the following derivatives are included to complete those found by G. Kurylowich.
CX
H 0.4 to 0.75
CX
e
~ 06.5 Ground Board Boundary Layer
The flow over the ground board was found to be turbulent at the model location at a wind speed of 15 ft/sec. A laminar flow would be expected
over most of the board from theory. This means that an adverse pressure gradient
existed along the board due to misalignment of the board with the flow stream-lines. However, with the model in the tunnel considerable choking of the flow occurred so that the actual effect of a misalignment would be hard to deter-mine without considerable testing.
At 15 ft/sec the transition point from a laminar to a turbulent
boundary layer was found to be 2 to
3
inches ahead of the leading edge of themodel. At the lowest speeds used the entire board had a laminar boundary layer.
Since all boundary layer measurements were made without the cushion jets in operation their value for analysis purposes is questionable.
6.6 Leading Edge Flow Description
A description of the leading edge flow about an air cushion vehicle is presented in this section using the results of Ref. 14 and Ref. 12. The UTIAS model was tested in the subcritical and transitional regions described
in Ref. 12. In the first regime the flow field is dominated by the jet efflux
and the free stream does not conform to the boundary of the vehicle. In the
second regime a stabIe flow pattern is formed with the front jet being split, part being deflected aft, and part penetrating upstream to a stagnation point from where it curves up and back to the leading edge of the model enclosing a separation bubble or vortex as shown in Figs. 18a, band c.
Figure 18 was drawn from some photographs presented in Ref. 14 and shows the flow pattern at the leading edge of a model with fixed and moving
grounds. The effect of the boundary layer on the leading edge flow pattern can
be seen in these diagrams. At the lowest forward velocity (C~
=
2.5) the flowis in the early stages of the transitional regime and shows the separation
bubble ahead of the model. The boundary layer allows the forward jet to
pene-trate further upstream than with a moving ground because of the lower veloeities in the boundary layer. The result is a larger separation bubble ahead of the model with the fixed ground.
As the speed increases from C~
=
2.5 to C~=
1.08 a definitevor-tex is formed ahead of the model. The fixed ground still maintains a larger vortex si ze than the moving ground because of the lower forward resistance near the ground board.
In Fig. 18d the moving belt flow has now reached the second cri-tical velocity where the jet is blown under the model since the statie pressure
ahead of the jet is now greater than the base pressure. However, the fixed
ground flow is in the Poisson-Quinton critical condition in which the curvature
of 'the leading edge jet is zero implying no pressure differential across it.
This is due to some of the leading edge jet flowing forward in the boundary layer and bursting out ahead of the model to reduce the pressure ahead of the jet.
In Fig. 18e (C~
flap regime. The fixed ground
far as the moving belt flow.
0.51) both diagrams show the model in the jet diagram shows that its jet is not blown back as
Thus we see that the flow process is strongly influenced by the moving belt in order to fulfil the no slip condition at the ground surface. The diagrams show that the greatest effect of the boundary layer should be at the higher forward speeds and is substantiated by the results of most of the references used in this report.
VII. CONCLUSIONS
This study has shown that the UTIAS air cushion vehicle could be adequately tested in a wind tunnel without appreci~ble effècts'of the
ground board boundary layer on the reactions of the vehicle.
Tests conducted with self propelled ACV's and externally fed ACV models should not be compared directly unless the effect of the jet intake is understood. Models with intakes will experience an increase in lift with for-ward speed whereas externally fed models without air intakes will show a
de-crease in lift with increasing forward speed. Also the intake momentum drag will be much greater than the remaining vehicle drag.
The ground board boundary layer has little effect on the lift, drag, and pitching moment until the critical speed is reached where the lead-ing edge jet curtain begins to be blown back under the vehicle and the base pressure equals the static pressure ahead of the leading edge jet; At and beyond this critical speed, considerable differences will occur between fixed and moving ground test results. Thus the boundary layer effect on the reactions of the vehicle should be small unless the model enters the jet-flap flight
regime.
Glow plug model aircraft engines can be successfully used for powered model wind tunnel tests of long duration.
1. Gowans, B.W. 2. Kurylowich, G.
3.
Liiva, J.4.
Dau, K.5.
Davis, J.M.6.
Surry, D.7.
Etkin, B.8
.
