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NASA Contractor Report ^ , _^ 



Helicopter Tail Rotor Blade-Vortex Interaction Noise 



Albert R. George and S.-T. Chou 



Sibley School of 

Mechanical and Aerospace Engineering 

Upson and Grumman Halls 

Cornell University 

Ithaca, New York 

(**S1-CB-183178) BELICCPfEE lill ECIOB N89-16167 

EiAD£-¥CBIEX INIIfiCTICS «C1£I final 
Ifchnical Bepcit, 1 F€t. 1S86 - 3 1 Bar- 1987 
(Ccrnell Dniv.) 26 p CSCL 20A Unclas 

G3/71 01912<}3 



Final Technical Report 

Prepared for NASA Ames Research Center 

Grant NAG 2-379: Helicopter Tail Rotor Noise 

Principal Investigator; A. R. George 

Period; February 1, 1986 - March 31, 1987 



Contents 

Summary 

Introduction 

Acoustics of Blade- Vortex Interactions 

Illustrated Examples and Discussion 

Additional Examples 

Conclusions 

References 

Publications Issued During the Research 

Figures 



Summary 

A study is made of helicopter tail rotor noise, particularly that due to the interactions 
with main rotor tip vortices. This report summarizes the present analysis, the computer 
codes, and the results of several test cases. 

Amiet's unsteady thin airfoil theory is used to calculate the acoustics of blade-vortex 
interaction. The noise source is modelled as a force dipole resulting from an airfoil of 
infinite span chopping through a skewed line vortex. To analyze the interactions between 
helicopter tail rotor and main rotor tip vonices, we developed a two-step approach. In the 
first step, the main rotor tip vortex system is obtained through a free wake geometry 
calculation of the main rotor using CAMRAD code developed by Johnson. This allows us 
to define each individual tail rotor blade-vortex interaction. In the second step, acoustic 
analysis takes the results from the aerodynamic interaction analysis and calculates the far- 
field pressure signatures for the interactions. 

It is found that under a wide range of helicopter flight conditions, acoustic pressure 
fluctuations of significant magnitude can be generated by tail rotors due to a series of 
interactions with main rotor tip vortices. This noise mechanism depends strongly upon the 
helicopter flight conditions and the relative location and phasing of the main and tail rotors. 



Introduction 

During the forward flight of helicopters, tail rotors are operated in a very 
complicated and gusty environment containing wakes from the main rotor, fuselage, 
stabilizer, tail boom, etc. These wakes have significant influences on both the harmonic 
and the broadband noise tail rotor generates (reference 1-3). 

Because of the size of the main rotor wakes, they are the most important sources of 
tail rotor noise. The main rotor wakes, which include the overall downwash flow field, the 
shedding vortices from blades' tips, and a number of other components have been studied 
in our past research (references 1 and 3). We were able to show that the overall downwash 
(mean and turbulent wake) of main rotor are important sources of the tail rotor harmonic 
and broadband noise. The main rotor tip vortex does not seem to be important to the tail 
rotor broadband noise. But it remains a question to be answered whether main rotor dp 
vortex is important to tail rotor harmonic noise and is the subject of this present study. 

Because the main rotor tip vortices are strong and concentrated, they are capable of 
generating significant velocity perturbations to the flow field seen by the tail rotor. These 
strong velocity perturbations can result in large unsteady loading fluctuation on the tail rotor 
blades; significant noise can therefore be generated. Figure 1 shows the geometries of tail 
rotor blade-vortex interaction and the main rotor blade-vortex interaction. These 
interactions are similar in a sense since, in both cases, the blades are interacting with the 
vortices shed from the main rotor. However they are different because of the relative 
orientations. The main rotor blade-vortex interactions generally occur with vortices of 
near-parallel orientations (Figure lb). They are basically transonic and have been treated 
by several investigators (references 4 - 8). The tail blade-vortex interactions, on the other 
hand, generally occur with vortices of near-normal orientations to the tail rotor plane 
(Figure Ic). 

In the present study, the acoustics of tail rotor blade-vortex interactions is treated 
by the analysis of Amiet (references 9 and 10). The acoustic analysis requires input 
informations such as the ingested vortex properties and the vortex-blade geometric 
orientation. In the present study, these inputs are calculated based on a main rotor free 
wake geometry analysis developed by Johnson and Scully (references 1 1 and 12). 



