4 

NASA Contractor Report -178172 

~~ 

[N&SA-CR-178172) DEVELOP ME NT OF BUTOfiCii APT N88-14762 

INTERIOR NOISE COUTROL CONCEPTS. PRASE 3 : 

DEVELOPIYENT OP NOISE CONTROL CONCEPTS Final 

Beport (united Technologies Corp,) 148 p Unclas 

CSCL 208 G3/71 01Q3572 

Development of Rotorcraft Interior 

Noise Control Concepts 

Phase : Development of 

Noise Control Concepts 

C.A. Yoerkie 

P.J. Gintoii 

S.T. ingraham 

United Technologies Corporation 

SIKORSK Y AIRCRAFT DIVISION 

STRATFORD, CT 06601 

J.A.Moore 

Cambrldge Collaborative, Inc. 

CAMBRIDGE, MA 02138 

. 

Contract NASI-I6932 

July 1987 

National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665-5225 
NASA Contractor Report -178172 

Development of Rotorcraft Interior 

Noise Control Concepts 

Phase IIT : Development of 

Noise Control Concepts 

C.A. Yoerkle 

P.J. Gintoli 

S.T. lngraham 

United Technologies Corporation 

SIKORSKY AIRCRAFT DIVISION 

STRATFORD, CT 06601 

J.A.Moore 

Cambridge Collaborative, Inc. 

CAMBRIDGE, MA 02138 

Contract NASI-16932 

July 1987 

NASA 

National Aeronautics and 

Space Administration 

Langley Research Center 

Harnpton, Virginia 23665-5225 
TABLE OF CONTENTS 

DESCRIPTION PAGE 

LIST OF FIGURES ....................................................... iv 

LIST OF PHOTOS ........................................................ viii 

LIST OF TABLES ........................................................ x 

FOREWORD .............................................................. 1 

INTERIOR CABIN NOISE CONTROL CONCEPTS ................................. 12 

Conventional Treatment approaches ................................ 14 

Advanced Treatment Approaches .................................... 18 

Introduction to Modeling Concepts ................................ 21 

Dissipative Type Treatments ...................................... 24 

Free extensional damping layers ............................. 24 

Constrained layer damping ................................... 27 

Acoustic absorption (cabin) ................................. 29 

Transmission Type Treatments ..................................... 35 

Trim panel acoustic cavities ................................ 35 

2-D acoustic modes .......................................... 36 

Effect of cavity absorption treatments ...................... 36 

Trim panel isolation mounts ................................. 41 

Interior Configurations of the S-76 Aircraft ..................... 43 

Utility interior ............................................ 56 

Executive (VIP) interior .................................... 57 

Advanced interior ........................................... 57 

SEA Model of the Treated Interior ................................ 57 

SEA models of trim panel transmission loss .................. 58 

Utility interior ............................................ 63 

Executive interior .......................................... 73 

Advanced interior ........................................... 77 

SEA Predictions for the Treated S-76 Cabin ....................... 84 

Power flow description of transmission paths ................ 84 

ISOLATION AS AN ALTERNATE NOISE CONTROL CONCEPT ....................... 92 

Gearbox Isolation Mounts ......................................... 92 

Preliminary Design Definition .................................... 95 

Physical constraints ........................................ 95 

Shape and size effects ...................................... 96 

Predicted isolation characteristics ......................... 98 

Isolation effectiveness ..................................... 110 

Plan for Validation .............................................. 114 

Isolation ................................................... 114 

ii 
8 

TABLE OF CONTENTS (Cont'd) 

DESCRIPTION PAGE 

CONCLUDING COMMENTS ................................................... 118 

RECOMMENDATIONS ....................................................... 119 

REFERENCES ............................................................ 120 

APPENDIX A ............................................................ 121 

STANDARD BIBLIOGRAPHIC PAGE ........................................... 138 

iii 
LIST OF FIGURES z 

FIGURE 

1 . 

2 . 

3 . 

4 . 

5 . 

6 . 

7 . 

I 8 . 

9 . 

10 . 

~ 

11 . 

12 . 

13 . 

14 . 

15 . 

16 . 

17 . 

18 . 

19 . 

Flow of Activities for the Development of Rotorcraft 

Interior Noise Control Concepts ............. 4 

Laboratory Experiment Schematic for Evaluation of 

Typical Constrained Layer Damping Cross Section ..... 15 

15 

Acoustic Treatment .................... 8 

4 

Typical Free Extensional Damping Layer Cross Section ... 

Typical Trim Panel. Cavity. Airframe Cross Section .... 16 

Typical Trim Panel Isolation Cross Section ........ 17 

Trim Panel Isolation Mount Nomenclature ......... 17 

Main Gearbox/Airframe Isolation Schematic ........ 18 

Main Gearbox/Airframe Isolation Mount Nomenclature .... 18 

Typical Composite Panel Sandwich Construction 

Crosssection ...................... 19 

Typical Constrained Layer Damping Composite Panel 

Cross Section ...................... 20 

Typical Double Honeycomb Constrained Layer Damping Composite Panel Cross Section .............. 20 

Loss Factor for Free Extensional Damping Treatment .... 26 

. 

Carpet Absorption .................... 32 

Carpet and Pad Absorption ................ 33 

Flow Resistivity of Glass Fiber Products ......... 37 

Compressibility of Air Within a Porous Material ..... 39 

Transmission Loss for a Utility Type ........... 60 

Interior Panel . 

Transmission Loss for an Executive Type Interior Panel ...................... 61 

iv 
LIST OF FIGU ES (Cont'd) 

FIGURE 

20a. 

20b. 

20c. 

21a. 

21b. 

21c. 

22. 

23. 

24. 

25. 

26. 

27. 

28. 

S-76 Overhead Panel Vibration 

Utility Type Interior .... 

S-76 Overhead Panel Vibration 

Executive Type Interior ... 

S-76 Overhead Panel Vibration 

Advanced Type Interior ... 

Levels for 

.............. 

Levels for 

.............. 

Levels for 

.............. 

S-76 Overhead Cavity Levels for Utility 

Type Interior ...................... 

Type Interior ...................... 

Type Interior ...................... 

S-76 Overhead Cavity Levels for Executive 

S-76 Overhead Cavity Levels for Advanced 

Utility Treatment Effectiveness Relative to Overhead 

Cavity Levels and Untreated Cabin Levels 

SEA Predictions for the Relative Distribution of 

Overhead Skin, Panel and Cavity Levels at 1000 Hz .... 

Coupling Loss Factors for Cavities to Cabin Through 

Panels and Holes (Leakage) for a Utility Type Interior. . 

Effect of Air Conditioning Openings and Panel 

Leakage on the SEA Predictions of a Utility 

Influence of Cavity Loss Factor on the SEA 

Predictions of a Utility Type Interior ......... 

Comparison of the SEA Predictions for a 

Utility Type Interior With S-76 Flight Data 

at the Standard Cruise Condition ............ 

Coupling Loss Factors for Cavities to Cabin 

Through Panels and Holes (Leakage) for an 

........ 

Type Interior ...................... 

Executive Type Interior ................. 

PAGE 

64 

65 

66 

67 

68 

69 

70 

71 

72 

74 

75 

76 

78 

V 
LIST OF FIGURES (Cont'd) 

FIGURE 

29 . 

30 . 

31 . 

32 . 

33 . 

34 . 

35 . 

36 . 

37 . 

38 . 

39 . 

40 . 

41 . 

42 . 

43 . 

44 . 

45 . 

Effect of Air Conditioning Openings and Panel 

Leakage on the SEA Predictions of an Executive 

Type Interior ...................... 79 

Comparison of the SEA Prediction. for an 

Executive Type Interior With S-76 Flight Data 

at the Standard Cruise Condition ............ 80 

S-76 Frame Transfer Function Levels ........... 85 

S-76 Skin Panel Transfer Function Levels ........ 86 

S-76 Overhead Cavity Noise Levels ............ 87 

S-76 Cabin Noise Levels ................. 89 

Typical Isolator Installation .............. 101 

Single Isolator Geometry ................ 102 

Gearbox Attachment Points and Sign Conventions ..... 103 

Airframe Attachment Points and Sign Conventions ..... 103 

Block Diagram of Analytic Model ............. 104 

Main Gearbox Driving Point Frequency Response . Port 

Forward . Vertical ................... 105 

Airframe Driving Point Frequency Response . Port 

Forward . Vertical ................... 106 

Main Gearbox Driving Point Frequency Response . Port Forward . Lateral ................. 107 

Airframe Driving Point Frequency Response . Port Forward . Lateral ................. 108 

Isolator Bench Test Results ............... 109 

Variation of Material Properties at Constant 

Frequencies ....................... 111 

vi 
LIST OF FIGURES (Cont'd) 

PAGE 

4 

FIGURE 

46. 

47. 

48. 

49. 

50 

51. 

52. 

53. 

54a. 

54b. 

54c. 

54d. 

55a. 

55b. 

55c. 

55d. 

Influence of Temperature on the Predicted Isolation 

Effectiveness with Two Isolators in the Vertical 

Direction ........................ 112 

Influence of Preload on the Predicted Transmissibility 

with Two Isolators at 8OoC in the Vertical Direction ... 113 

Predicted Isolation Effectiveness for Two vs. Four 

Isolators at 8OoC in the Vertical Direction ....... 115 

Predicted One-Third Octave Band Attenuation vs. 

Isolator Stiffness ................... 116 

Relative Comparison of Bare Cabin Levels to the 

Utility, Executive and Advanced Overhead Cavities 

and Cabin Levels ...................... 124 

Effect of Increased Leakage Area on Panel TL ....... 125 

Absorption Provided by Several Typical Cabin 

Typical Total Absorption of Cabin Furnishings ...... 126 

Furni shings ....................... 126 

Measured Intensity Contours for a Utility Type Panel 

at 0.5 kHz ........................ 127 

Measured Intensity Contours for a Utility Type Panel 

at1kHz.. ....................... 128 

Measured Intensity Contours for a Utility Type Panel 

at2kHz.. ....................... 129 

Measured Intensity Contours for a Utility Type Panel 

at4kHz.. ........................ 130 

Measured Intensity Contours for an Executive Type 

Panel at0.5 kHz ..................... 131 

Measured Intensity Contours for an Executive Type 

Panel at 1 kHz ...................... 132 

Measured Intensity Contours for an Executive Type 

Panel at 2 kHz ...................... 133 

Measured Intensity Contours for an Executive Type 

Panel at 4 kHz ...................... 134 

vi i 
LIST OF PHOTOS 

PHOTO 

1 

2 I 

9 

10 

11 

I 12 

15 

PAGE 

Sikorsky S-76A .............................................. 3 

Typical Utility Interior, Aft Passenger Cabin ............... 6 

Typical Executive Interior, Aft Passenger Cabin ............. 7 

Panel Testing - Reference Specimen, Reverberant 

Room Side .............................................. 9 

Panel Testing - Reference Specimen, Anechoic 

Room Side .............................................. 16 

Typical Utility Interior Seating ............................ 44 

Typical Utlity Interior Sidewall, Ceiling, and 

Aft Bench Seat ......................................... 45 

Typical Executive Interior with Soft VIP Seats, 

Forward Curtains, Hard Ceiling and 

Sidewalls .............................................. 46 

Executive Interior Aft Divan Section, Hard 

Surface Sidewall and Duct Covers ....................... 47 

Executive Interior with Table and Credenza .................. 48 

Executive Interior with Short Credenza. 

Curtains Provide Forward View for 

Passengers ............................................. 49 

Executive Interior with Soft Sidewalls and Duct 

Covers, Hard Center Celing ............................. 50 

Totally Soft Surface Executive Interior with 

High Back Seats and Extra Thick Padded 

Headrest .............................................. 51 

Totally Hard Surface Executive Interior, Aft 

Section ................................................ 52 

Totally Hard Surface Executive Interior with 

Full Forward Bulkhead Containing Sliding 

Windows ............................................... 53 

viii 
LIST OF PHOTOS (Cont'd) 

PHOTO 

16 

17 

18 

19 

20 

21 

22 

PAGE 

Executive Interior with Soft Surface Full Forward 

Bulkhead Containing Sliding Windows .................... 54 

Executive Aircraft Cockpit Treatment ........................ 55 

Isolation Test Rig .......................................... 99 

Typical Isolator ............................................ 100 

Typical Utility Type Panel Being Tested for TL in the 

Test Chambers from Anechoic Side. Sound Intensity 

Probe Shown ............................................ 135 

Typical View of Panel TL Testing . Anechoic Side ............ 136 

Typical View of Panel TL Testing . Reverberant Side ......... 137 

ix 
LIST OF TABLES 

TABLE 

1 

- 

2 

3 

4a 

4b 

5 

6 

PAGE 

Noise Control Treatments: Extensions to the SEA Model 

of the S-76 ............................................ 23 

Major Contributors to Power Flow into the Cabin ............. 73 

Major Contributors to Advanced Cabin Model at 1000 Hz ....... 81 

Summary of S-76 SEA Model: Bare Interior ................... 83 

Summary of S-76 SEA Model: Utility Interior ................ 83 

Power Flow Description for an Overhead Cavity 

(OC56L) ........................... ..................... 90 

Power Flow Description of the Utility Interior Cabin 

Noise Environment for LF14LO Vibratory Input ........... 91 

I h 
FOREWORD 

The goal of this NASA contract in broad terms is the understanding of helicopter internal noise mechanisms and the development, design, and testing of 

noise control concepts which will produce significant reductions in the acoustic environment to which the passengers and pilots are exposed. The Phase I 

and Phase I1 efforts [l, 2]* have produced much progress toward this goal. 

Phase I provided a basic understanding of the complex helicopter cabin noise 

environment and also produced an analytic method for prediction of the relatively high-frequency structure-borne vibrations which lead to the generation 

of bare cabin acoustic energy levels. The success of any large-scale analytic 

method clearly depends on validation of its accuracy. The Phase I1 effort 

provided the full-scale validation required to build confidence in the application of Statistical Energy Analysis (SEA) technology to helicopter cabin 

noise and structure-borne energy predictions in a bare aircraft configuration. 

The Phase I11 effort described herein involved the identification and evaluation of current and advanced treatment concepts, including isolation of structure-borne paths. In addition, a plan was devised for the full-scale test and 

evaluation of an isolation concept during Phase IV: Validation of Concepts. 

*Numbers in brackets refer to the Reference list. 

1 
PHASE I11 OBJECTIVES 

The Phase I11 development program general objectives are: 1) incorporation of 

various noise control treatment concepts into the analytic SEA model devel oped 

for the S-76A (see Photo 1) in Phase I [l] of this contracted effort and 

validated in Phase I1 [E], 2) prediction of the relative effectiveness of these 

noise control concepts, 3) definition and evaluation of a preliminary acoustic 

isolator design, and 4) formulation of a full-scale validation plan for the 

Phase IV effort. Specific items which will allow these objectives to be met 

are as follows: 

identification and characterization of various noise control 

concepts, 

implementation of noise control concepts within the S-76 SEA 

model, 

experimentation involving some aspects of these noise control 

concepts, 

definition and evaluation of a preliminary acoustic isolation 

design to reduce structure-borne transmission of acoustic 

frequency main gearbox gear clash vibrations into the airframe , 

formulation of a plan (to be implemented in Phase IV) for the 

full-scale validation of the isolation concept, and 

prediction of the cabin noise environment with various noise 

control concepts i nstal 1 ed. 

Completion of these objectives allows comparisons to be made showing the 

relative effectiveness of various noise control concepts on a dB per pound or 

dB per dollar basis. The Phase 111 effort and an overview of its place in the 

total program is shown in the Flow of Activities (Figure 1). Phase I11 provides a framework for the follow-on, full-scale evaluation of the most effective noise control concepts. Isolation of the airframe from acoustic frequency 

gear mesh loading tones and harmonics still appears to rank first in effectiveness, followed by various skin damping techniques, acoustic isolation of 

panels, and airframe modification. As such, the Phase IV activities shown in 

Figure 1 anticipate validation of an isolation concept as the primary effort, 

with airframe modification effects as an interesting approach for future aircraft. 

