CF6 jet engine performance improvement program. Short core exhaust nozzle performance improvement concept

TABLE OF CONTENTS Page Section 1.0 SUMMARY 2.0 TNTRODUCTTON 2 3.0 DESCRIPTION OF SHORT CORE NOZZLE CONCEPT DESIGN AND FABRICATION 4.0 4 1 Overall Design Approach 4.2 I)es[gn DuscrLptIO 12 4 3 Core Cowl and Pylon Design 14 4.4 Maintainablity 14 4.5 Reliability 4.6 Safety 5 0 PERFORMANCE TFST 17 5.1 Test Configurations 5.2 Test Facility 5.3 [nstrumcntaLion 5 4 Test Procedure 25 5.5 Test Results 6 0 ACOUSTIC TEST 35 6.1 Test Configuration 35 6.2 Test Facility 6 3 Instrumentation 39 6.4 Test Procedure 6 5 Test Results 43 7 0 ENDURANCE TEST 92 7.1 Test Configuration 92 7.2 Test Facility 92 Procedure 92 7.3 Test 7.4 Test Results 92 Iv TABLE OF CONTENTS Sect Lon Page ASSESSMENT 8.0 ECONOMIC SUMMARY OF RESULTS 9.0 - QUALITY ASSURANCE APPENDIX A OF SYMBOLS APPENDIX B - LIST REFERENCES V 1.0 SUMMARY The purpose of the Short Core Exhaust Nozzle Program was to develop the technology and to demonstrate the technical feasibility of this performance im provement concept on a CF6-50 turbofan engine Back-to-back performance and acoustic tests were conducted in order that direct comparisons could be made between the short core and the long core exhaust nozzles. In addition, development endurance testing was performed for mechanical design assurance.

The sea level performance testing substantiated within test accuracies the expected uninstalled performance improvement-(-0 3% improved nozzle thrust coefficient) which is predicted to be 0.9% internal sfc improvement at cruise. Flight tests conducted outside this program indicate a cruise sfc reduction of at least 0.9% can be obtained with the Short Core Exhaust Nozzle on the Airbus Industrie A300B and the Douglas DC-10-30 aircraft. The acoustic tests demonstrated that this performance gain was achieved without an increase in engine noise. The nozzle hardware successfully completed 1000 simulated flight cycles of endurance testing without any signs of distress.

An economic assessment of the improvement applied to engines on the Boeing 747 and the Douglas DC-10 is included in the report I 2.0 INTRODUCTION National energy demand has outpaced domestic supply creating an increased U.S dependence on foreign oil.

This increased dependence was dramatized by the OPEC oil embargo in the winter of 1973 to 1974 In addition, the embargo triggered a rapid rise in the cost of fuel which, along with the potential of further increases, brought about a changing economic circumstance with regard to the use of energy. These events, of course, were felt in the air transport industry as well as other forms of transportation As a result of these experiences, the Government, with the support of the aviation industry, has initiated programs aimed at both the supply and demand aspects of the problem.

The supply problem is being investigated by looking at increasing fuel avail ability from such sources as coal and oil shale. Efforts are currently under way to develop engine combustor and fuel systems that will accept fuels with broader specifications.

Reduced fuel consumption is the other approach to deal with the overall problem A long-range effort to reduce consumption is to evolve new tech nology which will permit development of a more energy efficient turbofan or the use of a different propulsive cycle such as a turboprop.

Although studies have indicated large reductions in fuel usage are possible (e.g., 15 to 40%), the impact of this approachin any significant way would be 15 or more years away In the near term, the only practical propulsion approach is to improve the fuel efficiency of current engines. Examination of this approach has in dicated that a 5% fuel reduction goal starting in the 1980 to 1982 time period is feasible for the JF6 engine This engine is, and will continue to be, a significant fuel user for the next 15 to 20 years.

Accordingly, NASA is sponsoring the Aircraft Energy Efficiency (ACEE) program (based on a congressional request) which is directed at reducing fuel consumption of commercial transports. The Engine Component Improvement (ECI) program is the element of the ACEE program directed at the fuel efficiency of current engines.