Acoustics of Blade- Vortex Interactions 

In the present study, tail rotor blade-vortex interaction is treated as a 2-D blade 
vortex interaction problem (a flat plate of infinite span chopping through a moving skewed 
vortex). The analysis was originally developed by Amiet (references 9 and 10), who 
assumed that the noise source is an unsteady loading fluctuation (i.e. dipole) when a thin 
airfoil chops through a near normally incident vonex. The loading fluctuations can be 
calculated using unsteady thin-airfoil theory, and the far-field noise can then be calculated 
using Lowson's moving dipole theory (reference 13). Figure 2 shows the definition of the 
coordinates and the vortex orientation. According to Amiet, the far field pressure-time 
history can be calculated by 



2 -H>° 

inp^CzU j* i(k,Ut+^l(Mx-o) 

2aoa2 



i 71 p c z U (• i( 

p(t) = 2 k w(k ,k ) L(k ,k ,M) e 

^ ^ _2 J X ^ x' y-' ^ x' y' ' 



where po = the density of the acoustic medium, 

c = the tail rotor blade chord, 

ao = the speed of sound, 

M = U/ao, 

[i = kxM/(l-M2), 

a2 = x2+(l-M2)(y2+z2), 

L = the effective lift function, see reference 6 for details, and 
w = the Fourier decomposed vortex velocity field. 

In the present study, the effect of a moving vortex is included numerically, so U is 
the vonex normal velocity relative to the blade and can be obtained from the free wake 
geometry analysis (discuss later). 

To be consistent with the free wake geometry analysis, a different vortex model is 
used compared to Amiet's original analysis. In the present study, the tangential velocity for 
a concentrated 2-D vortex is defined by the widely used model: 

Ve = (r/27n-)T2/(r2+r^) 

where F is the vortex strength, and re is the vonex core radius. Figure 3 shows the 
comparison of the vortex model used in the present study and that used by Amiet. 

Since a different vonex model is used, w, the Fourier decomposed vortex velocity 
which is normal to the tail rotor plane, is found to be 



w 



= itane rr k K,( r ./k^+k^/cos^G ) / ((27u)\/ k^+k^/cos\ ) 



where Ki is the mcxiified Bessel function of the second kind, 0v is defined in Figure 2. It 
should be noted that w only accounts for the effect of the tangential velocity of the vonex; 
the axial flow in the main rotor tip vortices is neglected in the present analysis as discussed 
in the conclusion. 

The acoustic pressure-time history for a given blade-vortex interaction, calculated 
with the present vortex model, is compared to that obtained using Amiet's original 
formulation. The comparison is shown in Figure 4. The results for the two different 
vortex models show close similarities and only minor differences. More details on the 
acoustics analysis can be found in references 9 and 10. 



Illustrated Example and Discussion 

Main Rotor Tip Vortex Free Wake Geometry 

Since the main rotor tip vortex system is generally highly distorted (reference 11), 
classical rigid wake analysis (e.g. reference 14) can not predict the accurate locations and 
orientations of these vortices. The calculation of the free wake geometry of main rotor tip 
vortices is very important because the trajectories of vortices directly affect the 
characteristics of the interactions and subsequently the noise generated. In the present 
study, we use a comprehensive rotorcraft aerodynamics and dynamics analyses program 
(CAMRAD) of Johnson (reference 12) to calculate the main rotor tip vortex wake 
geometry. The CAMRAD code is based on a rotor-free wake geometry computation model 
of Scully (reference 11). Details on the CAMRAD code can be found in reference 12. 

In our application, the main rotor airloads are calculated assuming non-uniform 
inflow at the main rotor disk, but the presence of the tail rotor is assumed to have no effect 
on the main rotor tip vortex system. The wakes from other part of the airframe are also 
neglected in the present study. 

The aerodynamic analysis is aimed to provide the following information for the 
forthcoming acoustic analysis: the normal incident velocity of the ingested vortex relative to 
the tail rotor blade (U), the strength of the ingesting vonex (F), the skew angle between the 
vortex and the tail rotor axis (9v), and the vortex orientation with respect to the blade (<l)v). 
Refer to Figure 2 for the definitions of U, 0v, and (^y. It should be noticed that these 
values are generally not constant as a vortex convects through the tail rotor disk. 