2 
ORIGINAL PAGE IS 

G!! EOOR QUAJJm 

Photo 1. Si korsky S-76A. 

3 
PHASE I 

LITERATURE 

OEFlNlllON 

NASA 

PHASE I1 

TRANSMISSION MOBILITIES. 

AIRFRAME MOBILITIES. 

OYNIMICS ACOUST!~ SOURCE 

ANALYSIS I TRANSMISSION 

STRUCTURE 

I I RESULTS 1 

I MOOEL MOBILITIES CONDUCT 

SUBSCALE 

ISOLATOR 1 ISOLATOR OYNAMIC CHARACTERISTICS ’ 

TESTS ISOLATOR ATTENUATION 

I 

I I 

I I I 

PHASE IV 

I PHASE Ill 

- SEPARATELY v~LIOATEO 

VALIOATE ’ ISOLATOR AN0 FUSELAGE ISOLATOR -ANALYSIS 0 

FULL SCALE 

I 

I 

I 

- 

.---1 

CONDUCT 

PARAMETRIC 

STUOIES USING 

PHASE I 

ANALYSIS 

TEST GROUNO 

ISOL IR 

AL E 

UN 

SELECTED NOISE PREOICTEO ISOLATOR CHARACTERIST~CS. 

CONTROL CONCEPTS 

6 REO’M’TS 

(XSMN ISOLATlONl 

DEFINE - ~~~~~~ OETAILEO 

ISOLATOR 

V:iONA- PRELIMINARY VALIDATION OESIGf4 

~ OESIGN “f” PlpN 

I-7--t- ----- 

AIRCRAFT 

FABRICATE 

ISOLATOR 

I 

I 

I 

I 

I 

- ---- 

AIRCRAFT 

FEASIBLE 

Figure 1. Flow of Activities for the Development of Rotorcraft 

Interior Noise Control Concepts. 

4 
These various tasks are described as: 

Task 1: Identification of noise control concepts 

In this task the practical considerations taken into account during the design, 

fabrication, and installation of an interior cabin treatment will be outlined. 

Included will be concerns for cost, weight, producibility, installation, 

accessibility, flammability and toxicity. These considerations place certain 

constraints on the treatment types. Current treatment approaches, including 

trim panels, panel isolators, skin and web damping, and absorption, will be 

identified. Advanced concepts which have attractive acoustic and weight 

benefits will also be identified. A mini case study will aid in these 

identifications by depicting real-world applications. 

Task 2: Characterization of the Task 1 noise control concepts 

The next step in the process is to characterize the acoustic and dynamic 

behavior associated with the various noise control concepts outlined in Task 1. 

These models incorporate any experimental/empirical information appropriate for 

the situation. Based on the relative confidence levels for these models, a 

small scale experiment was conducted to test the model validity. 

Task 3: Implementation of noise control concepts 

Once the noise control concepts have been identified and characterized, they 

must then be incorporated within the S-76 analytic model. This involves the 

transformation of information generated under Task 2 into a SEA format consistent with the S-76 SEA model [l]. In particular, the model was modified to 

include the standard interior shown in Photo 2 (light weight, hard surface trim 

panels), the executive (VIP) interior shown in Photo 3 (higher surface density, 

generally with soft surface trim panels), and an advanced noise control concepts interior variation. These additions transform the Phase I1 bare aircraft 

model into a model capable of predicting and evaluating the complete aircraft 

including the relative merits of acoustic treatment concepts. 

Task 4: Experimentation to support development of SEA noise control treatment 

model s 

During the identification and characterization of various noise control concepts, limited laboratory experiments are required to support this development 

work. Mounting characteristics for both trim panel isolation and gearbox 

isolation present two areas where experimental work is useful. These experiments involve shake test measurements to characterize the frequency response of 

the isolator material. In addition, experimentation on the sound blocking 

characteristics of the various trim panel types was performed. This was done 

in a special side by side chamber arrangement shown schematically in Figure 2 

and in Photos 4 and 5. Information gained in this task is used to enhance the 

S-76 SEA model developed in Task 3. 

Task 5: Prediction of the effect various noise control concepts have on S-76 

cabin noise levels 

5 
Photo 2. Typical Utility Interior, Aft Passenger Cabin. 

ORIGINAL PAGE TS 

POOR QZJ.4LITY 
Photo 3. Typ 

7 1 
I REVERBERANT 

TEST CHAMBER 

I 1-f I I 

I I i: 

I )’/‘ 

// I I // 

L’h’ ANECHOIC 

TEST CHAMBER / 

Figure 2. Laboratory Experiment Schematic for Eva1 uation of 

Acoustic Treatment. 

8 
Photo 4. Panel Testing - Reference Specimen, Reverberant 

Room Side. 

ORIGINAL PACE rs 

OF POOR QUALITY 

9 
10 

Photo 5. Panel Testing - Reference Specimen, Anechoic 

Room Side. 

ORIGINAL PAG3 IS 

OF. POOR QUALITX 
The SEA model, as modified to include these noise control concepts outlined in 

Tasks 1 through 4, is now ready to be exercised. This task includes the generation of additional parameters related to specific noise control treatments 

incorporated within the model by the Task 3 improvements. Parameters required 

for the basic bare aircraft were developed under Phase I1 and, where possible, 

are retained for this task. Predictions are generated for various noise 

control configurations. These predictions are compared with available measured 

aircraft data. The performance of these noise control concepts are then 

interpreted and evaluated relative to a benefit parameter such as cost and/or 

weight. 

Task 6: Definition of a preliminary isolation design and validation plan 

In this task the preliminary design definition of an isolator is accomplished. 

This involves the physical constraints present in the helicopter, such as: 

attachment point geometry and size limitations, static and dynamic shaft 

alignment (engine and tail drive shaft) for various flight maneuvers, control 

linkage motion and system sensitivity, airframe dynamic effects, main rotor 

loads transfer, and environmental integrity. 

The isolation desigdprediction procedure is exercised to determine candidate 

designs consistent with the above constraints and material characteristics. 

The shape and size of an isolator is then defined to achieve these desired 

characteristics. The final portion of this effort involves the formulation of 

a practical plan, which would form the basis of the Phase IV program, to 

establish, on a full-scale basis, the validity of this isolation design. The 

general intent is to ground test the isolator on a modified aircraft. 

11 
An interesting, albeit subtle, modeling change involves the acoustic radiation 

from the beam webs and flanges. In the bare cabin configuration these web and 

flange surfaces exhibit an unbaffled radiation into the cabin acoustic space. 

This is a very inefficient mode of radiation and, in the bare configuration, 

has negligible impact on cabin sound pressure levels. However, when trim 

panels are attached to the inboard side of the beam flanges, the web and flange 

radiation changes to a baffled configuration. Not only does the web and flange 

now become a more efficient radiator, but it also now radiates into the acoustic cavity created between the aircraft outer skin and the trim panel rather 

than into the cabin. When modeled in this manner, the influence of add-on beam 

damping will be seen not only as an energy loss mechanism for that particular 

subsystem, but also as a reduction in radiated energy contributing to the 

reverberent build-up in the acoustic cavity. 

INTERIOR CABIN NOISE CONTROL CONCEPTS 

Damping as a treatment type is being modeled in several distinct configurations, including: (1) constrained layer damping where a viscoelastic material 

is sandwiched between a base structure and a stiff non-structural member to 

provide a shearing action within the elastomer, (2) free extensional damping 

which is just a free damping layer applied to a surface, and (3) a damping 

layer within a built-up panel. The modeling of these three damping types 

involves changes to the mass and stiffness properties of the particular subsection as well as the energy propagation within that subsection. In some 

construction types, the coupling loss factors to other subsections may also be 

modified. 

Acoustic panel treatments are initially being subdivided by their conceptual 

impact on cabin acoustics. One type is the relatively thin trim panel, closely 

mounted to an adjoining panel (e.g.: an airframe outer skin panel). In this 

configuration there is a very small acoustic space created between the skin and 

treatment panel. Essentially the treatment can be considered as a septum which 

modifies the radiation efficiency of the skin panel while assuming that the 

trim panel response is of a non-resonant nature. The difficulty with this 

approach is that batting material stuffed in the bay between these two panels 

would not be properly modeled, since a change in radiation efficiency alone 

would not account for the energy dissipation mechanism of the intervening 

batting. The damping loss factor effects of batting material in larger bays is 

taken into consideration because the larger bays are considered to be acoustic 

spaces in the model (i.e.: an additional subsectioddegree of freedom). A 

larger space between the trim panel and the outer skin has several distinct 

differences: (1) the acoustic space will have resonant modes in the frequencies of interest, (2) the trim panel cannot be modeled as a radiation 

efficiency modification, and (3) the batting material can be treated as a 

modification to the acoustic space loss factor. 

12 
Another type of modeling consideration is the isolation of panels where the 

frames are coupled through point stiffnesses to panels which have their own 

resonances. Point impedance models are used at the attachment points looking 

into each subsystem. Consideration is given to the beam attachment and whether 

the energy transfer is from out-of-plane modes or the combined in-plane/ 

torsion modes. 

Included are typical conventional approaches, such as: 1) cabin acoustic 

absorption, 2) panel damping treatments , 3) trim panel and cavity absorption, 

and 4) trim panel isolation, as well as advanced treatment approaches, such as: 

1) gearbox i sol ation mounts , and 2) advanced composite materi a1 designs. 

13 
The following sections summarize these approaches and indicate areas of concern: 

Conventional Treatment Approaches 

1. Cabin Acoustic Absorption 

a) Use of absorbing "carpet" materials 

walls, floors, ceilings, furnishings (seating, credenzas, etc.), 

bul kheads 

b) Effect on total wall absorption in the cabin (absorption coefficient) 

c) Cabin acoustic loss factor for SEA acoustic subsystem 

d) Effect on pressure level distribution within the cabin: 

- direct field levels near radiating panels 

- reverberant levels in central region of cabin 

- non-uniform distribution of absorbing materials 

e) Information inputs from: 

carpet manufacturers, product suppl iers , reverberant room absorption 

testing, sample testing, impedance tube measurements. 

2. Panel Damping Treatments 

a) Constrained layer shear treatment 

14 
Con st rai ni ng Layer 

Viscoelastic Shear Layer 

Base Panet 

7 

L 

t 'T* 

31 -L 

7- 

-t 1 

Figure 3. Typical Constrained Layer 

Damping Cross Section 

b) Free extensional treatment 

Viscoelastic Layer 

\\\ -Neutral Axis of Composite 

t Base Panel 

Figure 4. Typical Free Extensional 

Damping Layer Cross Section 

c) Modeling to include: 

- increase in panel damping 

- effects on propagation behavior 

changes in bending rigidity of base panel and increase in total 

mass [3] 

d) Inputs to the model: 

- material properties of the viscoelastic layers 

mass, shear and Young's moduli, loss factor, frequency dependence of properties 

- geometries 

15 
3. Trim Panel and Cavity Absorption (Acoustic Path) 

Acoustic Cavity 

Skin Panel 

Trim Panel 

Absorptive Material (Fiberglass, 

Open Cell Foam) 

Figure 5. Typical Trim Panel, Cavity, 

Airframe Cross Section 

Acoustic transmission into the cabin 

Acoustic resonances in the cavity 

a) Absorbing cavity fill material 

- flow resistivity of blanket material 

- effect on sound propagation 

isothermal compression of air, 

porosity, flow resistivity, 

determination of propagation phase speed, 

attenuation [4] 

- damping of acoustic cavity resonances 

- typical cavity dimensions: 0.1 to 0.2 m 

b) Trim panel 

- mass law transmission 

slowly propagating bending waves - non-coincident 

coupling of acoustic cavity modes to 

acoustic radiation 

the cabin acoustic modes 

- resonant response of trim'panel 

coupling to acoustic cavity modes 

16 
d) Flanking transmission path due to gaps in trim panel 

junctions and penetrations (e. g. : 1 ights, vents, etc. ) 

4. Trim Panel Isolation (Structureborne Path) 

LTrim Panel Isolation Mount 

Frame Member 

Figure 6. Typical Trim Panel Isolation 

Cross Section 

Acoustic radiation due to point excitation of trim panel at mount 

support (radiation from flexural near field) 

a) Radiated power (P) due to point excitation 151 

b) Trim panel point impedance (Z ) 

c) Trim panel mount point velocity 

P 

Related to mount impedance characteristics and support 

frame velocity 

Isolator Mount Impedance Matrix 

Frame77P/- panel 

Figure 7. Trim Panel Isolation Mount Nomenclature 

17 
Required inputs to the model: 

d) Mount characteristics 

impedance measurements 

low frequency simple stiffness model 

SEA model of mount at higher frequency 

e) Final result is in the form of a coupling loss factor between 

the frame (or skin panel) subsystem, to which the mount 

attaches, and the cabin acoustic subsystem. Power is 

radiated by nearfield flexural vibrations of the trim panel. 

Advanced Treatment Approaches 

1. Gearbox Isolation Mounts 

a) Mount reduces power inflow to the airframe - (red-uced velocity at 

attachment point on the airframe) 

Mount 

.-\Airfame 

Figure 8. Main Gearbox/Airframe Isolation Schematic 

b) Impedance description at attachment point assuming uncorrelated 

motions/forces in the different directions. 

- Gearbox velocity in the i-th direction 

- Airframe velocity 

- Mount impedance matrix 

Figure 9. Main Gearbox/Airframe Isolation Mount Nomenclature 

18 
* 

Ratio of airframe velocities with mount, Va, to case with 

Change in input power into airframe: 

Modeling of attached components 

no mount, Va 

pGwer is proportional to: /Valz 

- Airframe: already in SEA Model 

SEA description of airframe impedance 

Measured mobility on actual airframe 

- Gearbox: not currently modeled 

Measured mobility on actual gearbox 

Further consideration required in deciding approach 

- Isolation mount: 

Low frequency simple stiffness model 

Higher frequency - SEA model of propagation resonances 

Measured impedances on actual mount 

Transmission line modeling of mount impedances 

within mount 

2. Advanced Composite Material Designs 

a) Composite panel sandwich construction 

Figure 10. Typical Composite Panel Sandwich 

Construction Cross Section 

19 
b) Composite panel constructions with constrained shear layers 

- Four layer construction 

Constrained Viscoelastic Shear 

Layer (Damping) 

Composite Honeycomb Core 

Composite Skin Panels 

Figure 11. Typical Constrained Layer Damping 

Composite Panel Cross Section 

- Five layer construction 

Composite Honeycomb Core Layer 

Constrained Shear Layer 

\ 

\Composite Skin Panels 

Figure 12. Typical Double Honeycomb Constrained Layer 

Damping Composite Panel Cross Section 

20 
Introduction to Modeling Concepts 

The previous section described conventional and advanced noise control treatments applicable for use in helicopter airframes to reduce cabin noise. This 

section addresses the development of models of their behavior for inclusion 

within the overall SEA model of the S-76 airframe and cabin noise environment. 

The cabin noise control treatments fall within two general categories based on 

their primary intended effect: 1) those effecting the dissipation of vibratory 

and acoustic energy; and 2) those that alter the transmission of energy from 

the source to receiver, in this case from the gearbox to the bounding panel and 

frame surfaces which radiate into the cabin. 

Dissipative treatments include add-on free extensional (Figure 4) and constrained layer (Figure 3) damping treatments on skin panels and frame webbing. 

Dissipative treatments within acoustic spaces include the surface absorption 

provided due to added fiberglass or open cell foam materials within the different cavities (Figure 5) other than the cabin that, in part, are created by 

the addition of trim and ceiling panels. 

Transmission type treatments have the primary effect of reflecting energy back 

towards the source, with the intent of reducing the energy which reaches the 

receiving subsystems. An important consideration is whether source levels 

increase as a result of the treatments, thereby negating their effects. It is 

often important to incorporate additional damping at the source in order to 

minimize the tendency for increased source levels. 