The ECI program consists of two parts: engine diagonstics and performance improve ment. The engine diagnostics effort is to provide information to identify the sources and causes of engine deterioration. The performance improvement effort is directed at developing engine performance improvements and retention components for new production and retrofit engines The initial effort consisted of a feasi bility analysis which was conducted in cooperation with the Boeing and Douglas aircraft companies and American and United Airlines. The study consisted of: * The identification of engine and component modifications which exhibited a fuel savings potential over current practice in CF6 engi no's.

* The technical and economic assessment of the modifications, including the impact on airline acceptability and the probability of production introduction of the concepts by the 1980 to time period as well as their retrofit potential * The assessemnt of fuel savings for the DC-10-10, DC-10-30, and the B-747-200 aircraft.

* The selection of the most promising concepts and the preparation of Technology Development Plans for their development and evalu ation in ground test facilities analysis are reported in Reference 1.

The results of the feasibility One of the concepts selected for development was the Short Core Exhaust Nozzle, hereafter referred to as the Short Core Nozzle. This report presents the results of the development work on the concept.

In 1974/1975, General Electric and the Douglas Aircraft Company conduc ted a series of model tests directed at performance improvement of the CF6-50 core engine exhaust system The CF6-50 engine was initially designed to pro vide core engine thrust reversing, however, many airline operators subse quently deactivated the core reverser or adopted a fixed nozzle system that had the same aerodynamic flowpath but did not provide the reversing function.

The elimination of core engine thrust reversing capability on many aircraft provided the necessary flexibility for design changes and potential perfor mance improvements through reduced internal pressure losses and external drag reduction The model tests cited above confirmed the potential for improvement, and preliminary design studies were initiated by General Electric and Douglas.

Subsequent work effort included additional model tests and full-scale flight tests The additional model tests included wind tunnel tests with wing-on in which Douglas determined an interference drag reduction potential; full-scale tuft and pressure surveys conducted on a DC-10 by Douglas substantiated that the interference drag observed on the model actually exists on the airplane.

In August 1977, a static model test was conducted by General Electric to improve the internal flowpath to achieve the desired nozzle flow area for engine thermodynamic cycle matching. Further wind tunnel scale model tests were conducted by General Electric in October 1977 on the selected configura tion. These tests confirmed the results of the initial tests.

In late 1977, the Short Core Nozzle performance improvement concept was selected for development and evaluation in ground test facilities by the NASA Engine Component Improvement Program because of its high fuel saving potential and high payback for the DC-10-30 aircraft.

The objective of the program was to develop the technology of the Short Core Nozzle system and to verify the predicted fuel savings by full-scale engine ground tests. Mechanical, cycle, performance, acoustic, and instal lation design studies were conducted in support of the engine tests.

In the program, a back-to-back sea level static performance test was con ducted in a test cell on a CF6-50 engine equipped with a Long Fixed Core Noz zle and appropriate cowl doors and the same engine equipped with the Short Core Nozzle and a new core cowl. The acoustic test consisted of a back-to back test of a CF6-50 engine equipped with the Core Reverser Nozzle and the same engine equipped with the Short Core Nozzle in an outdoor noise test facility Endurance testing of a CF6-50 engine with the Short Core Nozzle was performed in order to establish the life capability of the new exhaust system.

The performance test objective was to demonstrate with a back-to-back the overall thrust coefficient improvement of the Short Core engine test Nozzle versus the Long Fixed Core Nozzle configuration to substantiate scale and subsequent data model test results. The objectives of the acoustic test analysis were * To establish the acoustical effect of the Short Core Nozzle on CF6-50 engine noise.