In this report, we will use an example to illustrate the general procedures developed 
in the present study. Our simulation begins with the calculation of the main rotor free wake 
geometry using CAMRAD. Figure 5 show the main rotor tip vortex system at four 
different times, this case closely represents a UH-ID helicopter in 100 knots level flight. 
(We use a modified version of GEOMPl subroutine to generate the vonex location 
database.) From these figures, the interactions between the tail rotor and main rotor tip 
vortices are evident. 

Trajectories of the Vortices on the Tail Rotor Plane 

After the free wake geometry calculation is completed, we can find the vortex 
trajectory/trajectories on the tail rotor disk using program VTRAJ. VTRAJ requires the tip 
vortex location database and the tail rotor location to perform the interpolations. Figure 6 
shows four possible patterns of the vortex trajectories on the tail rotor disk. For the 100 
knots UH-ID case, only one vortex trajectory exists on the tail rotor disk and is classified 
to be a single vonex interaction. The vonex trajectory for this case is plotted in Figure 7, 



the points shown in the figure are interpolated from the free wake geometry analysis 
results, and they show the locations of the vortex at different times. 

Notice that the tip vortices involved in the interactions are originally shed from main 
rotor blades at approximately 0° azimuthal location. Figure 8 shows the definitions of 
azimuthal angle for both the main and the tail rotors. They are relatively "young" (less than 
180° for this case), which implies that the ingested vortices are not fully rolled- up 
(Johnson, reference 12, had suggested that a vortex is not fully rolled-up until the vortex is 
older than 180° or so). Since a vortex is not fully rolled-up, the strength of the ingesting 
vortex should be less than the maximum bound circulation on the main rotor blade; we 
follow the assumption made by Scully (reference 11), and discount the strength of this 
partially rolled-up tip vortex by 20% from the fully grown tip vortex. For the vortex core 
radius, re we used 0.0025 of the main rotor tip radius following the assumption of Scully 
(reference 11). 

The results shown so far also indicate that a vortex may be chopped by more than 
one blade. It is quite possible that the vortex will break down immediately after the first 
interaction, thus subsequent interactions with the same vortex will be much weaker 
compared to the first one. However, the free wake geometry calculations also shows that 
the main rotor tip vortex is generally drifting sideways through the tail rotor disk. 
Therefore the next blade is likely to interact with a fresh piece of tip vortex segment, thus 
no effect of possible interactions with burst vortices has been included in the present study. 

Detailed Definition of Tail Rotor Blade- Vortex Interactions 

The informations defining a series of blade-vortex interactions can then be 
determine numerically by interpolating the results from main rotor free wake geometry 
analysis (using VINTER). Since the tail rotor RPM is generally not an integer multiplier 
of the main rotor RPM, the location of the blade-vortex interaction is different for each 
main rotor revolution. We have to locate all possible interactions until the interaction 
pattern repeats itself (this process usually has to go through several main rotor revolutions 
and is different for every helicopter configuration) to capture the general features of this 
phenomenon. For the first case, both #1 blades of the main and the tail rotors are set such 
that both blades start from V|/ = 0° positions initially and the computations continue through 
approximately four main rotor blade revolutions. Refer to Figure 8 for the definitions of 
main rotor and tail azimuthal angles. 

Acoustic Signature Calculation 

For the 100 knot UH-ID case, the data defining a series of blade-vortex interactions 
(which we had obtained in the previous section) are now used as the inputs for the noise 

7 



calculation (program VNOISE). The results are plotted in Figure 9. The results are given 
in term of the pressure-time history, the pressure is in Pascal (N/m2), and the horizontal 
axis is the time axis (tick marks are 0.1 seconds apart). 

It should be notice that the pressure peaks are not separated by equal time intervals; 
therefore if one Fourier analyzed the pressure-time history, the resulting acoustic spectrum 
will behave more like broadband noise with widened spectrum peaks rather than pure 
harmonics. Also an interaction with higher normal velocity (advancing blades or near the 
blade tip) will result in a stronger but relatively short perturbation pressure peak; while an 
interaction characterized by smaller normal velocity (retreating blades or near the blade hub) 
will result in a weaker but more lasting pressure perturbation. 

In references 2 and 3, we have shown that these types of tail rotor blade-vonex 
interactions exist even at lower helicopter speeds. Although the overall vortex convection 
speed decreases with the lower helicopter airspeed, the interactions are more frequent. 
Also because the strength of the vonex has to increase at lower airspeed to generate the 
required lift, the magnitude of the acoustic pressure fluctuation does not seem to change 
much compared to the high speed case. Details of this study can be found in references 2 
and 3. 