The primary transmission type treatments for the helicopter cabin are the side 

and ceiling trim panels that cover the structural skin panel and frame surfaces 

(Figure 5). Acoustically, their overall effect is to reduce the radiation into 

the cabin due to panel and frame vibration (Figure 6). They define additional 

acoustic spaces that are coupled to the skin and trim panels and which are 

coupled directly to the cabin through non-resonant mass controlled response of 

the trim panels. While not a treatment, per se, the presence of leaks or gaps 

in the trim panel coverage must be accounted for in describing the coupling 

across the trim panel between the cavity and the cabin. 

The isolation mounts, shown schematically in Figure 6, used to suspend the trim 

panels are transmission type treatments that affect the direct coupling of skin 

panel or frame vibratory energy into the trim panels. Isolation mounts at the 

gearbox attachment to the airframe (Figure 8) are a potentially important 

application of a transmission treatment directly affecting the dominant source 

of power flow into the airframe. Other types of treatment, including the 

dissipation treatments, affect individual connections or response levels often 

in parallel with many other paths having comparable contributions, so that 

significant reductions in cabin noise levels will require extensive application 

of such treatments. 

21 
Dissipative type treatments can a1 so affect the propagation behavior in the 

subsystems to which they are applied. Added panel damping treatments will 

effect the bending rigidity and surface mass of the panel. Fiberglass and foam 

treatments within the acoustic cavities created by the trim panels provide 

dissipation and in addition alter the propagation characteristics in the air 

within the cavity due to the porosity, thermal capacity and flow resistivity of 

the fi 11 material. 

Noise control treatments potentially effect all of the parameters of an SEA 

model of the cabin noise environment. Treated subsystems will have altered 

damping levels, mode densities and also coupling loss factors to other subsystems due to the effects of the treatments on propagation characteristics. 

In addition, modifications to the existing S-76 SEA model need to include the 

additional subsystems associated with particular treatments. Table 1 summarizes the specific items to be discussed in this section. 

22 
Table 1. Noise Control Treatments: Extensions to the SEA Model of the S-76 

1. Dissipative Type Treatments 

1.1 Free extensional damping 1 ayers 

1.2 Constrained layer damping 

1.3 Acoustic absorption 

2. Transmission Type Treatments 

2.1 Trim panels 

Resonant response 

Mass law response 

2.2 Trim panel acoustic cavities 

Resonant acoustic modes 

Cavity absorption 

Gaps or holes 

2.3 Trim panel isolation mounts 

2.4 Gearbox isolation mounts 

23 
Dissipative Type Treatments 

Free extensional damping layers. - The addition of viscoelastic layers to panel 

surfaces increases the overall damping level as a result of flexural and 

extensional deformation in the added -layers. The amount of damping increase is 

dependent on the relative energy storage within the viscoelastic material and 

base panel during bending. The greater the relative energy storage within the 

viscoelastic material, the greater will be the increase in overall loss factor. 

In addition to the material loss factor of the viscoelastic material, the 

stiffness properties and thicknesses of both layers are important in scaling 

the effectiveness of the damping treatment. 

The SEA model of a panel subsystem relies on a description of bending wave 

propagation characteristics in order to determine coupling loss factors between 

the panel and adjacent frames and panels, and also to adjacent acoustic spaces 

as a result of panel radiation. Bending propagation behavior a1 so determines 

the mode density of the panel. Panel damping loss factors are important, 

explicit parameters in the model. The added viscoelastic layer potentially 

effects the SEA model in each of these areas. 

The evaluation of bending rigidity of the composite, Figure 4, presumes that 

cross-sections remain plane during deformation. The neutral axis for bending 

is shifted as a result of flexural and extensional stiffness of the viscoelastic layer: 

- 1 Eb - Ev tt 

tn - z tb + tV 

where Eb, tb and Ev, tV are the elastic modulus and thickness of the base panel 

and viscoelastic treatment, respectively. The overall bending rigidity, Bc, 

is as follows: 

where tn is the distance between the viscoelastic 1 ayer/base panel interface 

and the neutral axis of the combined composite structure. The surface mass of 

the treatment must also be accounted for in determining the bending wave speed: 

24 
and 

BC 

Cb = fi C 

(4) 

where Mc is the surface mass per unit area and Cb the bending wave speed of the 

panel. 

The expression for the loss factor of the composite is taken from work by Ungar 

and Ross [3]. 

where 

BV 

rb = tb/12, rv = tv/12 

The deoendence of the comDosite 

loss factor of viscoelastic material 

ratio of extensional stiffnesses 

radii of gyration of base panel, 

viscoelastic layer 

distance between vertical planes of 

the layers 

loss factor on relative thicknesses and stiffnesses' is illustrated in' Figure 13 [4]. Increasing thickness or stiffness 

modulus of the viscoelastic material relative to the value for the base panel 

increases the energy storage within it resulting in a greater loss factor for 

the composite. 

25 
0.1 2 4 6 81.0 2 6 810 2 4 6 8100 

Relative Thickness of Damping Layer, H, - H,IH, 

Figure 13. Loss Factor for Free Extensional Damping Treatment. 

26 
The above description does not include the effects of temperature and frequency 

on the material properties of the viscoelastic material layer. If material 

loss factor and modulus are known for individual frequencies and temperatures, 

the above analytic expressions can be used to estimate propagation and damping 

behavior for the composite. 

Material property data can be obtained to varying degrees of completeness from 

the manufacturers and suppliers of viscoelastic material layers for damping 

purposes. Empirical damping loss factor data for specific treated panels may 

also be available from the manufacturers. Damping measurements on actual 

sample panels is an alternative source of loss factor data for input to the SEA 

subsystem model. 

Empirical data characterizing propagation behavior is relatively rare so that 

the above expressions would be needed to determine bending wave speed. The 

principle effect of the treatment is often due to the mass it adds to the base 

panel as stiffness moduli of typical viscoelastic materials are relatively 

small compared to those for base materials and do not significantly increase 

the bending rigidity. Also, the required composite damping level, while 

significantly greater than for the base panel material, is not typically on the 

order of that for the viscoelastic material but at least a factor of ten less. 

Treatments requiring larger composite loss factors often involve composite 

constructions with constraining layers or internal shear layers where the 

design focus is to induce more strongly damped shear deformation in the constrained shear layer, as will be discussed in the next section. 

Constrained 1 ayer damping. - Constraining 1 ayers are used i n conjunction with 

lossy viscoelastic material layers, as shown in Figure 3, to induce greater 

shear deformation, thereby taking advantage of the inherently larger damping of 

such materials in shear. The relative proportion of potential energy stored in 

shear deformation of the constrained layer is again important in determining 

the effective damping level of the composite. The degree of shear deformation 

that occurs is dependent on the curvature or wavelength of the panel bending 

de f ormat i on. 

At very long bending wavelengths little shear deformation occurs within the 

constrained layer resulting in reduced composite loss factor. The energy 

associated with shear deformation for very short wavelengths decreases relative 

to the energy in bending of the base panel, also resulting in a loss of effective damping. Constrained layer treatments exhibit a broadly defined optimum 

effective loss factor, as a function of frequency, related to the wavelength 

for bending wave propagation in the composite panel. 

The analysis of constrained layer damping treatments is also adapted from the 

work of Ungar and Ross [3] and which is described in Reference [4]. The 

composite panel loss factor is given by: 

27 
- ‘12 yx 

22 

- 

‘lc 1 + (2+Y)X + (l+Y) (1+P2)X 

where Y is a stiffness parameter given by: 

1 E t3 

3 

L[ Eltl+33 1 (- + -13 

Y 12 til E3t3 (7) 

and El, E3, and tl, t3 are the elastic moduli and thicknesses of the base panel 

and the constraining layer, and t31 is the distance between neutral planes of 

these individual layers. The quantity X is referred to as the shear parameter 

defined as: 

where G2 and t2 are the storage modulus in shear and thickness of the viscoelastic constrained layer. The wavenumber for bending wave propagation in the 

composite, k, is given by: 

where cb is the bending wavespeed: 

pc is the surface mass density of the composite: 

PC = Plt2 + $5 + P3t3 

28 
and B is its' bending rigidity, equal to the real part of the complex bending 

rigidity, B*, which is given by: 

* 

the complex shear parameter X* equals 

B1 and B3 are the bending rigidities of the base panel and constraining layer, 

respectively. 

The process of determining the effective loss factor requires an iterative 

approach to evaluating the shear parameter. The shear parameter depends on the 

bending wavenumber, k, which at a particular frequency, depends on the bending 

rigidity of the composite panel. Iteration is required because the bending 

rigidity depends on the shear parameter whose value we are seeking. A useful 

starting point for the iteration is to assume a long wavelength limit characterized by rigid shear coupling between the base panel and constraining 

layer. This yields an initial value for the bending rigidity of: 

B' = (B1+B3)(1+Y) (13) 

which is then used to determine a value for k and subsequently, X. Using this 

value of X determine B* and compare its real part with B. If it is close in 

value the iteration is complete, otherwise continue using the real part to 

determine a new value for k, etc. 

As was the case for free extensional damping treatments, the effects of a 

constrained layer treatment on both the damping level and bending wave propagation are required for the SEA model of the treated panel subsystem. 

Acoustic absorption (cabin). - This treatment refers to materials that are 

added to an acoustic space which result in the dissipation of acoustic energy 

within the space. The treatments include both materials placed against a 

bounding surface of the space, such as floor or wall carpeting and overhead 

panels, and also materials that behave as space absorbers. The latter refers 

primarily to porous seat upholstery which acoustically is characterized by its 

flow resistance. Porous materials within the cavities defined by trim panels 

are discussed in the section dealing with transmission treatments in that they 

occupy a larger volume fraction with potentially significant effect on acoustic 

propagation within the cavity. 

29 
Absorbing surface treatments are most straightforwardly accounted for using 

standard approaches from room acoustics. The loss factor for acoustic resonances within a space is related to the absorption of energy at the wall 

surfaces according to the following: 

cs - 

q=8nVfa! (14) 

where 

C - sound speed 

S - total wall surface area 

V - volume of space 

f - frequency, Hz 

and 6 is the area weighted average absorption coefficient for all bounding 

surfaces of the space defined by: 

-1 a=-I. s. a! Siii 

where Si, a! are the individual wall surface areas and corresponding absorption 

coefficients. 

i 

The absorption coefficients, a!., should account for the energy dissipated upon 

reflection from the surface aAd not include the transmitted energy which is 

accounted for in the SEA model in coupling to resonant panel response or direct 

coupling to adjac?nt spaces through non-resonant response of interviewing 

panels, or holes. When transmission of this type is not otherwise accounted 

for within the SEA model then it should be included in ai in addition to the 

dissipation at that surface. 

The most common source of absorption data is reverberation time measurements 

where the decay of acoustic pressure levels is recorded when the sound source 

is abruptly turned off. The reverberation time, TR, the extrapolated time required for the pressure levels to decay 60 dB, is related to the average absorption coefficient according to the Sabine equation: 

- 60V 

TR 1.086 cSG 

- 

30 
Reverberation time measurements are utilized to determine E which then gives q 

according to Equation (14). The loss factor for the treated cabin configuration may be directly estimated from reverberation time data, if available, 

according to the 

2.2 

q=- oTR 

If treated space 

absorption coeff 

fol 1 owi ng: 

data is unavailable then an independent 

cient of the treatment, aa, is required a 

(17) 

evaluation for the 

ong with Equations 

(14) and (15). Presuming also that information exists for Gut, the average 

absorption coefficient in the untreated case, then at in the treated case is 

based on aa and the treated area, St, according to the following: 

Absorption coefficient data is generally obtained from reverberant room measurements in specially designed test chambers. Representative data for two 

carpet materials are shown in Figures 14 and 15. The test materials are placed 

on an acoustically rigid floor in the chamber. It is presumed that effects due 

to potential differences in backing impedance between the reverberant room 

measurements and the actual cabin installation can be neglected in the above 

evaluation for it. 

While primarily intended for other functions, porous seating upholstery is also 

a source of acoustic dissipation in the cabin. Sound waves impinging on the 

relatively thin porous material layers will undergo dissipation within the 

pores of the material. Depending on the configuration of the seating both 

faces or sides of the material may be accessible to sound waves in the cabin. 

The modeling of a free standing porous absorbing layer accounts for the flow 

resistance, porosity and mass per unit area of the material. As sound propagates through the material , viscous dissipation occurs as a result of the 

motion of the air particles relative to the fibers making up the material. The 

viscous forces that are generated by the relative motion tend to accelerate the 

fibers along with the air particles. The fiber motion is inhibited by its mass 

reactance, presuming a limp structure. At lower frequencies the mass reactance 

may be insufficient to prohibit fiber motion resulting in a loss of relative 

motion between fiber and air, and an associated reduction in absorption provided. 

31 
1 .oo 

0.80 

0.60 

0.40 

0.20 

0 

FREQUENCY, Hz 

Figure 14. Carpet Absorption. 

32 
1 .oo 

0.80 

0.60 

0.40 

0.20 

0 

125 250 500 1000 2000 4000 

FREQUENCY, Hz 

Figure 15. Carpet and Pad Absorption. 

33 
The absorption coefficient is given by the following expression: 

IZa + ZmlL 

a 

L 

where: 

- flow resistance of material 

- 

Rm 

Mf surface mass per unit area 

ta is an acoustic radiation impedance including the air on both sides of the 

layer, and 

is the acoustic impedance of the material. It is the parallel combination of 

surface mass reactance and flow resistance. To characterize the absorption 

occurring in a diffuse or reverberant acoustic environment,an average over 0, 

the angle of incidence of the plane wave, is required to use the expression for 

CY. 

The loss of absorption at low frequencies or for very light layers is readily 

identified in the above expressions. When the surface mass reactance is small 

compared to the flow resistance, i.e., when 

<< 1 - I 

Rm 

34 
then 01 is also sma 

Zm approaches Rm in 

Rm 2- 

1. 

value, and: 

At higher frequencies or for heavier absorb 

At normal incidence, Za = 2pc and 

ng blankets 

which exhibits a maximum value for 01 equal to .5 when Rm =2pc. 

The effects of free standing absorbing layers on the acoustic loss factor of 

the cabin are accounted for according to the previously described approach 

based on the above expression for CY. Where sound waves in the cabin impinge on 

the material from both sides the appropriate area is twice the area of the 

material. Reverberation time measurements in the cabin with such materials in 

place provide a direct measure of their effect. 

Transmission Type Treatments 

Trim panel acoustic cavities. - The addition of trim panel treatments creates 

acoustic cavities between the trim and skin panels, which are segmented by 

frame members. Additional acoustic subsystems are straightforwardly included 

within the SEA model to describe the acoustic levels within these spaces. At 

lower frequencies the thickness dimension of the cavities becomes small compared to an acoustic wavelength and the resonances are those of flat two 

dimensional acoustic spaces. 

The acoustic cavities are coupled to the frame, outer skin, and trim panel 

structures that form the bounding surfaces. There is a direct coupling between 

adjacent cavities through holes in intervening frames and non-resonant mass 

controlled response of the frame webs. The cavities also couple directly into 

the cabin through non-resonant trim panel response and as a result of gaps in 

35 
the trim panel treatment. At higher frequencies and for the larger cavities in 

the cabin overhead, a transition occurs to three dimensional acoustic spaces 

for which all of the above coupling types are described in the current SEA 

model. 

2-0 acoustic modes. - At low frequencies, the two dimensional nature of the 

acoustic field in cavities created by trim panels effects the evaluation of the 

mode density. The mode density for a flat thin space, where S is the surface 

area and P is the perimeter around that area, becomes: 

P sw + - 

2nc; 2nc0 n (tu) = - 

This expression is valid when the thickness dimension is less than one-half the 

acoustic wavelength. 

Effect of cavity absorption treatments. - Acoustic treatments in the form of 

fiberglass or open cell foam materials are often added to control noise levels 

in cavities. In such cases the volume may be substantially filled with the 

absorptive material. Its effect is more than the absorption due to a simple 

surface treatment, and must be assessed in terms of its affect on acoustic 

propagation within the space. A1 though sound propagation in porous material s 

has been studied extensively by numerous investigators, a practical engineering 

description of the cabin noise environment for use within the SEA model is 

desired. The following approach is based on the discussion given in Reference 

[4] and accounts for the thermal capacity, porosity, and flow resistivity of 

the material in evaluating a complex propagation constant. 