* To assess the impact of the engine modification on community noise levels for typical aircraft approach and takeoff flight conditions The objective of the endurance test was to demonstrate the structural integrity of the Short Core Nozzle by subjecting it to 1000 simulated flight 3.0 DESCRIPTION OF SHORT CORE NOZZLE CONCEPT The Short Core Nozzle is a replacement for a deactivated Core Reverser Nozzle or the Long Fixed Core Nozzle, both of which are in use on the CF6-50 high bypass turbofan engine (Figure 1). A comparison of the Short Core Nozzle with the Long Fixed Core Nozzle is shown in Figure 2. A comparison of the nacelle and pylon for the short exhaust system with the production DC-I0 30 installation is shown in Figure 3.

The Long Fixed Core Nozzle was introduced for DC-10 and 747 aircraft for those airlines which do not require core stream reversing to meet their land ing requirements.

The lighter weight A300B aircraft do not require a core exhaust reverser, and Long Fixed Core Nozzles are used. These nozzles have essentially the same flow lines as the Core Reverser Nozzle Both the Long Fixed Core Nozzle and Short Core Nozzle systems provide significant weight reductions by removal of the deflector structure, blocker doors, and actu ation and position sensing hardware.

As can be seen in Figure 3, the Short Core Nozzle system requires re duced diameter fan flow lines aft of the fan reverser, therefore, recontouring the engine core cowl as well as the core nozzle is needed. The reduced diameters are due to the elimination of the exhaust reverser function. The reverser hardware, in particular the stationary deflectors and reverser actu ators, requires a larger cowl diameter at the engine turbine rear frame.

This requires the boattail angle in the core nozzle region to be approximately 12 degrees with the reverser rather than the 15 degrees, which is possible with the deflector structure removed. The reduced diameter cowling and shorter nozzle, therefore, reduce weight, core pressure loss and scrubbing drag.

This drag and pressure loss reduction along with a recontoured lower pylon fairing was estimated to result in a significant sfc reduction during cruise A weight reduction of 45 kg (100 lbs) over the Long Fixed Core Nozzle, and kg (325 ib) over the Core Reverser Nozzle would be achieved with the Short Core Nozzle.

An assessment of Short Core Nozzle performance improvement was obtained from isolated nacelle model tests at FluiDyne in March 1978. The model test included evaluation of both the Long Fixed Core Nozzle and the Short Core Nozzle to obtain a direct measure of the improvement with the Short Core Nozzle.

The gross thrust coefficients for these nozzles are presented in Figure 4 for the static testing and in Figure 5 for the external flow wind tunnel testing. It can be seen from Figure 4 that the static test demon strated improvements in gross thrust coefficient with the Short Core Nozzle of 0 0036 and 0.0037 at maximum cruise power and normal cruise power pressure ratios, respectively. At lower nozzle pressure ratios, there is more scatter in the test data and the improvement is approximately 0.0035 in gross thrust coefficient.

From Figure 5, it can be seen that the improvement with the short nozzle is 0.0039 in gross thrust coefficient at M 0.82 cruise. The insert in Figure 5 shows that this improvement is approximately 1% net thrust (-1% sfc) at 40,000 N (9,000 ib) of net thrust and 10,668 m (35,000 ft) altitude. This improvement in cruise thrust coefficient is exactly the Mechanism Reverser Nozzle ,ore Reverser Fixed Core Nozzle Long CF6-50 Core Nozzles.

Current 1.

Figure Nozzle Fixed Core Long Nozzle Short Core Fixed Core Nozzle to Long Short Core Nozzle of Figure 2 Comparison A -Core Reverser Nozzle or A-A Nozzle Section Long Fixed Core - CF6-50 Engine with Core DC-10-30 Nacelle and Pylon a) Production Nozzle.

Nozzle or Long Fixed Core Reverser

B C

Section B-B Section C-C > Short Core Nozzle Short Core Nozzle.

CF6-50 Engine with Pylon Fairing - Nacelle and Nodified b) Nozzle Comparison.