Colin, P.E.9.
Boehler, G.D.la. Tinajero, A.A.
11. Kuhn,R.E.
Carter, A.W.
12.
Walker, N.K.REFERENCES
Experimental Study of the Aerodynamic Characteris-tics of a Model of an Air Cushion Vehicle in
Hovering Flight. UTIAS TN
74,
Feb.,1964.
The Light-Line Tethering Technique for Determin-ing the Flight Derivatives of an Air Cushion Vehicle. UTIAS Report No. 110.
A Facility for Dynamic Testing of Models of Air-borne Vehicles with Ground Effect. UTIA Tech. Note No.
53,
act.1961.
Characteristics of a Rectangular Wing with a Peripheral Jet in Ground Effect, Part I. UTIA Tech. Note No.
56,
Sept.,1961.
Characteristics of a Rectangular Wing with a Peripheral Jet in Ground Effect, Part 11. UTIA Tech. Note No.
59,
May1962.
Characteristics of a Rectangular Wing with a Peripheral Jet in Ground Effect, Part 111. UTIAS Tech. Note No.
77,
Aug.,1964.
Dynamics of Flight-Stability and Control. John Wiley and Sons, N.Y.,
1959.
Powered Lift Model Testing for Ground Proximity Effects TCEA TN
14,
Oct.,1963.
Forward Flight Characteristics of Annular Jets -Symposium on Ground Effect Phenomena. Princeton
University, act.,
1959.
Comparison of Experimental and Theoretical Design
Parameters of a
6 -
inch Diameter Annular JetModel with a Jet Angle of
_45
0 Hovering in Proximity to the Ground; and Experimental Results for For-ward Flight at Zero Angle of Attack. DTMB Aero Reprt.954,
May,1959.
Research Related to Ground Effect Machines Symposium on Ground Effect Phenomena, Princeton
University, Oct.,
1959.
Some Notes on the Lift and Drag of Ground Effect Machines - Proceedings of the lAS National Meet-ing on Hydrofoils and Air Cushion Vehièles, Sept.
17-18, 1962.
13. Thunholm, C. Härge G. 14. Thunholm, C. Har~e 11 G. 15. Greif, R.K. Kelly, M.W. Tolhurst, Jr., W.H. 16. Otis, Jr., J.H. Goodson, K.W. 17. Williams, J. Butler, S.F.J. 18. Earl, T.D. 19· Davies, H.J. 20. Cockerell, C.S. 21. Eames, M.C. 22. Stanton Jones, R. 23. Loekspeiser, B.
An Experimental lnvestigation of Eight
Axi-symmetrie Annular Nozzles in Proximity to Ground, Part I: Apparatus and Method of Testing. Re-sults from Tests with Wind Off (Hovering), KTH Aero TN 54, June, 1964.
Effects of Moving Ground Surface on the
Characteristics of Annular Jets as Found in Wind Tunnel Tests. Hovering Craft and Hydrofoil, Vol. 3, No. 10, Kalerghi Publications, July 1964. Wind Tunnel Tests of a Circular Wing With an Annular Nozzle in Proximity to the Ground. NASA TN D-317, May, 1960.
Low-Speed Wind-Tunnel lnvestigation of an Annular-Jet Configuration in Ground Proximity. NASA TN D-1779, April, 1963.
Further Developments in Low-Speed Wind Tunnel Techniques for V/STOL and High-Lift Model
Test-ing. RAE TN No. Aero 2944, Jan., 1964.
Ground Effect Machines, AGARDograph 67, Jan., 1962.
General Principles _of the Hovercraft: Hovering Craft
&
Hydrofoil, Oct. 2, No. 7, Kalerghi Pub-lications, April, 1963.An Introduction to the General Principles of Hovercraft, Rovering Craft
&
HYdrofoil, Vol. 3, No. 4, Kalerghi Publications, Jan., 1964.Basic Principles of the Stability of Peripheral Jet Ground Effect Machines. lAS Paper No. 61-71, Jan. 1961.
The Development of the Saunders-Roe Hovercraft SR-Nl - Symposium on Ground Effect Phenomena, Princeton University, Oct., 1959.
Ventil~tion of 24 ft. Wind Tunnel. ARC R
&
M No. 1372.L
M
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Internal Fuel Tank
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