The flow resistivity of common porous materials is dependent on the cross 

dimension of the filaments or fibers and the bulk density of the material, a 

description of the fiber content. This dependence is shown for common glassfiber materials in Figure 16. The relationship is described by the following 

algebraic expression: 

3 1.531d2 

f r = 3.18~10 pb 

where r is the flow resistivity in mks rayldm, pb is the density of the bulk 

material in kg/m3, and df is the fiber diameter in microns (lO-%n). The bulk 

density is related to the density of the material making up the fibers, pf, and 

the porosity of the bulk material, ho, according to: 

36 
1 2 3 4 5 6 7 0910 2 3 4 5 6 7 89100 

3 Density, kglm 

Figure 16. Flow Resistivity of Glass Fiber Products. 

37 
Thermal conduction between the air and the solid material can significantly 

influence the compression of the air as sound propagates through the material. 

At sufficiently low frequency the sound propagation occurs isothermally as heat 

transfer to and from the solid compensates for the temperature rise and fall 

which would otherwise occur during compression and rarefaction of the air. At 

high frequency the fluctuation occurs too rapidly for heat transfer to have a 

significant effect and the process is adiabatic. 

The gas compressibility relates pressure and volume changes according to the 

following expression: 

where K is the compressibility of the air and K' the compressibility of the 

porous material. K' accounts for the reduced volume of air available to be 

compressed as a result of the presence of the material. The volume, V, and 

associated volume change, dV, refer to the bulk porous material. The solid 

portion of the material is assumed to be volumetrically incompressible in comparison with the air. The variation of K with frequency due to thermal effects 

is shown in Figure 17. 

Flow resistivity directly affects the propagation by introducing dissipation 

due to viscous losses. Its effects are conveniently accounted for in terms of 

a complex density, p*, for propagation within the porous material: 

* i 1.2 r 

where p is the density of air and 

1.2 r 2 fl = 1 + (- 1 WJ 

and 

Pm 1.2 r 2 f2 = 1 + (ho + -1 (- P Pm" 

38 
/ I 1.3 Ox 1 O5 

1.2ox1o5 

1.1oX10 

7t 

I / I 1 

0.001 0.01 0.1 1.0 

f/Rl,m3kg-’ 

Figure 17. Compressibility of Air Within a Porous Material. 

39 
For materials with rigid frames: 

f =f =1 12 

The expressions, when applied to soft or semi-rigid materials, show an increase 

in effective air mass, i.e., the real part of the complex density, which is 

dependent on flow resistivity. The effective mass is that of the air alone for 

low flow resistivity. As flow resistivity increases the viscous forces accelerate the mass of the frame material along with the mass of the air particles 

resulting in an increase in effective air mass. 

-iwt A complex propagation constant or wavenumber for the porous material, for e 

time dependence, is defined by: 

where 

In terms of real and imaginary components: 

where cm, Am are sound phase speed and wavelength in the porous material, and d 

is the attenuation constant in nepers/m. 

The attenuation constant, d, can then be related to the loss factor for acoustic resonances within the porous material filled cavity by modifying the total 

sabine absorption to account for propagation loss in addition to wall absorption: 

Combining Equation (36) with Equation (14), replacing Sa in Equation (14) with 

(SCr) from Equation (36), gives the desired loss factor. t 
Alternatively, the component of loss factor associated with propagation losses 

can be obtained directly from the complex propagation constant by recognizing 

that the complex phase speed is related to the loss factor according to: 

and that 

by comparison with Equation (35), it can be seen that, 

*‘m 

‘7=, (39) 

The loss factor, given by Equation (39) , associated with propagation attenuation is added to the value for q corresponding to wall absorption to account 

for both dissipation mechanisms. This approach is preferred to that indicated 

by Equation (36) in that it does not depend on details as to whether the cavity 

is a two dimensional or three dimensional acoustic space. Equation (36) is 

based on a three dimensional diffuse field representation of the acoustic 

space. 

Trim panel isolation mounts. - Trim panel mounts are important to the SEA model 

in that they are a potential flanking transmission path to the airborne transmission loss provided by the trim panels themselves. They constitute a flanking transmission path in two respects, firstly power flow through the mount 

transmits into resonant motion of the trim panels with subsequent radiation 

into the cabin. Secondly, the point attachment of the mount to the trim panel 

generates a flexural near field that radiates directly into the cabin. This 

section describes the modeling of the mount dynamic behavior and calculation of 

coupling loss factors for both mechanisms. 

Commonly, mounts attach to frame members making up the cabin airframe structure 

as shown schematically in Figure 6. The desired coupling loss factors are between the frame and trim panel subsystems in one case and directly between the 

frame and cabin acoustic subsystems for the mechanism involving radiation from 

the flexural nearfield. Coupling loss factors are evaluated for out-of-plane 

motion of the frame and compressional motion through the mount. 

The acoustic power radiated by the flexural nearfield of a point excited plate 

is given by the following [4]: 

41 
Where: Ac - bending wavelength of trim panel at coincidence 

frequency 

N 

V$ - mean square velocity of the trim panel at the point 

where it attaches to the frame through the isolation 

mount. 

The trim panel velocity is related to the mean square 

the mount attachment, V2, by the following expression which depends on impedances characterizing th8 dynamic behavior of the mount and the trim panel: 

velocity of the frame at 

where Z is the point impedance of the trim panel: P 

zP = % 

where D M are the flexural rigidity and surface mass density of the trim 

panel and Z12, 

P’ P 

ZZ2 are transfer and input impedances for the mount. 

FO zll -z12 vO 

F1 z12 -z22 v1 

- 

When the mount is modeled as a simple spring stiffness, these impedances 

become: 

(43) 

The mount behaves as a simple spring stiffness until high frequencies relative 

to its dimensions where wave propagation and resonance effects occur within the 

mount. 

42 
The coupling loss factor between frame out-of-plane motion and the cabin 

acoustic space is obtained by combining Equations (40) and (41) with the 

additional relation that: 

Ef = Mf g2 0 (45 1 

where Mf is the total mass of the frame subsystem, to then obtain: 

- %ad - wqf ,aEf 

1 212 1 

‘If,a w n3 1Zp+Z22 l2 - Mf - --- (47) 

The coupling loss factor between frame motion and resonant trim panel motion 

describes the power input to the trim panel as a result of frame motion. 

The trim panel point impedance is purely resistive (i.e., Z = R ) so that: 

‘If,t’ 

P P 

=1R I 

‘If,t w p R +Z 

P 12 

(49) 

For mount attachments to skin panels the coupling loss factor expressions in 

Equations (47) and (49) are used where the frame subsystem mass is replaced by 

the skin panel mass. For mount attachments to both frames and panels the SEA 

mean square response velocity is presumed to adequately represent the velocity 

at the mount attachment location. The isolation mount is presumed not to 

significantly load the airframe structure to which it is attached. 

Interior Configurations of the S-76 Aircraft 

General: As discussed earlier, typical conventional approaches for noise 

control that we intend to focus on are: 1) cabin acoustic absorption, 2) panel 

damping treatments, 3) trim panel and cavity absorption, and 4) trim panel 

isolation mounts. These concepts are incorporated into the S-76 aircraft 

interior design to the greatest degree by the construction and installation of 

43 
Photo 6. Typical Utility Interior Seating. 

44 ORIGINAL PAGE IS 

OE POOR QUALITY 
I 

Pho-o 7. Typical It1 i ty Interior Sidewa 

Aft Bench Seat. 

1, Cei ing, and 

45 
46 

Photo 8. Typical Executive Interior with Soft VIP Seats, 

Forward Curtains, Hard Ceiling and Sidewalls. 

ORIGTNAL PAGE IS 

OE F!OOR QUALITY 
Photo 9. Executive Interior Aft Divan Section, Hard 

Surface Sidewall and Duct Covers. 

47 

I 
8P 
i 

I 

Photo 11. Executive Interior with Short Credenza. 

Curtains Provide Forward View for 

Passengers. 

a 

49 
Photo 12 

50 

Executive Interior with Soft Sidewalls and Duct 

Covers, Hard Center Celing. 
Photo 13. 

.ORIGINAL PAGE IS 

DE POOR QUALITY 

Totally Soft Surface Executive Interior with 

High Back Seats and Extra Thick Padded 

Headrest. 

51 
Photo 14. Totally Hard Surface Executive Interior, Aft 

Secti on. 

52 
Photo 15. Totally Hard Surface Executive Interior with 

Full Forward Bulkhead Containing Sliding 

Wi ndows. 

53 
Photo 16. Executive Interior with Soft Surface Full Forward 

Bulkhead Containing Sliding Windows. 

54 
Photo 17. Executive Aircraft Cockp t Treatment. 

55 
the trim panel system. The panel is basically an isolated mass barrier with 

absorptive materials added. The absorptive material is added to both sides of 

the panel to face into both the cabin and outer cavities. Additional absorption in the cabin comes from material selection of the seats and carpeting on 

the deck. 

The basic construction of an Interior system as modeled in this report consists 

of three overhead panels and four side panels (two on each side). The panels 

and seats are all upholstered with material best suited for their mission, i.e. 

durability and ease of maintenance in the Utility type and visually aesthetic 

plus highly noise absorptive for the quieter Executive type. 

The overhead panels are inserted into H-channels (designated as Edge Panels in 

the model) which are supported from the overhead frames again through isolators. The H-channels also support on their outboard side the overhead air 

conditioning ducts. Overhead lighting and further ventilation is built into 

these panels with loaded vinyl boots behind them to prevent acoustic leakage. 

The air conditioning ducts supply air from the Environmental Control Unit (ECU) 

located in the luggage compartment forward to the cabin and cockpit. The 

ducting splits after going through the rear bulkhead and runs up across the 

overhead outboard of the new overhead trim panels. Venting of the air takes 

place along the length of the overhead duct and is then brought into the 

cockpit overhead and down the sides behind the forward side trim panels. 

Return air is drawn from the junction area of the cockpit and cabin. The ducts 

are supported by the H-channels on one side and by isolators attached to the 

overhead frames on the other. The rest of the interior layout consists of a 

wool rug covering all deck space and upholstered bench seats for a seating 

capacity of up to 15 persons. 

The side panels are attached directly to the airframe through isolators (four 

for the rear panels, six for the forward panels). Built into the side panel is 

an interior window which is isolated from the outer window. This creates a 

double pane effect for further noise attenuation. 

Utility interior. The S-76 Utility aircraft is intended to be a low cost 

aircraft for general transportation such as shuttling of crews to and from 

off-shore oil drilling platforms. The interiors are designed for quick and 

efficient removal and maximum payload capacity (see Photos 2, 6 and 7, typical). Therefore, the panels and treatments are of basic construction and 

lightweight. Accordingly, attenuation of noise levels is nominal, roughly 10 

to 15 db down from bare aircraft levels. 

The structural member of the interior panels used both for the overhead and 

down the sides is made of 1/4" ABS plastic with a cover of 1/4" foam and vinyl 

trim cloth. The outboard side of the panels have an acoustic treatment of 1/2" 

fibrous glass batti ng (see Figure 5). 

56 
Executive (VIP) interior. The Executive Interior is designed for comfort of 

the passengers. Person capacity is limited to generally six passengers. Seating and interior layout is plush with many amenities for passenger comfort. 

Low noise levels in this type of aircraft is of mime imDortance and accordingly much attention to detail to prevent 

levels in an Executive Interior equipped c 

down from bare aircraft levels. 

The interior trim panels are laid out as 

interior layout. The trim cloth covering is 

as visually by using fabric material (genera 

characteristics. Transmission loss of the 

using a 1 psf loaded vinyl. Leakage between 

acoustic leakage is made. Noise 

ircraft are generally 25 to 30 dB 

described above for the general 

improved both acoustically as well 

ly wool) which improves absorption 

acoustic treatment is improved by 

panels is prevented by overlapping 

the loaded vinyl across panel junctions. The acoustic spaces behind all- trim 

panels is filled with a fiberglass batting to improve absorption in those 

cavities. Seats are made of the same trim fabric with thick foam for comfort 

and increased cabin absorption. 

Advanced Interior. The intent of the Advanced Interior concept was to take the 

lessons learned from the model about the Utility and Executive Interiors and 

apply treatment improvements to the model that could be again compared with 

equivalent treatments on a physical aircraft. Hopefully, the treatments in the 

aircraft would show positive results similar to the model prediction as a 

demonstration of the diagnostic capabilities of this method and as further 

varification of the model. 

The improvements attempted in the model incorporated: 1) increased damping due 

to a constrained layer damping package applied on eight frames and four skin 

panels located near the two forward foot attachments; 2) increased absorption 

in available exterior cavities using fibrous glass material, and 3) increased 

absorption on outer material of the trim panels due to improved absorption 

capability of materials used. 

Sea Model of the Treated Interior 

The SEA model of the bare interior consists of 130 SEA subsystems connected to 

each other at 245 junctions [Ref. 1,2]. The subsystems include frame sections, 

skin panels, and acoustic spaces (principally the cabin). The subsystems for 

the treated interior i ncl ude those for the bare aircraft plus additional panel 

subsystems that make up the trim panel treatments of the cabin overhead and 

sidewalls, as well as additional acoustical spaces for the cavities formed 

between the trim panels and outer skin panels. The trim panels are supported 

by soft rubber isolators which are modeled by a stiffness connection between 

the airframe structure and trim panel. Resonant transmission through the mount 

is also included, adding a subsystem for each isolation mount. Panel and space 

subsystems are added to describe the air conditioning ducting and duct cavities. The subsystem total increases to 307 for the treated interior with a 

total of 945 junctions. 

57 
Verification of the model predictions on an overall basis is done by comparison 

of the cabin subsystem predictions to actual interior noise levels from flight 

aircraft. This has been done in Reference 2 for bare interiors. Comparisons 

within this report are done relative to those bare levels - actual and predicted. 

The SEA program takes individual subsystems in the aircraft and models them 

using standard subsystem types, such as cavities, panels, and frames with 

specific junction types defined at a subsystem end or middle. This report 

concerns itself with how close this comes to real world response by comparison 

with the laboratory experimentation and measured aircraft data. 

The predicted response of frames and skin panels as part of the airframe has 

been shown to compare favorably to measured data for the bare aircraft configuration [Ref. 23 and is the basis of this model. The predicted flow of 

energy through the interior system as modeled by SEA techniques and how those 

predictions compare with the response of interior subsystems as measured 

through experimentation is discussed here within. Modeling of the treatment 

panels had not been done before and had to be compared to actual response. 

This was accomplished by measuring panel TL in the laboratory and then comparing with predicted SEA models. The panel was created using the SEA model 

for a panel and incorporating the parameters of the interior panel. These will 

be shown to compare favorably with actual panel responses. From this the 

response of the various panels could be modeled into the aircraft. This helps 

with the prediction of the Advanced Interior model which incorporated panels 

that have been laboratory tested, but not yet flight tested. 

Laboratory experiments were performed generally within two test chambers 

designed to allow a variety of different acoustic measurements. The chambers 

are positioned side by side, yet isolated dynamically from each other and from 

surrounding structure. One room is a reverberant type chamber and the other is 

an anechoic type chamber. The rooms can be used individually or together by 

opening the test section between the rooms. 

Other subsystem verification was performed either directly on the various 

aircraft, such as cabin absorption, or calculated, such as cavity absorption. 

Sea models of trim panel transmission loss - In the SEA cabin noise model the 

trim panels are separate structural subsystems that couple directly to adjacent 

acoustic space. In addition, non-resonant panel response results in a direct 

coupling between adjacent spaces on either side of the trim panel, independently of the coupling to resonant panel response. An area junction between a 

trim panel subsystem and adjacent spaces consists of three separate coupling 

loss fsctor evaluations: resonant panel response separately coupled to each 

space, and space to space coupling through non-resonant panel response. 

58 
To substantiate the SEA modeling of trim panel behavior a series of laboratory 

experiments were carried out to evaluate the resonant and non-resonant transmission through trim panel designs for the Utility, VIP, and Advanced interiors. Leakage around the edges of trim panels also constitutes an additional direct coupling between the adjacent acoustic spaces. A series of 

experiments were also performed to support the modeling approach for leakage. 