Figure 3. CF6-50 Nacelle-Pylon-Core

O Long

Fixed Core Nozzle O Short Core Nozzle la-uo Max. Cruise Ac = 0.0036 T 0.99 S AC T .0.0035 T UU 0 0.98003 HT 0 0.97 (Avg.

of Cruise Points) o E (0 0.96 Remarks 0.95 1. FluiDyne Channel 7 Facility 2. Sting Mounted Installation 3. 9.174% Scale Model 0.94 I I I I 1.4 16 1.8 2.0 2.2 2.4 2.6 2.8 3.0 Fan Nozzle Pressure Ratio, pTFANPo Figure 4.

CF6-50 DC-10 FluiDyne Test Results, Static Test Data, Isolated Nacelle.

Nozzle Fixed Core o Long C Short Core Nozzle 0.25 Data 0.99 Mach power Approach Takeoff CT =0.004 AF = 0.5% AC T = 0.05 AF- = 0.8% 0.98. 0 U L o45 0.97 - 0.82 Data Mach to C)C TAvg.

ACT T 0.9592 o 0.96 - - .-- 0 .0039 o 0.9553 Mach 0.6 Data Hold Condition o 0.95 0 A C = 1 1 10.7 km (35,000 ft) AFN =1.1% AFN 1.0% 0.94 0 I I 0.9% Remarks 6000 8000 I0,000 lb I I I in) Transonic Tunnel 1. FluiDyne 1.68 m (66 40,000 50,000 N 30,000 2, Sting Mounted Installation 0.93 EN 3. 9.174% Scale Model 1 - I 1 1 1 1 0,92 2.8 2.2 2.4 2.6 2.0 1.6 1.8 1.2 1.4 PTFAN/Po Fan Nozzle Pressure Ratio, Test Data, Wind Tunnel Results, FluiDyne Test DC-10 Figure 5. CF6-50 Nacelle.

Isolated improvement that was obtained in the 1975 model test of an earlier version of the Short Core Nozzle. At lower Mach numbers (Mach 0.6 and 0.25 on Figure 5), there is more data scatter but the Short Core Nozzle shows an improvement at all conditions Installation of the Short Core Nozzle is readily adaptable to all CF6-50 series engines on the A300, DC-10-30, and 747 airplanes Utilization of the Short Core Nozzle requires a different core cowl and lower pylon fairing.

4 0 DESIGN AND FABRICATION DESIGN APPROACH 4.1 OVERALL to the CF6-50 Long The Short Core Nozzle is functionally similar Fixed Core Nozzle. In designing the Short Core Nozzle, core cowl and core nozzle contours were established which result in reduced scrubbing area and, the fan reverser and core nozzle therefore, drag The cowl diameter aft of length and diameter were reduced The effect is that the boattail (diver gence) angle aft of the fan was increased from 12 to 15 degrees providing the drag improvement The nozzle is converging-diverging with an area ratio of 1.035.

a minimum of modification to the engine The Short Core Nozzle requires No changes in the fan reverser are made. Bolting flanges for the Short Core are common.

Nozzle, the Core Reverser Nozzle, and the Long Fixed Core Nozzle However, since the core cowl diameter over the turbine rear frame is reduced, the envelope available for the supply and scavenge tubing in the vicinity of the turbine rear frame is also reduced. These tubes were formed to more closely follow the turbine rear frame Lontour than was previously the case Also, the core cowl hinge line on the pylon and pylon apron required modification to accommodate the new flow lines Since the Short Core Nozzle does not translate as does the Core Reverser Nozzle, it no longer is necessary to pro vide a horizontal split line at the fairing juncture with the pylon. Therefore, it is advantageous to support the total fairing directly from the pylon rather than splitting the fairing and carrying half on the nozzle and half on the pylon.

4 2 DESIGN DESCRIPTION The Short Core Nozzle is depicted in Figure 6 The design includes an outer cowl with an integral core cowl support ring, exhaust nozzle liner, forward centerbody, aft centerbody, replaceable core cowl wear pads, baffle, and associated mounting hardware There are no provisions for pylon fairings or pylon fairing attachment structure, these fairings are part of the air frame pylon structure The material used in this lightweight, high temperature structure is Inco 625 for the sheet metal and stiffeners and the rolled and welded flanges.