Additonal Examples 

Effect of Tail Rotor Location 

As discussed previously, the vortex trajectory on the tail rotor disk is very 
important to the blade-vortex interaction noise. The tail rotor location relative to the main 
rotor, and the helicopter operating conditions are two primary variables that change the 
vortex trajectories on the tail rotor disk. To illustrate the effect of tail rotor location on the 
blade-vortex interaction noise, we will artificially lower the UH-ID tail rotor by 0.5 m. 

For die same 100 knots level flight case, the main rotor tip vortex trajectory on the 
tail rotor disk is now shown in Figure 10. Notice that the vortex path is higher than that 
shown in Figure 7 due to a lowered tail rotor and hence causing the blade-vortex 
interactions to occur with advancing tail rotor blades. Following the procedures described 
earlier, the interaction locations and vortex properties can again be determined using 
VINTER. The acoustic pressure signatures of the tail rotor blade-vortex interactions can 
also be calculated, the results are shown in Figure 11. 

There are considerable differences between the results shown in Figures 9 and 11. 
Since the vortex is passing through the advancing side of the tail rotor, this results a higher 
relative velocity between the tail rotor blade and the ingesting vortex element, so in general 
the pressure perturbation has higher peaks. Also the interactions are more frequent than in 
the previous cases. Unquestionably, with this configuration (lowered tail rotor), the 
overall tail rotor noise will be more significant than that from a standard tail rotor. 

Comparison with Experiments 

The most important helicopter tail rotor noise experiments we have examined are 
those of Balcerak (reference 15). In a parametric study, he found that the noise increased 
as the main rotor wake passed the tail rotor. His tests employed a model-scale (1/1 6th) 
UH-1 helicopter and input parameters are well-defined. Here, we will compare the present 
analysis to his experimental data. 

The case we chose to simulate corresponds to an advance ratio of 0.2. First, the 
main rotor free wake geometry was calculated using CAMRAD. The results are presented 
in Figure 12. Unlike from the full-scale UH-ID case, the model-scale calculations show 
multiple vortex trajectories on the tail rotor disk (Figure 13). The two vortex trajectories 
present on the tail rotor disk look similar but are quite different. The first trajectory (the 
upper one) was shed originally from approximately 0° main rotor blade location, the vortex 
segments on this path are relatively young, just Uke that in the full-scale UH-ID case. The 
second (the lower) trajectory, on the other hand, was shed from the 180° main rotor 
azimuthal position. The vortices on this path are much older than those on the upper 



passage (their ages are approximately 540* or older), and will be fully grown vonices 
(compared to those partially roUed-up vortices as on the upper passage). 

In most of Balcerak's tests, the model scale UH-1 tail rotor is rotating at an integer 
multiple of main rotor speed (main rotor 2120 RPM, tail rotor 10600 RPM for our case). 
The phasing of the tail rotor may have some impact on the overall noise generation; this will 
be discussed later. In the first example, we will assume zero phasing difference between 
the two rotors (both #1 blades start from 0° azimuthal locations). With this phasing, a set 
of blade-vortex interactions can be defined and the acoustic signatures calculated. Figure 
14 shows the acoustic results for this zero phasing case. 

Figure 15 shows the pressure signature and the corresponding noise spectrum for 
another zero phase difference case with a slightly lower tail rotor location, along with the 
experimental results of Balcerak (reference 15). The results compare fairly well with the 
experiment, considering that the experiment's tail rotor phasing was not known. 

Effect of Tail Rotor Phasing 

In this section we will examine the effect of different main/tail rotor phasings on the 
tail rotor blade-vortex interaction noise. Four more different tail rotor/main rotor phasings 
are examined for the case of Figure 14. In all cases, the #1 blade of the main rotor starts 
from the same azimuthal position (0°), but the tail rotor #1 blade stans from 30°, 60°, 90°, 
and 135° respectively. Since the main rotor free wake geometry stays the same as the zero 
phasing case, the first step (CAMRAD) and the second step (VTRAJ, locating the vortex 
trajectories) will be skipped. Only VENTER (to define blade-vonex interactions sequence) 
and VNOISE (to calculate the acoustic pressure signatures) are to be rerun. The acoustic 
results are presented in Figures 16-19. It is obvious that tail rotor/main rotor phasing can 
be very important; in the test cases presented, it not only affects the frequency of the 
interactions, it also affects the signal strengths of each interactions. 