The experimental procedures are described in Appendix A. The measurement 

involves placing the test trim panel in an opening between a reverberent room 

and an anechoic chamber. With an acoustic source in the reverberent room the 

power transmitted through the panel to the anechoic chamber is measured using 

an acoustic intensity probe scanning over the panel surface to determine the 

total power. 

An SEA model was developed of the experimental test setup. The room volumes 

and the trim panel size were identical to those for the measurement. The 

output of the SEA model was interpreted in terms of the transmission loss (TL) 

in the same manner as the data reduction that was applied to the experimental 

results to determine TL. The acoustic intensity incident on the test panel was 

determined from the space average pressure level in the reverberent room. In 

identical fashion the SEA model pressure response in the source space was 

converted to an incident acoustic intensity. The intensity probe was used to 

directly measure the transmitted intensity, and the SEA model directly predicts 

the transmitted power. The transmission loss is the ratio of transmitted to 

incident intensities. 

The Utility trim panel design is basically a limp mass construction in that 

panel bending and coincidence effects do not occur until high frequencies above 

the band of primary interest. The measured TL results in Figure 18 following 

the conventional mass law behavior with a 6dB change in TL for each doubling or 

halving of frequency. The SEA result follows closely the measured data. 

As previously described, the VIP trim panel is fundamentally a double wall 

construction with two mass layers separated by a compliant foam core. The 

outer foam and facing layers can be expected to have little influence on the 

TL. Their function is to provide absorption in the adjacent acoustic spaces. 

Such constructions can be modeled in detail with the addition of SEA subsystems 

to describe each mass layer, which must also include flexural stiffness effects, and the foam core. 

The primary TL feature of double wall constructions is a dip in TL in the 

frequency region near the double wall resonance of the masses of the mass 

layers against the stiffness of the foam core. Above the double wall resonance 

the TL increases rapidly at a rate approaching a practical value of 12dB per 

doubling of frequency. The TL exceeds mass law based on the total mass of the 

panel design. This behavior is observed in the experimental results in Figure ia 

59 
50 

40 

TRANSMISSION 

LOSS, 

dB 

250 500 1000 2000 4000 8 

FREQUENCY, HZ 

0- 

X MEASURED 

SEA PREDICTION 

-MASS LAW 

i 

100 

Figure 18.Transmission Loss for a Utility Type 

Interior Panel. 

60 
7( 

6C 

5c 

TRANSMISSION 

LOSS, 

dB 

4c 

2c 

0 

io 500 1000 21 

Y 

I 

FREQUENCY, Hz 

X MEASURED 

SEA PREDICTION 

- MASS LAW 

I 

0 4000 81 0 

Figure 19. Transmission Loss for an Executive Type 

Interior Panel. 

61 
A simplified SEA model of the VIP panel was developed in order to avoid the 

additional complexity in modifying the cabin noise model to include the increased number of subsystems required to describe the double wall constructions. The double wall dip in TL is described by introducing panel coincidence 

for conventional panel bending at the frequency of the double wall resonance. 

The rapid increase above double wall resonance is achieved by adding a simple 

distributed stiffness at the area junction between the panel and the adjacent 

spaces. In qualitative terms the added stiffness accounts for the stiffness of 

the foam core. Increased panel damping levels also contribute to modeling the 

rapid increase. No additional SEA subsystems are required to describe the VIP 

panel when modeled in this fashion. The comparison of the SEA model results 

and the measured data is good. 

A final comparison involved the effects of leakage as a flanking transmission 

path between adjacent spaces. For the experimental measurements holes, in the 

form of narrow slits, were cut in the utility trim panel construction. The 

results in Figure 55 show the adverse effects of leakage in reducing the TL. 

An otherwise negligibly small leakage area that is .OS% of total panel area 

results in an up to lOdB decrease in TL at higher frequencies. 

The SEA model of the experimental setup was augmented to include the transmission through the leakage by adding an area junction equal in area to that of 

the opening in the panel. In the cabin noise model the leakage area can only 

be estimated. The SEAM code estimates impedances characterizing the radiation 

at the opening into the adjacent spaces. The predicted results agree with the 

measured data at lower frequencies, particularly for the smaller opening. At 

lower frequency the panel TL is not as great so that transmission through the 

panel dominates and additional transmission through the opening, which is less 

frequency sensitive, is not as significant. At higher frequencies the panel TL 

is greater so that transmission through openings will be of greater consequence. 

The SEA model at higher frequencies overpredicts the total transmission, with 

lesser TL values in comparison with the data. The predictions level out to 

constant TL consistent with a ray acoustics characterization of the opening 

where the total transmission at high frequency becomes proportional to the area 

ratio of the opening to the total panel area. 

Sources for the discrepancy may lie with the modeling in not appropriately 

accounting for geometry effects in the shape of the opening, or possibly, in 

accounting for flow resistance as the air in the opening oscillates back and 

forth due to the incident acoustic wave. 

62 
Utility interior - Sea predictions of cabin and cavity levels are shown in 

Figure 22 along with the cabin levels for the bare aircraft. The same arbitrary input power (to the frame section that the gearbox attaches to on the 

left side) was used for the bare and treated cases without attempting to scale 

the results for actual flight input levels. An 11 to 18 dB reduction in cabin 

levels from 500 to 4000 Hz is predicted. This is consistent with differences 

in flight levels for the Bare and Utility interior aircraft. 

As described above, the interior treatment is basically add-on in character and 

would not be expected to significantly change the vibration transmission in the 

airframe structure. SEA predictions of frame and skin panel levels confirm 

this, showing negligible changes (<l de) with the addition of the treatment. 

However, the trim panels trap acoustic energy radiated by the skin panels and 

framing, thereby raising the levels in the cavities in comparison with the bare 

airframe cabin levels. Added acoustic absorption in the cavities is insufficient to significantly limit this buildup of acoustic energy. The overhead 

cavities are strongly coupled by openings in the frame webbing that allow for 

the passage of wiring, control cables, etc. As a result the cavity levels are 

more uniformly distributed throughout the cabin overhead in comparison with the 

skin panel vibration levels which are the primary source of radiation into the 

cavities, as shown in Figure 23. 

Figure 20a and 21a show the predicted energy of the various cavities and 

associated skin panels in the overhead region. As expected levels are highest 

around the input source and generally decrease as the subsystems get further 

removed. OP14D is misleading as it is a very "thin" cavity (only 1/2"). 

The increase in acoustic levels in the cavities with the trim panels in place 

diminishes the panel effectiveness in reducing cabin noise levels. The noise 

reduction potential of a panel system is dependent on the panel's transmission 

loss (TL). An independent measurement of trim panel TL was carried out in a 

reverberant room/anechoic chamber facility. Power transmitted from the reverberent room through the panel was measured using an acoustic intensity probe on 

the anechoic chamber side of the panel. 

The data is compared in Figure 18 with SEA predictions for the experimental 

setup of a 4.1 kg/m2 panel. The behavior is governed by mass law TL for the 

frequencies of interest. 

In the aircraft leakage around the trim panels can significantly reduce the 

noise reduction actually achieved if not carefully limited by proper design and 

installation. Estimates of leakage area were incorporated in the SEA coupling 

between the cavities and the cabin. Coupling loss factors between cavity and 

cabin characterizing panel TL and transmission through leakage openings are 

shown in Figure 24. At 500 Hz transmission through the trim panel is greater 

than through the leakage openings. The relative contributions are nearly equal 

near 1000 Hz. Above 1000 Hz leakage contribution dominate due to the increase 

in panel TL associated with mass law behavior. 

63 
FORWARD 

LOOKING DOWN 

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

. . . . . . . . . . . . . . . . . . e... ..... . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . dB 

Figure 20a. S-76 Overhead Panel Vibration Levels for 

Utility Type Interior. 

64 
LOOK 

I////////// 

’////////// 

........... 

........... ........... 

.......... ........... .......... ........... .......... ........... .......... 

.......... 

.......... .......... 

.......... .......... .......... .......... .......... .......... .......... .......... .......... ......... .......... ......... .......... ......... .......... 

::OCA 5 R:: ......... ......... ......... ......... .......... ......... .......... ......... .......... ......... .......... ......... ......... ......... ......... ......... a ......... ......... 

Figure 20b. S-76 Overhead Panel Vibration Levels for 

Executive Type Interior. 

65 
FORWARD 

LOOKING DOWN 

Figure 20c. S-76 Overhead Panel Vibration Levels for 

Advanced Type Interior. 

66 
FORWARD 

LOOKING DOWN 

I I 

Figure 21a. S-76 Overhead Cavity Levels for Utility 

Type Interior. 

67 
.OOK 

FOI 

:ING 

RWAA 

DOW 

Figure 21b. S-76 Overhead Cavity Levels for Executive 

Type Interior. 

68 
OC67M OC67R 

FORWARD 

LOOKING 

Figure 21c. S-76 Overhead Cavity Levels for Advanced 

Type Interior 

69 
SOUND 

PRESSURE 

LEVEL 

(dB) 

500 1000 2000 4000 

0 OVERHEAD CAVITY LEVELS 

X UNTREATED CABIN LEVELS 

A TREATED CABIN LEVELS 

FREQUENCY, Hz 

Figure 22. Utility Treatment Effectiveness Relative to Overhead 

Cavity Levels and Untreated Cabin Level s. 

70 
SUBSYSTEM 

RESPONSE 

LEVEL 

(d B) 

X OVERHEAD PANELS 

0 OVERHEAD CAVITIES 

Figure 23. SEA Predictions for the Relative Distribution of 

Overhead Skin, Panel and Cavity Levels at 1000 Hz. 

71 
COUPLl NG 

LOSS 

FACTOR 

(d B) 

FREQUENCY, Hz 

PANEL AREA % OPEN 

00C5YM MASS LAW TL 

OOC5YM HOLES 0.08 

msc14~ MASS LAW TL 

mSC14L HOLES 0.13 

0.305 m2 

0.488 m2 

Figure 24. Coupling Loss Factors for Cavities to Cabin Through 

Panels and Holes (Leakage) for a Utility Type Interior. 

72 
As a result, the effects of reducing leakage would be expected to be of somewhat greater significance at higher frequency, as can be seen in Figure 25. A 

doubling or halving of leakage area for all trim panels produces a 2 to 3 dB 

change in the cabin level at high frequencies. Completely removing all of the 

leakage reduces the cabin levels by nearly 10 dB at 4000 Hz, with a diminishing 

effect at lower frequencies where panel transmission becomes of greater importance. Closing of the air conditioning duct openings into the cabin has a 

minimal effect for the Utility interior. 

The effect of changing cavity absorption, as shown in Figure 26, are also 

somewhat greater at higher frequency. A factor of 2 increase in the Sabin 

absorbing area in either the cavities or the cabin produces the same reduction 

in cabin noise level. Increases in cabin acoustic absorption will not be 

effective in reducing cabin noise levels unless the loss factor associated with 

the absorption is comparable to or greater than the coupling loss factor 

between the cavities and the cabin. Greater coupling between the cavities and 

the cabin (i.e., TL plus leakage) at lower frequencies reduces the effectiveness of increased cabin absorption. 

A comparison of SEA predictions with flight data is shown in Figure 27. The 

flight data is an average of many typical Utility interior installations (as in 

Photo 1) in S-76 aircraft. The SEA prediction for 0% leakage is shown only as 

an indication of the maximum cabin noise reduction potential since, by the 

nature of its usage, a Utility type interior cannot be sealed that tightly. 

The measured data are shown to agree well with the expected leakage of approximately 0.08%. 

Table 2. Major Contributors to Power Flow into the Cabin* at 1000 Hz 

Location Bare 

OC23L 13.2 

OC45L 8.8 

OC4YM 10.8 

OCY5M 10.7 

Uti 1 i ty 

5.2 

9.2 

8.2 

11.1 

SC14L 9.9 18.0 

SC56L 3.3 7.3 

Executive 

6.4 

10.2 

7.8 

5.9 

14.4 

10.6 

Executive interior - The contributors to the cabin level in the Executive model 

are quite similar to those of the Utility. Energy flow from the overhead 

cavities accounts for 47.8% of the cabin level. Again SC14L is the greatest 

single contributor providing 14.4% of the total. Because of the similarity of 

the treatment models the effect is still negligible on the overhead frames and 

panel s. 

73 
CABIN 

NOISE 

REDUCTION, 

dB 

500 1000 2000 4000 

FREQUENCY, Hz 

- AIC % LEAKAGE 

CLOSED 0.00 

A OPEN 0.00 

e 

A ‘I 0.08 

0.16 

0 11 0.3 2 

Figure 25. Effect of Air Conditioning Openings and Panel 

Leakage on the SEA Predictions of a Utility 

Type Interior. 
R 

16 

14 CABIN 

NOISE 

dB 

t I I I I 

500 1000 2000 4000 

FREQUENCY, Hz 

Figure 26. Influence of Cavity LOSS Factor on the SEA 

Predictions of a Utility Type Interior. 

75 
CABIN 

NOISE 

REDUCTION, 

dB 

FREQUENCY, HZ 

0- - -0 SEA PREDICTIONS 

H FLIGHT DATA 

Figure 27. Comparison of the SEA Predictions for a 

Utility Type Interior With S-76 Flight Data 

at the Standard Cruise Condition. 

76 
The Executive noise control treatment package incorporates an improved trim 

panel design (7.3 Kg/m2) with greater TL and increased absorption levels in the 

cavities (up 36%) and the cabin (up 18%). Reduced leakage around the trim 

panels is also achieved. The distribution of power flow into the cabin is very 

similar to that for the Utility interior (see Table 2). The power flow from 

the cavities to the cabin occurs with similar relative contributions but with 

reduced absolute levels for the Executive interior. 

TL data for a typical Executive type trim panel is shown in Figure 19, along 

with results from the SEA model of this panel. The panel involves a double 

wall construction with a foam layer between the structural panel and a (lead) 

vinyl septum. The TL data shows a double wall resonance dip near 315-400 Hz 

with rapid increases in TL at the rate of 15 dB/octave at higher frequencies. 

The improved panel TL behavior results in a greater significance for transmission through leakage openings around the trim panels. Coupling loss factors 

for panel TL and leakage in the Executive interior are shown in Figure 28. The 

leakage coupling loss factors are proportional to area. Cavity to cabin 

coupling through panel transmission is significantly less for the Executive 

interior than that associated with leakage, particularly at higher frequencies. 

The SEA predictions show insertion loss values ranging from about 17 dB at 500 

Hz to 24 dB at 4000 Hz (see Figure 29). 

A comparison of SEA predictions with flight data is shown in Figure 30. Again 

the flight data is an average of many typical Executive interior installations 

(as in Photos 2 and 3) in S-76 aircraft. The measured data are in reasonable 

agreement with the expected leakage in this type interior. 

Advanced interior - As mentioned previously, the intent of an Advanced interior 

was to flight test improvements discovered through use of the model. Unfortunately, aircraft avai 1 abi 1 i ty was such that the "Advanced" concepts could 

not be flight tested as a complete set. Many noise control systems have been 

designed and tested in laboratory conditions with good success but have yet to 

be installed onto a flight vehicle. Using the laboratory responses, these 

concepts were applied to the Executive model and a prediction of their effects 

made. 

Measurements were performed on a bare configured S-76 with and without an 

improved frame and panel damping package applied specifically to those subsystems located near the transmission foot mounts. Loss factors of those 

subsystems were derived both through rap testing while on the ground and 

through in-flight measurements. This damping treatment was specifically 

designed to work in the 1000 Hz region, as this is the critical region of input 

frequencies from the main gearbox. The effect of damping appears straightforward enough that it could be applied to the advanced Interior treatment with 

confidence in the results. Verification of this system in an SEA model could 

not be performed as allowances for updating of the previously used bare model 

[Z] to include improvements in the code used for the present models had not be 

made. 