Wear surfaces are provided by plasma spraying of a wear-resistant coating, tungsten carbide Bolting hardware is Inco 718 due to the high temperature locations and the frequent removal/reasembly operations.

Both the inner and outer nozzle sections and the centerbodies are full 360 degree bodies of revolution The sheet metal components are butt-welded together, and reinforcing "hats" and doublers are attached by brazing /See View E wCowl, Outer Assembly fTi E "U" Channel Linei j Sound Treatment Centerbody, Aft Assembly CF6-50 Short Core Nozzle.

Figure 6.

H4 The sound treatment design Lb the same type as that used on the Core Reverser Nozzle, i.e , the corrugated "top hat" type treatment In this design, the face sheets are perforated with 1.6 mm (0 063 in ) diameter holes onln 60 It gi ,c nt ujg('redt paIl , i In to ohi 'in lhe reclutred 9% _t 2% open area To provide acoustic chambers, 0.19 mm (0 0075 in ) Inco 625 sheet metal is formed into circumferential corrugates to a height of approximately 9.5 mm (0.375 in ) dependent upon the tuning required and brazed to the nonflowpath bide of tie face sheet Fujl height 9 5 mm (0 375 in ) sheet metal partitions placed at every 76 mm (3 in.) of circumference in the corrugate serve to reduce any rotational wave propagation Drain holes are provided in each corrugate at bottom vertical to prevent accumulation of unburned fuel during "hot" starts In the concept utilized, there are approximately 1.44 m (15 2 2 ft ) of sound treatment area in the outer liner and 0 88 m (9 5 ft ) in 2 2 the centerbody for a total of 2.32 m (25-0 ft ) which is comparable to that for the current exhaust system.

4.3 CORE COWL AND PYLON DESIGN Douglas Aircraft Company designed the core cowl and pylon modifications.

The nacelle modification, exclusive of the pylon, involves the area aft of the fan reverser New, steeper external loft lines for the core cowl and core nozzle necessitated a change to the core cowl attachment to the pylon The present center and aft core cowl hinges were redesigned to a lower loca tion so as not to protrude into the airstream. The pylon apron which seals the interface between the core cowl and the pylon was redesigned to relocate it to follow the new loft line The aft pylon fairing was redesigned to match the new short nozzle contour and the aerodynamic lines developed for this shorter exhaust system.

Stress analyses were conducted to establish material gages for strength requirements.

Material selections utilizing aluminum, steel, and titanium were made to minimize cost and weight consistent with the temperature environ ment Detailed design drawings were made for fabrication of the hardware.

Three sets of prototype core cowl doors were furnished by General Elec tric for engine test and the flight test program on the Airbus A300B The Douglas flight test program was conducted with Douglas production cowling.

These flight test programs are described in Section 4.6 Tnqtallation/rework drawingq were made for all the new components so that the production airplanes used for flight testing could be readily con figured for test and reworked back to the original production quality for later delivery to the airplane customer 4.4 MAINTAINABILITY Compared to the Core Reverser Nozzle, maintainability is improved, be cause with the Short Core Nozzle, there are no actuating components, and, therefore, rigging after assembly is not required.

No parts have to be free to translate, thus eliminating fretting and wear.

is no major With regard to the Long Fixed Core Nozzle, however, there change in maintainability with the Short Core Nozzle. Reducing the weight of the outer cowl of the nozzle is expected to simplify handling of that compo nent The cowl door support land on the Short Core Nozzle has replaceable wear strips whereas the Long Fixed Nozzle core cowl support has wear coating plasma sprayed directly on the support cone land and is more difficult to refurbish.

Access to the core cowl compartment is achieved in the same manner as before, through opening the core cowl doors The nozzle is a true body of revolution, and there are no pylon fairings mounted to it. The pylon fairlngs attached to t1f& pylon are not provided with "skirt" extensions, and there is no contact between the cowl surface and the pylon-mounted fairing.