10 



Conclusions 

Tail rotor blade-main rotor tip vortex interaction is a very important tail rotor noise 
mechanism. The noise generated depends strongly on the main rotor operating conditions 
and on the tail rotor location and its phasing. Major parameters governing this blade- vonex 
noise generation are the ingested vortex strength, the ingested vortex skew angle relative to 
the blade, and the relative velocity of the ingested vortex to the tail rotor blade. More 
detailed study should be devoted to the problem considering a vortex chopped by an airfoil 
of finite span. 

It is clear that the main and tail rotor relative phasing and location have a strong 
influence on the noise tail rotor generates. They offer opponunities to reduce tail rotor 
blade-vortex interaction noise by controlling the phasing and the relative location of the 
main and tail rotors. However, more work is needed to demonstrate the full potential of 
these techniques. 

The present study does not include the possibly major effect of the axial flow in the 
main rotor tip vortex. This can be another strong contributor to the unsteady loading 
fluctuation on a tail rotor blade. The result of free wake geometry analysis does indicate 
some evidence of the main rotor tip vortex drifting normal to the tail rotor disk. Also the 
strength of main rotor tip vortex is not constant, this will result in an axial pressure gradient 
inside the vortex, thus inducing some axial flow. These important problems should be 
addressed in future studies. 



11 



References 

1. George, A. R., Chou, S.-T.: A Comparative Study of Tail Rotor Noise 
Mechanisms. Journal of the American Helicopter Society, Vol. 31, No. 4, 1986, 
pp. 36-42. 

2 . Chou, S.-T., George A. R.: Progress in Helicopter Tail Rotor Noise Analysis. 
Proceedings of the 42nd Annual Forum of the American Helicopter Society, 
Washington, D.C., June 1986; also AIAA-86-19(X), AIAA 10th Aeroacoustics 
Conference, Seattle, Washington, July 1986. 

3 . Chou, S.-T.: A Study of Rotor Broadband Noise Mechanisms and Helicopter Tail 
Rotor Noise. Ph.D. thesis, Cornell University, Ithaca, New York, May 1987. 

4. George, A. R. and Chang, S. B.: Noise Due to Transonic Blade- Vortex Interactions. 
39th Annual National Forum of the American Helicopter Society, St. Louis, 
Missouri, May 1983. 

5. George, A. R. and Chang, S. B.: Flow Field and Acoustics of Two-Dimensional 
Transonic Blade- Vortex Interactions. AIAA-84-2309, AIAA/NASA 9th 
Aeroacoustics Conference, Williamsburg, Virginia, October 1984. 

6. George, A. R. and Lyrintzis, A. S.: Mid- and Far-Field Calculations of Blade- 
Vortex Interactions. AIAA-86-1954, AIAA 10th Aeroacoustics Conference, Seanle, 
Washington, July 1986. 

7 . Srinivasan, G. R., McCroskey, W. J., and Kutler, P.: Numerical Simulation of the 
Interaction of a Vortex with a Stationary Airfoil in Transonic Flow. AIAA-84-0254, 
AIAA 22nd Aerospace Sciences Meeting, Reno, Nevada, January 1984. 

8. McCroskey, W. J., Yu, Y. H., and Smetana, F. O.: Workshop on Blade- Vortex 
Interactions. NASA Ames Research Center, October 1984. 

9. Schlinker, R. H., Amiet, R. K.: Rotor- Vortex Interaction Noise. NASA CR-3744, 
October 1983. 

10. Amiet, R. K.: Airfoil Gust Response and the Sound Produced by Airfoil- Vortex 
Interaction. Journal of Sound and Vibration, Vol. 107, No. 3, June 1986. 

1 1 . Scully, M. P.: Computation of Helicopter Rotor Wake Geometry and its Influence 
on Rotor Harmonic Airloads. M.I.T. ASRL Report TR-178-1, Cambridge, 
Massachusetts, March 1975. 

12. Johnson, W.: A Comprehensive Analytical Model of Rotorcraft Aerodynamics and 
Dynamics, Part I, II, and III. NASA TM's 81182, 81 183, and 81 184, 1980. 

13. Lowson, M. V.: The Sound Field for Singularities in Motion. Proceedings of the 
Royal Society of London, Series A, Vol. 286, 1965. 

14. Heyson, H. H.: Nomographic Solution of the Momentum Equation for VTOL-STOL 
Aircraft. NASA TN-D-814, 1961. 

15. Balcerak, J. C: Parametric Study of the Noise Produced bv the Interaction of the 
Main Rotor Wake and Tail Rotor. NASA CR-145001, 1976. 