77 
-20 

-40 

a 

0 

0 

COUPL NG 

LOSS 

FACTOR 

(d B) -6C 

-8C 

FREQUENCY, Hz 

PANEL AREA % OPEN 

OC5YM MASS LAW TL 

OC5YM HOLES 0.02 

SC14L MASS LAW TL 

SC14L HOLES 0.03 

0.305 m2 

0.488 m2 

Figure 28. Coupling Loss Factors for Cavities to Cabin 

Through Panels and Holes (Leakage) for an 

Executive Type Interior. 
30 

CABIN 

NOISE 2o 

RED UCTl ON, 

dB 

10 

500 1000 2000 4000 

FREQUENCY, HZ 

A/C % LEAKAGE 

0 CLOSED 0.04 

CLOSED 0.08 

0 OPEN 0.08 

Figure 29. Effect of Air Conditioning Openings and Panel 

Leakage on the SEA Predictions of an Executive 

Type Interior. 

79 
CABIN 

NOISE 

REDUCTION, 

dB 

Figure 30. Comparison of the SEA Prediction, for an 

Executive Type Interior With 5-76 Flight Data 

at the Standard Cruise Condition. 

80 
Table 3. Major Contributors to Advanced Cabin" Model at 1000 Hz. 

Location 

OC23L 

OC45 L 

0 C4Y M 

OC5YM 

SCl4L 

SC56L 

FL 

DCTCZL 

DCTCZR 

Advanced Advanced 

W/Duct C1 osed W/Duct Open 

0.2 0.1 

0.2 0.1 

14.6 8.5 

11.1 6.4 

21.7 

14.5 

12.0 

0.0 

0.0 

12.6 

8.3 

6.0 

23.6 

10.1 

*Shown as a percentage of total power flow into the cabin. 

For use in the model, those specific frames and panels that were treated when 

the damping was tested were assigned a new material model with an increased 

loss factor term as well as specific changes in the subsystem densities and 

wave speeds. 

A low weight, cost, yet effective noise control, is the installment of fibrous 

glass batting into many of the overhead cavities. As was shown earlier, with 

the installation of trim panels, build up of noise in the overhead bays increases dramatically. With more control on leakage in the Advanced model, 

those levels are even greater, as can be seen in Figure 50. By placing glass 

batting into those bays, noise levels can be lowered, thus decreasing the 

amount of energy being forced through the areas of leakage. 

Some of the effects of the added damping can be seen in Figures 2Oc and 21c. 

The damped panels and cavities are much lower in level than their associated 

utility or executive interior level. Cavities OC34L and R are both surrounded 

on five sides with damped frames or panel which apparently is the cause for the 

large decrease in energy. The damping in the frame seems to be the most 

effective, at least according to the model as seen in the response of cavity 

OC4YM as the levels here are lower than the next cavity down, OC5YM, apparently 

due to the frame damping in CF4M. 

The results of the Advanced model shows cabin levels down an additional 6 - 8 

dB over the Executive model. This would be exceptional if these concepts can 

be installed into an interior design for full flight testing with comparitive 

results. So while experimental testing of these noise control concepts and 

installation into the model seems to indicate positive results, they have yet 

to be verified as an advancement over our present interior designs. 

81 
Because of the increased absorption in the cavities and the improved leakage 

control, energy flow from the ducts can now be seen to be quite critical. The 

single largest contributor for the advanced is no longer SC14L at 12.6%, but 

the main supply duct DCTC2L at 33.6%. Levels from the overhead cavities are 

down to 16.0% total contribution. Table 3 shows that with the opening of the 

overhead ducts, all other contributors decrease in significance dramatically as 

if a dam was opened. The ducts can be shown to be an effective noise block as 

total cabin energy increases over 2 db by opening them. 

82 
Table 4a. Summary of S-76 SEA Model: Bare Interior 

Subsection types Number 

pane 1 s 53 

frames 35 

acoustic spaces - 7 

Totals 95 

Junction types 

line 99 

area 86 

point 29 

frame 21 

Total 245 

- 

Table 4b. Summary of 576 SEA Model: 

Subsection types Number 

panel s 101 

frames 70 

i n-pl ane 59 

(trim panel isolators) 

acoustic spaces 64 

pipes - 7 

Totals 301 

Junction types 

1 ine 154 

area 330 

point 230 

frame - 21 

U 

# of D.O.F. in SEA Model 

53 

70 

7 

130 

- 

ility Interior 

# of D.O.F. in SEA Model 

101 

140 

59 

64 

7 

371 

- 

Total 735 

83 
SEA Predictions for the Treated S-76 Cabin 

The primary interest in comparing the bare and treated interiors is in relation 

to the reductions in cabin noise levels that are achieved. The degree of reduction must be compared with the added weight of the total treatments package 

in assessing its effectiveness. Insights into the vibration and acoustic 

transmission processes gained from the SEA model can provide insight that are 

helpful in designing a weight effective treatment. Specifically, regions of 

high modal energy or response can be targeted for treatment, as well as the 

important transmission paths by which energy reaches these regions. Output 

from an SEA model is particularly well suited for identifying contributions to 

the total cabin noise environment due to particular transmission paths and the 

immediate panel structures that are the dominate radiators into the cabin. 

I An important question is the degree to which the treatments affect the transmission in the basic airframe structure, if at all. The utility treatments are 

primarily of the transmission type; trim panels "block" the radiation of noise 

from the skin panels, preventing it from reaching the cabin, directly. Connected to the airframe structure by soft isolation mounts; they are not expected to mechanically load the frames or panels from which they are supported 

to any significant degree. There is the possibility that by preventing the 

noise from radiating directly into the cabin, thereby resulting in a buildup of 

noise levels in the acoustic spaces between the skin and trim panels, the 

higher acoustic levels will mean increased panel vibration levels. Frames, 

which feed energy to the skin panels, would not be expected to show as noticeable an increase as the panels because they are an additional connection 

removed from the cavity acoustic spaces and are therefore less sensitive to the 

increased noise levels adjacent to the panels. 

The preliminary results in Figures 31 and 32 show a trend that is counter to 

this argument, namely, that the frame and panel levels moving forward in the 

cabin away from the gearbox location are higher in the bare cabin than in the 

treated cabin. Panel levels adjacent to the gearbox are more closely comparable. The frames, as expected, show less of a difference between the two 

configurations. Acoustic levels in the cavities immediately beneath the 

forward skin panels, shown in Figure 33, are significantly greater than the 

cabin noise levels in the bare configuration consistent with expectations. 

Cavity levels adjacent to the gearbox are inexplicably less than the cabin 

levels in the bare configuration. 

The apparent anomalies in the comparisons between bare and treated interior 

predictions are currently being investigated. The junctions describing the 

connections for these subsystems are currently being reviewed as to their 

accuracy. 

84 
30 

m 

2 

0 

i= 20 

0 

z 

3 

L 

U 

W 10 

a 

Eo 

L 

2 

- OBARE INTERIOR 

Lfl4L Lf45L L1 

Figure 31. S-76 Frame Transfer Function Levels. 

85 
30 

m 

z' 

F 20 

W 

0 

0 

z 

3 

LL 

$ 10 

LL 

v) 

z 

a 

a 

co 

Figure 32. S-76 Skin Panel Transfer Function Levels. 

86 
Figure 33. S-76 Overhead Cavity Noise Levels. 

Qi’ 
Predicted cabin acoustic levels for the bare and treated configurations are 

shown in Figure 34 as a function of frequency. The levels are scaled relative 

to the same arbitrary input power which does not correspond to actual input 

source levels for either an in-flight measurement or a ground test with shaker 

excitation. The predictions with the utility treatment are from 4-5 dB less in 

the lower frequency octave bands and from 9-10 dB less at 2 and 4 kHz. These 

differences are not as great as expected based on actual measurements on bare 

and treated aircraft and would suggest that a further review of the SEA model 

is needed. 

Power Flow Description of Transmission Paths - The principal effect of the 

treatments in the utility interior is to block direct skin panel radiation into 

the cabin by the addition of trim panels. The VIP interior adds damping to 

panels and absorption to the cavity and cabin spaces to control the levels of 

these subsystems. An important consideration for the treated interiors is 

whether the treatments change fundamentally the mix of important contributors 

to the cabin noise environment or whether all are reduced in comparable fashion. Identifying possible flanking transmission paths that limit the actual 

effectiveness of the treatments is also important in assessing their performance. 

Flanking transmission can occur due to gaps or holes in mounting the trim 

panels or through the isolation mounts that support them. The areas of these 

gaps and/or holes would need to be included in the SEA model in order to 

account for acoustic transmission through them. The predictions would serve to 

quantify the point where they are significant in increasing the cabin noise 

levels, thereby establishing a criteria for quality control in the installation 

of the treatment. 

A similar assessment can be made with respect to the performance of the isolation mounts supporting the trim panels. Frame or skin panel vibration on the 

top of the mount is transmitted through to the trim panel where the resonant 

response of the trim results in a radiation into the cabin. The point attachment to the trim panel results in a nearfield vibration response that also 

radiates into the cabin. The significance of the contribution through the 

mount will indicate whether it is sufficiently soft in relation to the structures it connects. 

Power flow contributions into the cabin are given in Table 2. Cavities are 

directly coupled to the cabin through leakage around and mass law transmission 

through the trim panels. Overhead cavities account for 53% of the total power 

flow into the cabin with the more important contributions coming from cavities 

closer to the excited frame section. The Table 2 data show that the distribution of contributions to the cabin noise level does not fundamentally differ 

from that for the bare aircraft. 
63 125 250 500 1000 2000 4000 8000 

FREQUENCY IN HERTZ 

Figure 34. S-76 Cabin Noise Levels. 

89 
The aft side cavity is the single largest contributor to the cabin noise level. 

It receives essentially all of its power (98.6%) from the side skin panel. The 

power flow into the side panel from the side wall frame is 62% of its total 

power. This side wall frame is a primary structural member down the side of 

the cabin which connects to, and receives 100% of its power flow from, an 

overhead cross-frame just forward of the gearbox. Vibration transmission is 

strong along this important mechanical load path. 

Cavities near the gearbox receive power from the overhead skin panels. Further 

from the gearbox, contributions from adjacent cavities closer to the gearbox 

grow in importance relative to skin panel radiation. The furthest forward 

cavity receives nearly all of its power from the adjacent "upstream" cavity. 

The power flow contributions for a typical cavity acoustic space are shown in 

Table 5. The dominate contributor is the outer skin panel above the cavity, 

with important contributions from adjacent cavities that are closer to the 

source. Transmission from an adjacent cavity is related to mass law transmission through the web of the intervening frame or through openings in the 

web. 

The power flow path from the forward overhead skin panel, OP67L, back to the 

source frame section is shown in Table 6. The transmission behavior in the 

underlying airframe structure is the same as in the untreated S-76. The path 

leads to adjacent frame members and along the main longitudinal frames to the 

source frame section. 

*Shown as percentage of total power flow into the cabin for the particular configuration described. 

Table 5. Power Flow Description for an Overhead Cavity (OC56L) 

Power Inflow From: Percent of Total Inflow 

OP56L 

OC45L 

OC56M 

69.2% 

30.0% 

0.7% 

I 90 
Table 6. Power Flow Path Description for the Utility Interior Cabin Overhead 

for LF14LO Vibratory Input 

Power Inflow for OP67L, from: Percent of Total Inflow 

LF67LO 

CF6LO 

OP56L 

CF6LI 

Power Inflow for LF67L0, from: 

LF56LO 

CF6MO 

CF6LO 

CF6MI 

48.9% 

23.3% 

15.7% 

6.3% 

67.4% 

13.7% 

9.5% 

6.5% 

Power Inflow for LF67L0, from: 

LF45LO 100.0% 

Power Inflow for LF45L0, from: 

LF14LO 98.6% 

91 
ISOLATION AS AN ALTERNATE NOISE CONTROL CONCEPT 

Gearbox Isolation Mounts 

The design of isolation mounts at the attachment location of the gearbox to the 

airframe offers the possibility of potentially significant reductions in the 

cabin noise environment. Mounts are a transmission type treatment which are 

used most effectively when located as close to the source as possible; a 

condition satisfied by gearbox isolation mounts. This section discusses the 

modeling of mount dynamics, accounting for the impedance characteristics of the 

gearbox and airframe structures at the location where they are attached through 

the mount. 

An important presumption made in describing mount behavior is that motions or 

forces at the attachment points and in the different directions can be treated 

as being statistically independent or uncorrelated. This greatly simplifies 

impedance representations of gearbox and airframe behavior. The measured 

effects of source coherence which are described in the Phase I1 report [2], 

lend credibility to this presumption. 

Individual motions at an attachment location involve a single pair of generalized force and velocity variables. The structural connection is presumed to 

occur at a point. The gearbox is modeled as the source and the airframe as the 

receiving structure into which power flows. 

The source description of the gearbox is idealized in terms of an open circuit 

or free velocity, Vfree, where the subscript i refers to the different attachment locations and directions of motion. Treating the gearbox as a linear time 

invariant system with internal sources at gear mesh locations, the free velocity is that which would occur at the gearbox attachment location when the 

internal sources are active and the gearbox has been detached from the airframe. This description is an idealization since a measurement under such 

conditions is an impossibility. It is convenient from a modeling point of view 

because the source description does not depend on the dynamics of the structure 

to which the gearbox is attached. 

g,i 

In the attached condition the force and velocity at the connection are related 

to the free velocity source levels according to: 

Where Y is the input mobility, i.e., the ratio of velocity to force, looking 

into the gearbox at the attachment location with the internal sources turned 

off. 

9, i 

The airframe dynamics are also described by: 

92 
Without the mount V = -V so that the velocity of the airframe is: g,i a,i 

- - 'a.i ,,free 

'a,i - Y + Y g,i a,i 9, i 

The reduction of airframe motion at the attachment location which occurs as a 

result of the addition of an isolation mount is of particular interest. The 

mount connects on one side to the gearbox and on the other to the airframe. It 

is described by two force/velocity pairs of variables: 

-V -z12 g,i (53) 

a,i F a,i z12 -z22 V 

Combining the above impedance matrix description of the mount with the descriptions for the gearbox and airframe an expression is obtained for the ratio of 

airframe attachment location velocities in the case with the mount, V:,i to 

that without the mount, Va,i, as follows: 

V! i (Ya + Yq) z12 

-= 

a,i (1 + Y Z )(1 + Ya ZZ2) - Yg Ya Zf2 V 

9 11 

- When the mount behaves as a lightweight simple spring, then Z 

= Zll - = Z12 and the expression takes on the following simple fovm: z22 

m 

"a,i - - 1 

V 

1 - 

\I 1 I 

m l. 

I+Ya+Y g 

zm (Ya + Yg> 1+ V a, i 

(54) 

(55) 

Where Zm is the stiffness reactance of the mount and Ym is a mobility for the 

mount: Ym = l/zm. 

93 
When the mount is rigid compared to either the airframe or gearbox, i.e., when 

Ym << Ya or Ym << Y then the mount has no effect in reducing the airframe 

vibration. This points out an important feature when stated as a criteria on 

mount stiffness in order to obtain significant reductions; namely, that the 

mount must be soft compared to the softer of either the airframe or gearbox. 

g' 

For given airframe and gearbox structures the inclusion of a mount will result 

in a reduction in the power flow from the gearbox by the following amount: 

dB 1 M = 10 loglo 

'm 12 

9 

I1 + Ya + Y 

This is also the anticipated reduction in airframe response and cabin noise 

levels. 

This evaluation depends on separate estimations of mount , airframe and gearbox 

impedance characteristics. Depending on frequency, the mount can often be 

modeled as a simple spring stiffness, including the effects of cross-sectional 

geometry. At higher frequencies transmission line models of mount impedances 

can be developed. 

When many internal mount resonances are excited, an SEA model of the mount 

provides accurate statistical estimates of mean or average mount behavior. 

This applies also to the airframe structure where a statistical representation 

of attachment point impedance behavior depends on the motion type and details 

of the construction, for instance, whether the attachment is at a mid-point 

location along a frame member or at the intersection or junction of several 

such frame members. 

The evaluation of power flow reduction in Equation (56) can be carried out 

separately from the SEA model of airframe vibration transmission and cabin 

noise. Experimental measurements of structural impedances for the airframe and 

gearbox with shakers and force gages can be combined with estimates of mount 

behavior as modeled analytically or measured experimentally. The estimated 

reductions in power flow apply directly to predictions of vibration transmission and cabin noise response levels from the SEA model. 