The bolting hardware attaching the outer cowl and the forward and aft centerbodies is made of Inco 718 with silver-plated Waspalloy nuts to accommodate the high temperatures. The outer cowl has lifting brackets attached to facilitate handling.

The-sheet metal and flange material used in the nozzle is Inco 625 which is readily repair welded and requires no subsequent heat treatment to re establish its properties.

The thickness of the aft centerbody and the sound treatment face sheet has been increased from 0.36 xmm (0 014 in.) and 0.46 mm (0.018 in.), respectively, to 0.63 mm (0.025 in.) in order to lessen handling damage.

4 5 RELIABILITY Compared to the Core Reverser Nozzle, the reliability of the exhaust system is greatly improved because of the elimination of the translation mode.

All actuation and position sensors utilized in the turbine reverser were re moved The Short Core Nozzle was designed to achieve a total useful life, with repair, of at least 35,000 projected flight cycles or 50,000 aircraft opera ting hours, whichever occurs first It was designed to operate for the power 4(It (np ind wi thin the Fl iphi envolope defined In the Fngine Model Specifi r111 LOil Maneuver limit load factors for flight and landing, including landing impact, have been established Loads were established consistent with methods used on the Long Fixed Core Nozzle for combining translation accelerations, angular velocities, and thrust or drag. The additional constraint of engine operation with higher inertia loadings due to a fan blade-out was also met The temperatures and pressures to be encountered in the nozzle were taken from the 0F6-50 cycle deck. The applicable General Electric Design Practices were utilized in the design process 4.6 SAFETY A detailed stress analysts of the individual components utilizing the maneuver, pressure and thermal inputs has been completed and documented.

All flight envelope cases analyzed gave a positive margin of safety FAA Air-Worthiness Standards Aircraft Engines Part 33 Revised 10/31/74 and FAA Advisory Circular AC33-1B guided the design.

The construction features and materials utilized are quite similar to the well-proved Long Fixed Core and Core Reverser Nozzles.

In order to substantiate both reliability and safety, a series of engine tests was run.

The endurance test described in Section 7.0 is one such test.

Other tests included the following rype, of Test flposure Ground Test 88 simulated flight cycles Airbus Industrie 19 flights accumulating A3OOB Flight Test hours and 30 minutes on each nozzle plus a total of 10 hours and 18 minutes of ground running Douglas Aircraft Co 382 hours and 15 minutes of DC-10 Flight Test flight testing including 415 total engine cycles cumulative on all nozzles flown For the ground test, the Short Core Nozzle was instrumented with accelerom eters on both the nozzle and centerbody to establish response frequencies from ground idle to takeoff power.

Good agreement with calculated data was obtained and accelerations were well within capability.

This nozzle was then used in the Airbus Industrie A300B flight test pro gram. A total of 21 pressure taps installed on the outer cowl was monitored during flight testing to establish pressure distributions on the cowl.

5.0 PERFORMANCE TEST TEST CONFIGURATIONS 5,1 7 shows the Long Fixed test vehicle was the CF6-50 engine. Figure The wing the pylon fairing simulating the Nozzle configuration including Core Nozzle configuration Figure 8 shows the Short Core position on the airplane.

test The Long Fixed Core Nozzle Core Nozzle Core cowl doors.

with the Short consisted of: hardware * Fan Reverser Core Cowl Doors * Long Fixed Core Nozzle * - Outer Cowl - Forward Centerbody - Aft Centerbody * Nozzle Fairing install the Short Core Nozzle were: Changes made to * Core Cowl Doors Core Nozzle * Short - .Outer Cowl - Forwnrd Cent erbody - Aft Ceaterbody for measurement of A flight-type engine cowling, a test inlet bellmouth were an inlet screen for foreign object damage protection inlet airflow and used for the test.

FACILITY 5.2 TEST Building 500 of General engine test cell 7 in 9 is a schematic of Iigure test facility turbofan or turbojet It is a large Plant.