12 



Publications Issued During the Research 

1. George, A. R., Chou, S.-T.: A Comparative Study of Tail Rotor Noise 
Mechanisms. Journal of the American Helicopter Society, Vol. 31, No. 4, 1986, 
pp. 36-42. 

2. Chou, S.-T., George A. R.: Progress in Helicopter Tail Rotor Noise Analysis. 
Proceedings of the 42nd Annual Forum of the American Helicopter Society, 
Washington, D.C., June 1986. 

3. Chou, S.-T., George A. R.: Helicopter Tail Rotor Noise. AIAA-86-1900, AIAA 
10th Aeroacoustics Conference, Seattle, Washington, July 1986. 

4. Chou, S.-T.: A Study of Rotor Broadband Noise Mechanisms and Helicopter Tail 
Rotor Noise. Ph.D. thesis, Cornell University, Ithaca, New York, May 1987. 

5 . Chou, S.-T., George A. R.:Helicopter Tail Rotor Blade- Vonex Interaction Noise. 
Proceedings, NOISE-CON 1987, Penn State University, June, 1988. 

6. George, A. R., Chou, S.-T.: Helicopter Tail Rotor Blade-Vortex Interaction Noise. 
Final Technical Report, NASA Grant NAG 2-379, Period: February 1, 1986 - March 
31, 1987. 



13 



List of Figures 



Flight direction 



Strong main 
rotor blade 
vortex interaction 



\Neak 
main rotor 
blade vortex 
interaction 



Advancing 
side 



(a) 




Retreating 
side 




Vorticity vector normal to page 

Cross section 
of blade 



Vortex 




Tall rotor 
blade 



Figure 1. Geometry of Helicopter Blade- Vortex Interactions (from reference 9). 

(a) Illustration of the Relative Location of the Helicopter Main Rotor Tip Vortex 
System and the Tail Rotor. 

(b) Illustration of the Parallel Main Rotor Blade- Vortex Interaction. 

(c) Illustration of the Skewed Tail Tail Rotor Blade- Vortex Interaction. 



14 




Vortex 



Figure 2. Sketch of the 2-Dimensional Near-Normal Blade- Vonex Interaction. 



T 1 r 



-T 1 r 



Amiet 

Present 




Figure 3. Comparison of the Tangential Velocity Distribution for Different Vortex Models. 



15 



0) 

X. 

3 
V) 
CO -^ 

<U CM 

c 
o 

CO 



■20 



■15 



■10 



-5 



Amiet 

Present 



12 3 4 5 

Time (ms) 

Figure 4. Comparison of the Calculated Blade- Vortex Interaction Noise Signatures 
Using Different Vortex Models. 




< ::^ >c' >^' :>^ .^^ 



■***%»-_ y-_ ^ i r;;::Ln » ii^j^J 4, ill " * 




Figure 5. Calculated UH-ID Free Wake Geometry Results, 1(X) Knot Level Flight, 
(a) Main Rotor \(/ = 0°. 



fr k '-.i- 



Y 





''>-- -r^^^^"^^^::;^. 




A^^jiis*: 



Figure 5. Calculated UH-ID Free Wake Geometry Results, 100 Knot Level Flight, 
(b) Main Rotor Xj/ = 30°. 




~^ .-- '-~-^^-' 

^7-^* . ^"^ 




(c) Main Rotor \|/ = 60° 



17 



ORIGIi^AL ?MrZ B 
OF POOR Ql'ALH r 





"—T"^^^^ 



*^._^»K- 




Figure 5. Calculated UH-ID Free Wake Geometry Results, 100 Knot Level Flight, 
(d) Main Rotor \|^ = 90°. 



Vortex 




Vortex 1 



Vortex 1 ' 



Trajectory 



Vortex 1 



(c) Vortex 2 



k / Traj. 1 



Traj. 2 




Vortex 2 



Traj. 1 



Traj. 2 



Figure 6. Possible Main Rotor Tip Vortex Trajectory Patterns on the Tail Rotor Disk. 




Figure 7. Calculated Main Rotor Tip Vortex Trajectory on the Tail Rotor Disk, UH-ID 
100 Knot Level Flight. 



u 




Main rotor i z 



iy 



Tail rotor 




Figure 8. Definitions of the Main Rotor and Tail Rotor Coordinates. 