94 
Preliminary Design Definition 

Physical constraints - Installation of compliant elements between the main 

gearbox and the S-76 airframe requires that certain analyses be performed to 

determine their effect on various aircraft systems, and safety. These include 

spatial envelope, failsafe design, and allowable static and transient maneuver 

motion between transmission and airframe. Physical size limitations and ease 

of retrofit must also be considered. 

4 

The spatial envelope considered is bordered by the height of the main longitudinal beams, and the flange width of the same, which are approximately 8 and 

3 inches respectively. Given the elastomeric properties and loading information, a preliminary design can be developed. This appears in Figure 35. It 

basically consists of two "hat" style units configured end to end and through 

bolted to the transmission foot. This design provides the redundancy needed in 

case of elastomer failure, as well as high loading capabilities in the vertical 

direction needed to handle maximum maneuver loading. Safety considerations 

dictate the configuration of the encapsulation of the elastomer such that its 

failure precludes catastrophe. Static, transient maneuver and dynamic loadings 

define an envelope of practical designs. Maximum maneuver loading for this 

aircraft occur at the right front attachment location, where -14000 lb. limit 

tension load vertical is encountered during symmetrical dive and pullout, with 

-8000 lb. compression limit load at the left rear attachment in the vertical 

direction during a rolling pullout. 

Dynamic 1 per rev and 4 per rev rotor forces must also be considered for 

fatigue analysis and change in airframe response from the non-isolated to the 

isolated configuration. Normal modes analysis was performed on the af/gearbox 

system in both conditions with spring rates of 20000 lb/in vertical, 110000 

lb/in radial at each attachment, and twice more using 1.5 and 2 times these 

values. This was performed utilizing available finite element model for the 

S-76 to ensure that the vibratory response of the airframe was not detrimentally affected at these forcing frequencies. Hub loads measured at 4/rev during 

full scale wind tunnel testing are as high as -900 lbs. vertical shearing force 

and 12000 in-lb yaw moment during 150 knot forward flight. A more detailed 

analysis of the fatigue characteristics of this installation will be performed 

for validation, using these figures as conservative estimates of actual loadi ng. 

Various mechanical components that have attachment on both the airframe and 

gearbox must be studied so that an increase in relative motion between the 

gearbox and airframe does not hinder their performance. These include fligh 

control rods, input power shafts and power takeoff items such as the tail takeoff from the main gearbox. The minimum stiffness requirement lies in the 

allowable values of error or misalignment for the control system and rotating 

shafting. 

95 
The controls system has mounting locations for control rod pivots on both the 

airframe and gearbox from the controls deck, on the outboard side of OP4-6MY 

and the collective and cyclic pitch actuators mounted on the upper housing of 

the main gearbox. Relative motion here induces error into the flight control 

system, which must be kept to less than 0.1 degree resultant pitch change, 

collective or cyclic. A NASTRAN static analysis of airframehain gearbox installation with and without isolators installed was performed to approximate 

any added error that might be induced. Relative motion of grids adjacent to 

each end of each control rod with static and maneuver loading conditions 

applied is tracked, and resultant error calculated. Analysis of deflection 

between these components has shown that stiffnesses of 132000 lb/in vertical 

and 66000 lb/in radial at each attachment location provide control system error 

within allowable limits. 

Rotating components connecting between the airframe and main gearbox include 

engine input power shafts and tail driveshaft. These components are mounted 

vi a f 1 exi bl e coup1 i ngs designed to encounter a certai n amount of mi sal i gnment 

under normal operating conditions, and provide either a limit isolator stiffness criteria or the need to handle more misalignment via redesign of coupling/ 

shaft arrangements. The possibility of tandem installation is a feasible 

a1 ternate approach 

Shape and size effects - The shape, size, and encapsulement configuration of 

any given elastomer defines an isolators stiffness characteristics. Design 

equations for elastomeric springs along with properties of candidate elastomers 

are used for the final sizing of the units [Ref. lo]. Figure shows the preliminary design shape, with pertinent dimensions labeled. 

For radial loading, 

=k+k 

kY lY 2Y 

where k represents stiffness for each element. 

shear , The top section, in bulk 

and, 

22 AL = 71 [r3 -rl] + nrlh2 

therefore, 

71 G [(ri-rl) + rlh2 

h2 k= 

lY 

96 
where, G = static shear modulus, Pa 

A = Loaded area, m2 

h2 = thickness, m L 

The lower body, radially loaded, 

K1 

- 7.5 A LG - 

k2y In (r2/r1) 

where K1 is a form factor which depends on height to thickness ratio. 

makes the total radial stiffness, 

This 

For axial loading, 

X 

The top section i s i n shear and compression, 

where 

and A is the unloaded area, m2. 

The lower section is in pure shear, 

U 

Making the total axial stiffness, 

97 
Predicted isolation characteristics - The transmissibility and isolation effectiveness of an installed isolator system can be predicted if the dynamics of 

each attachment location and direction is known. Using the method described, 

estimations of the possible benefits of vibration transmission through isolation mounts can be studied. 

The structural characteristics of the gearbox and airframe are represented by 

compliance data obtained from frequency response measurements. Frequency 

response measurements were made on an S-76 gearbox (S/N A-081-00005) and a 

partially assembled S-76 airframe (A/C 244). Driving point frequency response 

measurements and cross frequency response measurements were taken at the four 

attachment locations of the gearbox and airframe in the vertical, lateral, and 

longitudinal directions (see Figures 40 through 44, typical). The data were 

generated using random excitation and were recorded on digital tape for further 

processing. A Hewlett-Packard 5420 two-channel Fourier analyzer was used to 

process the recorded data. The data were analyzed over four frequency ranges-- 

500-2100 Hz, 2100-3700 Hz, 3700-5300 Hz, and 5300-6900 Hz. Each frequency 

response measurement was i ntegrated twice and cal i brated to produce compl i ance 

data with units of meters/Newton. In order to reduce the noise content in the 

data, the frequency response measurements were averaged to produce frequency 

spectra from 500-6900 Hz with a 25Hz frequency spacing. 

Frequency response measurements were made on the gearbox at each of the four 

airframe attachment points in the three principal directions. The frequency 

response measurements are plotted in the form of compliance with units of 

meters/Newton. The attachment point numbering and sign convention for the 

gearbox are shown in Figure 37. Typical driving point measurements are shown 

in Figures 40 and 42. 

The measurement locations and sign conventions used for the airframes are shown 

in Figure 38. Figures 41 and 43 show typical driving point frequency response 

measurements on the airframe at the gearbox attachment locations. 

The isolator characteristics required for the model are the stiffness and damping of each isolating element. These characteristics can be obtained in three 

different ways. The simplest is to use stiffness and damping values which are 

constant or vary linearly with frequency. Another is to use experimentally 

measured values. The third way is to compute the isolator stiffness and damping values using the isolator geometry and actual material properties (modulus 

and loss factor). Experimental studies were performed on several isolator 

configurations. The tests were carried out in the isolation test rig shown in 

Photo 18. A typical isolator is shown in Photo 19 and the test results are 

shown in Figure 44. These experimental results were merged with the analytic 

prediction program to develop an overall methodology to predict isolation 

effectiveness. 

98 
I 

ORIGINAL PAGE B 

OF POOR QUPhoto 18. Isolation Test Rig. 

99 
ORPGI" PAGE IS 

OE P80R QULITY 

Photo 19. Typical Isolator. 

100 
TRANSMISSION FOOT, 

ELASTOMERIC ISOLATOR 

Figure 35. Typical Isolator Instal lation. 

101 
tY - 

I A'* 

STEEL ELASTOMER 

HOUSINGS 

Figure 36. Single Isolator Geometry. 

I 102 
X(+ FORWARD) 4 PORT FORWARD STARBOARD FORWARD 

PORT AFT STARBOARD AFT 

Figure 37. Gearbox Attachment Points and Sign Conventions. 

PORT FORWARD STARBOARD FORWARD 

X(+ FORWARD) 

t @&@ Y (+ STARBOARO) Z 

(+ DOWN) 

PORT AFT STARBOARD AFT 

Figure 38. Airframe Attachment Points and Sign Conventions. 

103 
la 

lb 

RECEIVER 

(AIRFRAME) 

t 

I+ F1 b kb 

VISCOELASTIC 

ELEMENTS 

Figure 39. Block Diagram of Analytic Model. 

SOURCE 

(GEARBOX) 

104 
Figure 40. Main Gearbox Driving Point Frequency Response - Port 

Forward - Vertical. 

105 
1 

1 

t 

E 

. 

FREQUENCY, Hz 

Figure 41. Airframe Driving Point Frequency Response - Port 

Forward - Vertical. 

106 
Figure 42. Main Gearbox Driving Point Frequency Response - 

Port Forward - Lateral. 

107 
10 -10 111t111111111111111111111111111I 

21 00 2900 3700 4500 5300 61 00 6900 500 1300 

FREQUENCY, HZ 

Figure 43. Airframe Driving Point Frequency Response - 

Port Forward - Lateral. 

I 108 
Figure 44. Isolator Bench Test Results. 

109 
Since the isolation system is subject to changes in temperat.ure and experiences 

excitation over a broad frequency range, the use of isolator geometry and 

material properties was found to give the best representation of the isolator 

characteristics. The isolation model was exercised for many constant stiffness 

and damping values and for many isolator geometries. 

The analytical model was used to predict the performance of many isolation 

systems. Each isolation system was composed of one or more identical isolators 

at each attachment point. Initially, the isolators at each attachment point 

were characterized by constant stiffness and damping values based on allowable 

gearbox displacements and desired levels of vibration reduction. Temperature 

and static preload levels were changed to investigate their effects on the 

vibration transmitted to the airframe. The variation of material properties 

over a wide temperature range is shown in Figure 45 over a range of frequencies. An output from the model is the relative displacement of the gearbox and 

airframe attachment points as a function of frequency. The ratio of the 

relative displacement of the gearbox (transmissibility) was calculated for each 

isolator configuration and was used as a basis for evaluating isolator designs. 

Another measure used to evaluate isolation system performance was isolation 

effectiveness. Several isolator designs were evaluated in terms of isolation 

effectiveness in addition to transmissibility. 

To investigate the temperature effects on the stiffness and damping of the 

isolators, transmissibility was predicted for the two isolator systems at 

temperatures of 8OoC, and -25OC. These results are shown in Figure 46. 

The effects of static preload on isolator stiffness and damping were likewise 

investigated. Static preload values chosen from the vertical in-flight isolator loads were 35kN and 70kN for forward and aft flight, respectively. 

Figure 47 shows the predicted transmissibility calculated for 0, 35, and 70 kN 

preloads on the two isolator systems at 8OoC in the vertical direction. 

Isolation effectiveness - Isolation effectiveness is defined as the ratio of 

the the receiver (airframe) displacement obtained when the receiver (airframe) 

is connected with a r i g i d l i n k t o the displacement obtained when an isolator is 

inserted between the source (gearbox) and receiver (airframe). Therefore, an 

isolation effectiveness value of one indicates a rigid connection with no 

isolation and higher values indicate better isolation. Isolation effectiveness 

is a more accurate indication of isolator performance than transmissibility 

because it takes into account the output characteristics of the source (gearbox). Isolation effectiveness equals the reciprocal of transmissibility only 

when the output of the source (gearbox) is independent of the attached load. 

The isolation effectiveness will become worse at structural resonance frequencies because the stiffness of the structure may be as low or lower than the 

isolator. Therefore, isolator performance at a resonance is not always dis- 
J 

3 

0 

0 

5 

w 

a 

a 

a 

E 

v) 

7 1 o2 lot I I I 11 I I 1 1 I I 1 I I 

TEMPERATURE. c 

Figure 45. Variation of Material Properties at Constant 

Frequencies. 

111 
-2 Ill 

10 lll"ll~~~~~~~~~~

500 1300 2100 2900 3700 4500 5300 6100 6900 

FREQUENCY, Hz 

Figure 46. Influence of Temperature on the Predicted Isolation 

Effectiveness with Two Isolators in the Vertical 

Direction. 

112 
2 

10 

70 KN PRELOAD 

35 KN PRELOAD 

--- 0 KN PRELOAD 

_---- 

101 

loo 

*. 

k 

d 

- rn 

g lis1 

3 

E ,p 

z 

U 

2100 2900 3700 4500 5300 6100 6900 10'~ 

500 1300 

FREQUENCY, Hz 

Figure 47. Influence of Preload on the Predicted Transmissibility 

with Two Isolators at 8OoC in the Vertical Direction. 

113 
played by poor transmissibility but is always displayed by poor isolation 

effectiveness. The isolation effectiveness of the isolation system designs 

were calculated relative to a rigid connection of 10l2 N/m. The predicted 

effectiveness of the two- and four-isolator designs are shown in Figures 48 and 

49 for the vertical direction. 

Plan For Validation 

Isolation - The preliminary isolator design has been established in the Phase 

I11 effort. During Phase IV the detailed isolator design will be completed and 

full scale tests carried out. Final definition of the loads envelope and 

aircraft component deflections would be used to set the limits for isolator 

design. These parameters, together with a final spatial envelope will be input 

to the Isolator Design/Prediction Methodology procedures which have been 

developed 'by Sikorsky and validated by model scale tests of small scale isolators. A parametric evaluation of isolator effectiveness will then be made. 

The parameters will include stiffness in axial and radial directions, preload 

conditions, low frequency dynamic quasi-static load effects, material properties, geometry, and temperature. 

A selected isolator set will be fabricated for bench testing and full scale 

evaluation. The bench test will involve a single unit in the half-scale 

dynamic airframe rig. These tests will involve several preload conditions and 

be repeated for several isolators to obtain representative results. These 

tests will provide the verification of the individual loading direction effects 

and isolation effectiveness in a controlled environment. The isolator design 

methodology will then be validated at full scale dimensions with representative 

preload conditions. 

A full scale ground shake test on an S-76 aircraft will determine the isolator 

effectiveness as a group rather than individually. In this test the interaction of all eight isolators (2 at each attachment location between the main 

gearbox and airframe) will be evaluated. Rigid mount inserts of the same size 

as the isolators will be used as a baseline to compare with the isolated case. 

Measurements will be made at the isolator interface (on each side in all three 

directions at each attachment) and on the frames, panels, and in the cabin 

acoustic space. These measurements will be compared to the Statistical Energy 

Analysis (SEA) predictions for frame and panel energy levels. The beneficial 

influence which the isolation system has on cabin noise level reduction will be 

determined. 

The following is an outline for the proposed plan: 

2.5 Phase IV - Validation 

2.5.1 Full Scale Validation of Isolator Analysis 

2.5.1.1 Detailed Isolator Design 

114 
v) 

v) 

W 

z 

W 

> 

c 

LL 

LA 

W 

- 

Y 

Figure 48. Predicted Isolation Effectiveness for Two vs. Four 

Isolators at 8OoC in the Vertical Direction. 

115 
8Ooc, ZERO PRELOAD, VERTICAL DIRECTION 

‘“T 

1000 Hz 113 OCT 

800 Hz 113 OCT 

1250 Hz 1/3 OCT 

ATTENUATION, 

dB 

-20 ’ I 

VERTICAL STIFFNESS PER FOOT 3x105 6xiO5 LBIIN. 

NUMBER OF ISOLATORS PER FOOT 2 4 ISOLATORS 

Figure 49. Predicted One-Thi rd Octave Band Attenuation vs. 

Is01 ator Stiffness. 

116 
2.5.1.1.1 Verify loads envelope at airframe/gerabox interface (quasi-static 

maneuver conditions). 

2.5.1.1.2 Verify allowable relative deflections for critical elements (e.g.: 

engine high speed shaft, swashplate controls, tail drive shaft). 

2.5.1.1.3 Define allowable spatial envelope, and the structural impact, associated with isolator installation. 

2.5.1.1.4 Conduct parametric evaluation of isolator effectiveness as a function 

of stiffness and material properties within the constraints of Items 

2.5.1.1.1 through 2.5.1.1.3 above. 