Electric's Evendale (2000 F) at 907 kg/sec of up to 66* C (150* heaters capable with inlet air sound-controlled are water-cooled with The cell exhausts lb/sec) airflow.

thrust frames can handle Overhead thrust intakes and discharges.

vertical for automtic data All cells are equipped N (100,000 lbs).

loads up to 445,000 speeds analog channels at up to 400 transient recording handling, including data are available per second. Printed to 10,000 channels ranging from 200 reading.

2 minutes after initiation of a within

I

Reverser ..

Nozzle Cowl Doors Core Core Nozzle.

Fixed 7. Long Figure * ow A Figure' 8. Short Care Nozzle.

to V W Sound Treatment

____ rScreen

--- 1-- ...

.

:v I!!, Exhaust Stack - 25 - O.91m

1 (36 in.)

,"AU~--, i \[ Dia. Tubes Tuber Test Stack Inlet Test Chamber u e .

6.1 m (20 ft) Wide x A Wide x 6.1 mnW 6.1 m (20 ft) 5.8 m (19 ft) Long (20 ft) High _-42.7 m (140 ft) Cascade Turning Vanes Figure 9. Cell 7, Building 500.

Engine control'and safety instrumentation is processed by the computer ized control console. This system is depicted in Figure 10. Data are con verted into engineering units and can be displayed on a cathode ray tube (CRT) by addressable call-up pages, bar graphs, or data plots. Limited performance calculations can be accomplished and the results displayed on the call-up pages. Bard copy data are obtained through a teletype unit with capacities as follows: * Pressure - 39 channels * Potentiometer Position - 6 * Frequency - 9 * Thrust - 1 channel * Thermocouple - 43 * Vibration - 8 Steady state performance data are acquired using the on-site Data Manage ment System (DMS). System capabilities are: 345 Pressures - The pressure system consists of eight 48-position scanner valves. Thirty positions are blanks or reference pressures with the number varying by scanner valve.

400 Temperatures - Either chromel-alumel or copper-constantan through the use of CATS blocks (copper alloy thermal sink) reference junctions at the engine facility interface.

There are also 10 frequency-to-voltage converters and 10 frequency counters.

Once the data are transmitted into the control room blockhouse, they are input into the cell computer. The data are converted into engineering units and basic performance calculations are made and printed out on-line. The data are transferred to the site computer at Evendale. The data base manager within the new Data Management System stores and retrieves each data item via its six character data base code..

data are maintained on a large data base within the DMS and the All central computer and can be recalled to be displayed or plotted on the inter active graphic terminals.

5.3 INSTRUMENTATION The following test instrumeantation was used to monitor engine operation and engine perfratnce during performance testing. (See Figure 11 for station designation.)

* Barometric Pressure (PBAR) - The local (control room) barometric was taken using an electronic barometer.

pressure * Cell Pressure (PO) - Test cell pressure using 0-7 N/cm (0-10 psi) differential transducer.

Control Console.

Figure 10. A Computerized * Humidity (HUM) - The absolute humidity in grains of moisture per dry air was recorded using a wet/dry bulb sling psychrom pound of eter * inlet Total Pressure (PT2) - Four 6-element total pressure rakes with 0-7 located in the engine inlet at the fan face and measured N/cm (0-10 psi) differential transducers and pressure scanning valves were used The circumferential locations of these rakes measured from the engine top vertical centerline were 45, 135, 225, and 315 degrees * Inlet Static Pressure (PS2) - Four 6-element rakes identical to the total pressure rakes were used * Compressor Inlet Static Pressure (PS25) - One static pressure tap located on the outer wall of the fan frame core flowpath was recorded. Measurements were made using a 0-10 N/cm (0-15 psi) differential transducer and pressure scanning valve.

* Compressor Inlet Temperature (TM25) - One ungrounded copper con stantan thermocouple, replacing one of the mounting bolts for the CIT sensor, was utilized.