19 



E 
z 1 

a> 

3 

<fl " 

o 

i_ 

Q. 

3 
O 

w 
-2 



0.1 s 



Figure 9. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, UH-ID 
100 Knot Level Flight. 




Figure 10. Main Rotor Tip Vortex Trajectory on the Tail Rotor Disk, UH-ID 100 Knot 
Level Flight, Tail Rotor is Lowered by 0.5 Meters. 



20 



3r 



E 
z 1 

3 

V) n 
<n " 
a> 

a. 

3 
O 
(/) 



■3L 



0.1 s 



Figure 11. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, UH- 
ID 100 Knot Level Flight, Tail Rotor is Lowered by 0.5 Meters. 





V-,-»^ 



Figure 12. Calculated Model Scale UH-1 Free Wake Geometry Results, |i = 0.2 Level 
Flight, 
(a) Main Rotor \|/ = 0°. 



21 



OF FOG?J 



:»:■.'»;...,, ] !■ 





Figure 12. Calculated Model Scale UH-1 Free Wake Geometry Results, |i = 0.2 Level 
(b) Main Rotor \i/ = 30°. 





(c) Main Rotor \\f = 60° 



10 



ORIGIiVAL PmE IS 
OF POOR QUAUjy 




y 







'-^f-". 



•-.^ 



Figure 12. Calculated Model Scale UH-1 Free Wake Geometry Results, |i = 0.2 Level 
(d) Main Rotor \|/ = 90° 




Figure 13. Main Rotor Tip Vortex Trajectories on the Tail Rotor Disk, Model Scale UH-1, 
H = 0.2 Level Flight. 



23 



^ 3 



0) 

(0 
(0 



T3 

C 
3 
O 



■11- 



kl 



Tw 



20 



Time (ms) 



30 



Figure 14. Calculated Noise Signature for the Tail Rotor Blade- Vonex Interactions, Model 
Scale UH-1, |i = 0.2 Level Flight, Synchronized Main and Tail Rotors (Zero Phase 
Shift). 



E 



k_ 

3 
(0 
(0 

0) 

^. 

Q. 
"O 

c 

3 
O 
(0 




Time (ms) 

Figure 15. Comparison of Calculated Tail Rotor Blade- Vortex Interaction Spectrum and 
the Experiment Results of Balcerak (reference 15), Model Scale UH-1, |i = 0.2 Level 
Flight. Slightly Lowered Tail Rotor Position. 

(a) Calculated Signature. 



24 



SPL 

(dB) 




Frequency (kHz) 

Figure 15. Comparison of Calculated Tail Rotor Blade- Vortex Interaction Spectrum and 

the Experiment Results of Balcerak (reference 15), Model Scale UH-1, |i = 0.2 Level 
Flight. Slightly lowered Tail Rotor Position. 

(b) Calculated Spectrum and Experimental Results. 



r- 3 - 



(0 



c 

O 



■IL 



1 10 



20 r 



Time (ms) 



30 



Figure 16. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, Model 

Scale UH-1, ^ = 0.2 Level Flight, Synchronized Main and Tail Rotors (30° Phase 
Shift). 



25 





5 


- 








4 


. 




^ 
















z 
















0) 


3 


- 




^ 








3 








(A 








(/) 








0) 








k. 








Q. 


2 


- 




■o 








C 








3 








O 








0) 


1 







L 1 I 






'10 




-1 









L 



loT 



Time (ms) 



30 



Figure 17. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, Model 
Scale UH-1, ^i = 0.2 Level Flight, Synchronized Main and Tail Rotors (60° Phase 
Shift). 



CM 

E 
z 
a> 

3 

tn 
in 



■o 

c 

3 

o 
CO 



■1 i- 



T — r-^ 
f 10 



lof 



Time (ms) 



30 ms| 



Figure 18. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, Model 
Scale UH-1, |i = 0.2 Level Flight, Synchronized Main and Tail Rotors (90° Phase 
Shift). 



26 



<u 

^. 
3 
(A 

<n 

<D 

a_ 

Q. 
"O 

c 

3 
O 
0) 



-1 1- 



Tio 



20 [ 



Time (ms) 



30 



Figure 19. Calculated Noise Signature for the Tail Rotor Blade- Vortex Interactions, Model 
Scale UH-1, \i = 0.2 Level Flight, Synchronized Main and Tail Rotors (135' Phase 
Shift). 



27