2.5.1.1.5 Develop the detail isolator drawings and fabricate a selected isolator set (10 units). 

2.5.1.2 Unit Isolator Bench Test 

2.5.1.2.1 Set up and instrument the half-scale dynamic airframe rig. 

2.5.1.2.2 Install single isolator in the rig fixture and conduct isolation 

shake testing at acoustic frequencies at a given preload condition. 

Repeat tests at two additional preload conditions. Repeat tests for 

three additional isolator units to test for fabrication variations. 

2.5.1.2.3 Analyze data from bench tests. 

2.5.1.2.4 Compare bench test data with predicted isolator performance and 

validate overall prediction methodology at model scale. 

2.5.1.3 Full Scale Ground Shake Test 

2.5.1.3.1 Formalize the detailed test and measurement plan including any bench 

test benefits. 

2.5.1.3.2 Fabricate inserts for rigid mount condition. 

2.5.1.3.3 Install rigid mount inserts on an S-76 airframe at the four attachInstrument the 

Measure the vibration 

ment locations between the main gearbox and airframe. 

aircraft and conduct the baseline shake tests. 

and acoustic levels. 

2.5.1.3.4 Install isolators on an S-76 airframe at the four attachment locations between the main gearbox and airframe in place of the rigid 

mount. Instrument the aircraft and conduct shake tests with isolators. Measure the vibration and acoustic levels. 

2.5.1.3.5 Reduce and analyze the full scale test data. Compare with predictions and make recommendations. 

117 
CONCLUDING COMMENTS 

The application of SEA methodology to the prediction of helicopter interior 

noise [l, 21 generated many interesting views of how energy travels from source 

to receiver. This phase of the NASA program has looked at various ways to 

interrupt these varied energy paths. These include interior treatments, panel 

isolation, leakage control, damping, and source isolation. The influence each 

has on aircraft noise performance has been presented in this report and some of 

the key items are as follows: 

e 

I e 

e 

e 

e 

For the first time, a SEA model has been demonstrated for a complete 

aircraft including two different current interior configurations and one 

experimental. 

The coupling loss factor provides the proper information to conduct trade 

assessments between panel TL and leakage control. 

Predictions of interior treatment parametric variations are reasonable and 

consistent with experience. 

SEA predictions of existing designs are in good agreement with flight 

data. 

The modeling process and the predictions generated provide insight which 

will foster intrinsically better designs for future aircraft. 

The frequency response measurements made on the S-76 reveals that the 

airframe is stiffer than the gearbox and contains fewer resonances. 

Resonances in both structures are heavily damped. 

Adequate isolation can be provided at an operating temperature of 8OoC 

using a two- or four-isolator design at each of the four gearbox-airframe 

attachment points. 

Additional isolation can be provided by making the isolators weaker and 

installing positive stops on the gearbox to limit its displacement while 

under peak in-flight loads. 

Stiffening the gearbox will improve the isolation system performance more 

than stiffening the airframe. 

118 
RECOMMENDATIONS 

The benefits of model i ng interior treatment types are demonstrated by the 

ability one has to do design studies and trade-offs. Since the SEA model 

contains information re1 ative to both structureborne and airborne energy f 1 ow 

via various paths, a study conducted early in the design stage of a new aircraft is now possible. The effect of changing the frame/junction points and 

beam terminations could be assessed for their impact on acoustics. 

Another area having strong potential is isolation. The dominant source of 

cabin noise has been shown to be the gear mesh forces generated in the main 

gearbox. These forces transmit through the gearbox housing to the airframe; 

subsequently radiating acoustic energy from those surfaces. The major path is 

structural and i sol ati on interrupts that path. Thus a significant noise 

reduction is possible depending on the constraints of the overall helicopter 

performance, e.g.: swashplate control inputs, engine/gearbox and gearbox/tail 

take-off alignments, and spatial compactness. These items play a role in 

sizing an isolation system, and hence they tend to limit the acoustic reductions achievable. Even with these constraints it is apparent that important 

noise reductions are possible. It is recommended that the initial isolation 

study reported herein be completed and isolators fabricated and tested. A 

recommended program for Phase IV has been described and, pending the outcome, 

plans should be initiated for full scale flight testing of this concept. 

119 
REFERENCES 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

Yoerkie, C.A. ; Moore, J.A. ; and Manning, J. E. : Development of Rotorcraft 

Interior Noise Control Concepts - Phase I: Definition Study. NASA 

CR-166101, May 1983. 

Yoerkie, C.A. ; Gintoli, P. J. ; and Moore, J.A. : Development of Rotorcraft 

Interior Noise Control Concepts - Phase 11: Full Scale Testing. NASA 

CR-172594, February 1986. 

Unger, E.E.; Ross, D.: Damping of Flexural Vibrations by Alternate 

Viscoelastic and Elastic Layers, Proceedings of the Fourth Annual Conference on Solid Mechanics, University of Texas, Austin, Sept. 1959. 

Beranek, Leo L., ed.: Noise and Vibration Control. McGraw-Hill Book Co., 

1971. 

Cremer, L. ; Heck1 , M. ; Ungar, E. E. : Structureborne Sound, SpringerVerlag, 1973. 

Ungar, E.E.: Loss Factors of Viscoelastically Damped Beam Structures, 

Journal of the Acoustical Society of America, Vol. 34, pp. 1082-1084, 

1962. 

DeJong, R.D.: A Study of Pump and Piping Generated Ship Noise Using 

Statistical Energy Analysis. Cambridge Collaborative, Inc., Report to 

DTNSRDC, Annapolis, MD. Contract No. 00600-77-R-1417, 15 May 1978. 

Swift, P.B.: The Vibrational Energy Transmission Through Connected 

Structures. Ph.D Thesis, The University of Adelaide, Australia, May 1977. 

Blake, W.K.: The Radiation from Free-Free Beams in Air and in Water. 

Journal of Sound and Vibration, Vol. 33, No. 4, 1974, pp. 427-450. 

10. Gobel, E. F. (A.M. Brichta, trans. and ed.): Rubber Springs Design, 

John Wiley 81 Sons, 1974. 

11. McGary, Michael C.: A New Measurement Method for Separating Airborne and 

Structureborne Noise Radiated by Aircraft-Type Panels. NASA TP 2079, 

September 1982. 

12. McGary, Michael C.: Noise Transmission Loss of Aircraft Panels Using 

Acoustic Intensity Methods. NASA TP 2046, August 1982. 

I 120 
APPENDIX A 

Laboratory Experimentation 

. 

Test Chambers - Laboratory testing for varification of Panel TL type junctions 

were performed within the Acoustic Test Chambers at Sikorsky's Stratford plant. 

The Chambers consist of two rooms, side by side, with a test section of variable size joining the two (see Figure 2). One room is a reverberant type 

diffuser. The other room is an anechoic type with aborptive wedges on all 

inner surfaces. Both rooms are isolated from their surrounding support stucture as well as from each other. The dimensions of the chambers is of sufficient size to allow for measurements in the required range of 500 to 4000 Hz 

octave bands. The dimensions of the test section opening used in the below 

measurements is 4' x 4'. The test section can be closed for use of individual 

chambers. 

Panel TL - TL of aircraft interior panels was determined using the two-room 

method in the test chambers. Typical panel buildups of the type used in this 

report were placed in the test section between chambers. Sound pressure levels 

were measured on both the incident side as well as the transmitted side. Sound 

intensity (Lit) was measured on the transmitted side directly using a two-microphone sound intensity analyzer. Intensity levels were derived on the incident 

side using 

Lii = Lpi - 6 

where Lii is the intensity level, incident side and Lpi is the pressure level, 

incident side and after checking of the phase error the transmission loss is 

calculated by 

TL = Lii - (Lit + .16). 

where Lit is the intensity level on the transmitted side. Typical responses of 

the panel layups used are shown in Figures 18 and 19. 

Mass law confirms the ability of the chambers to determine TL as when compared 

to a panel of limp mass construction, such as a typical Utility type structural 

panel (Figure 18) the TL response is quite comparitive. 

To verify trim panel response for transmission loss, SEA modeling techniques 

were used to recreate the chambers with an area junction of the size used in 

the actual chamber tests. 

121 
Panel Leakage 

Leakage measurements of typical leakage area in a panel was performed in the 

two test chambers. An ABS plastic panel of the same general construction as 

those of the Utility type panels was placed in the test section. A slot of 

variable total area was cut into the panel to represent a typical leakage between two panels. The leakage area covered for 0.03 to 0.12 percent of total 

panel area with a comparison to a baseline panel with no slot (0.00% leakage). 

Noise levels were generated again in the reverberant chamber and an incident 

pressure level determined. Transmitted noise was measured in the anechoic 

chamber using ASTM 336 type methods and TL determined for each leakage condition. 

In Figure 51 it can be seen that leakage around the trim panels has a strong 

effect on interior cabin levels. How energy flows through leakage area per 

frequency was measured experimentally and then compared to SEA predictions. 

Those predictions were based (as with the TL testing) on a model of the two 

chambers with a panel junction between the two and an additional junction 

directly from chamber to chamber of the specific leakage area. This again can 

be seen to compare quite we1 1. There were some difficulties in matching up 

some of the lower frequencies (below the range of interest) which could possibly be due to the mehod used. It would have been preferred to generate the 

noise on the anechoic side and measured power directly in the reverberant 

chamber, but the chmaber characteristics made it impossible to generate levels 

required for the proper signal to noise ratio. 

Cabin Absorption 

Initially, an attempt was made to measure absorption in the individual components in the cabin and the total cabin absorption would be determined from that. 

measurements of many of those components as they functioned in the aircraft I Some of the measurements made can be seen in Figures 48 and 49. But, practical 

i proved impractical. 

To determine the total absorption within the various cabin interior types it 

was decided that a direct measurement of the cabin as opposed to testing of the 

many various components would provide the mode accurate information. This 

method required only a space averaged sound pressure measurement with a known 

sound power source to derive the absorption. The other method required a very 

the cabin interior. 

, complicated series of absorption tests of the individual components making up 

A calibrated sound power source was placed in both a utility and executive type 

aircraft with interiors equi Val ent to those modeled and pressure measurements 

made. To determine the acoustic loss factor q, calculation of the cabin room 

constant R is required; 

Lp = Lw + lolog ( Qe + 2 ) 

4~r2 R 

122 
where Lp is the measured sound pressure in the cabin, Lw is the known sound 

power emitted from the source, Q6 is the directionality factor, and r is the 

distance from the source. By applying that value with the measured surface 

area St provides the average cabin absorption E using 

R 

St + R 

- 

ci= 

The absorption A in Sabins comes from 

A = SE 

and then applying the volume of the cabin V, the reverberation time T can be 

calculated 

and finally the loss factor 

is calculated for each individual frequency F. This was app 

the cabin material model loss factor. 

ied direct 

123 
-30 :::I UTlL 

VIP 

1 

ADV 

500 1K 2K 4K 

FREQUENCY, Hz 

Figure 50. Relative Comparison of Bare Cabin Levels to the 

Utility, Executive and Advanced Overhead Cavities 

and Cabin Levels. 

124 
-r 

40- 

n 

30-- Y 

m 

m 

0 

-I 

vj 

a 20- 

z 

+ 

10- 

Figure 51. Effect of Increased Leakage Area on Panel TL. 

Ill Ill 11 111 11") 

250 500 1000 2000 4000 8000 

125 
ABSORPTION OF S-76 INTERIOR IN WNS 10 

0 

7 

I 

Figure 52. Absorption Provided by Several Typical Cabin 

Furnishings 

TYPICAL S-76 CABIN ABSORPTION 

D(CUI0INO WNA *IN0 CPUNO TRUTMtM 14 , 

Figure 53. Typical Total Absorption of Cabin Furnishings. 

126 
\ 

\ 

\ 

I 

I 

I 

Figure 54a. Measured Intensity Contours for a Utility Type Panel 

at 0.5 kHz. 

12 7 
6: r 

62 

6 -. 

I 

I I6i 

I 

I 

I 

I 

I 

I 

Figure 54b. Measured Intensity Contours for a Utility Type Panel 

at 1 kHz. 
Figure 54c. Measured Intensity Contours for a Utility Type Panel 

at 2 kHz. 

129 
Figure 54d. Measured Intensity Contours for a Utility Type Panel 

at 4 kHz. 

130 
45 

47 

45 

Figure 55a. Measured Intensity Contours for an Executive Type 

Panel at 0.5 kHz. 

131 
Figure 55b. Measured Intensity Contours for an Executive Type 

Panel at 1 kHz. 
Figure 55c. Measured Intensity Contours for an Executive Type 

Panel at 2 kHz. 

133 
35 

3s 

3s. 

35 

3! 

35 

35 

39 

35 3s 

35 35 

3s 35 

3s 3s 

Figure 55d. Measured Intensity Contours for an Executive Type 

Panel at 4 kHz. 

134 
.. 

Photo 20. Typical Utility Type Panel Being Tested for TL in the 

Test Chambers from Anechoic Side, Sound Intensity 

Probe Shown 

135 
Photo 21. Typical View of Panel TL Testing - Anechoic Side 
-.. 

Photo 22. TYP cal View of Panel TL Testing - Reverberant Side 

137 
Standard Bibliographic Page 

1. Report No. 

NASA CR-178172 

2. Government Accsglion No. 3. Recipient's Catalog No. 

16. Abstract 

The application of statistical energy analysis (SEA) to the modeling and 

design of helicopter cabin and cockpit interior noise control treatment is 

demonstrated. The interior configurations modeled are typical of utility, 

executive, and advanced type treatments. 

SEA models are developed directly from existing S-76 aircraft designs 

currently in production. 

with reasonable acoustic benefits; for the executive interior, emphasis is 

primarily on improving the acoustic environment while a1 lowing for increased 

interior weight. The relative importance of panel transmission loss (TL), 

acoustic leakage, and absorption to the control of cabin noise is shown through 

SEA modeling parametric variation. 

1000 Hz comes from increased TL, while above u)DD Hz it's from reduced 

acoustic leakage and increased cabin and overhead absorption. and aircraft flight noise data are shown to correlate very well with SEA 

predictions. 

established from which new interior design concepts can be evaluated with 

relative ease. A preliminary design was carried out for an acoustic isolation 

of the structureborne energy coming into the airframe from the main gearbox. 

This vibration isolation design shows significant reductions in cabin noise are 

possible. The Phase IV effort will refine this design and include fabrication 

and ground testing. 

The uti1 ity and executive interior 

The utility interior design priority is light weight 

The major cabin noise improvement below 

Laboratory TL 

Thus a viable technique for aircraft noise control design is 

4. Title and Subtitle 

Development of Rotorcraft Interior Noise Control ConConcepts 

C.A. Yoerkie, P.J. Gintoli, S.T. Ingraham and 

cepts, Phase 111: Development of Noise Control 

~ 

7. Author(#) 

J.A. Moore 

3. Performing Organization Name and Address 

Si korsky Aircraft Division 

United Techno1 ogi es Corporation 

National Aeronautics and Space Administration 

Hampton, VA 23665 

Stratford, CT 06601-1381 

12. Sponsoring Agency Name and Addresa 

Langley Research Center 

L5. Supplementary Notea 

~~ 

7. Key Words (Sug eated by Authora(8)) 

Statistical Engergy Analysis Modeling, 

Helicopter, Transmission Isolation, 

Acoustics,Vibration, Cabin Noise, 

Structureborne Noise, Full Scale Vibro/ 

Acoustic Test 

5. Report Date 

July 1987 

6. Performing Organization Code 

78286 

8. Performing Organization Report No. 

, SER-510274 

10. Work Unit No. 

505-6 1-5 1 

11. Contract or Grant No. 

NAS1-16932 

13. Type of Report and Period Covered 

ontr tor R oy $an 85 - JUT! k7 

14. Sponsoring Agency Code 

18. Distribution Statement 

9. Security Classif.(of this report) 

IflPd 

.. 

I Subject Category 71 

20. Security Classif.(of this page) 21. No of Pages 22. Price 

148 Uncl assi f ied 

For sale by the National Technical Information Service, Springfield, Virginia 22161 

NASA Lan~l~y Form 83 (June 108.5)