* Compressor Discharge Temperature (TT3) - One ungrounded chromel alumel probe mounted in the condition monitoring port of the compressor rear frame was utilized.

Compressor Discharge Static Pressure (PS3) - A wall static pressure tap was located in a combustor borescope port and measured on a 0-345 N/cm (0 to 500 psi) absolute transducer.

* Low Pressure Turbine Inlet Total Pressure (PT49) - Five 4-element probes were manifolded by probe and measured on a 0-103 N/cm (0 to 150 psi) absolute transducer.

* Exhaust Gab Temperature (T49) - The low pressure turbine inlet tem perature indicating system consisted of 11 dual-immersion chromel alumel thermocouples electrically averaged.

The system was composed of four harnesses which were joined together by means of an aft lead which, in turn, connected to a forward lead. The foward lead had another electrical connector for transmission of the signal to the EGT indicator * Low Pressure Turbine Discharge Pressure (PT5) - Four 5-element rakes were manifolded together and located in the turbine rear frame PT5 was measured with a 0-10 N/cm (0-15 psi) differential transducer and pressure scanning valves.

* Low Pressure Turbine Discharge Temperature (TT5) - Two 5-element rakes were located in the turbine rear frame. The signals were electrically averaged thermocouples.

* Fan Discharge Pressure (PT13) - Four strap-on rakes Two each with four elements and two each with three elements for a total of 70 pressures were read by each immersion and located on the strut in the fan frame. The measurement was made with a 0-10 N/cm (0-15 psi) differential transducer and pressure scanning valve.

* Fan Speed (NI) - Low pressure rotor speed was measured using two fan speed sensors.

* Core Speed (N2) - High pressure rotor speed was measured using an C1gtLue Cott. Sp d sLcaor driv.n off tlie end of OIL tube and scavenge pump.

* Main Fuel Flow (WFM) - Facility engine fuel flow was measured on a volumetric turbine flowmeter.

* Verification Fuel Flow (WFV) - Facility engine fuel flow was measured in series with WFM.

* Fuel Temperature (TF) - Facility engine fuel temperature was measured at the flowmeters using a copper-constantan thermocouple.

* Thrust (FGM) - 222,400 N (50,000 lb) three bridge load.

5.4 TEST PROCEDURE The Long Fixed Core Nozzle configuration was tested first.

Normal pre fire checks, idle leak check, and mechanical checkout were completed. The power calibrations conducted for the test consisted of 15 steady state speed settings from 2093 to 3980 rpm or 61 to 116% corrected speed. At each speed point, two readings were taken after a stabilization time of 3 minutes The first power calibration was completed and the first six points were repeated before instrumentation problems with PT49 (low pressure turbine inlet pressure) and PS3 (borescope compressor discharge static pressure) occurred Further attempts to complete the second power calibration resulted in failure due to icing on the inlet airflow rakes Bad weather forecasted for the next three or four days was the deciding factor in the decision to use the inclement weather period to install the Short Core Nozzle configura tion. The six top points of the second power calibration demonstrated good repeatability of the data and the total points completed were, therefore, con sidered adequate for representing the Long Fixed Core Nozzle charcteristics The Short Core Nozzle configuration was installed and the power calibra tion was completed twice with no further problems.

5.5 TEST RESULTS conditions, the performance indicator At sea level test cell operating for the Short Core Nozzle improvement is primarily the difference in overall gross thrust coefficient. The thrust coefficient is defined as follows- FO CT = F iFan + FiCore) ' where .lcore - Ideal core nozzle thrust bascd on core measured pressures and calculated core gas flow.

FiFan = Ideal fan nozzle thrust based on fan duct measured pressures, inlet total airflow, and calculated core flow.

Two methods were used in calculating the overall gross thrust coefficient dif ference. The first method utilized the scale model data nozzle flow coeffici ent difference. The first method utilized the scale model data nozzle flow coefficients to determine core airflow and the second method utilized the low pressure turbine effective area to determine core airflow. The fan flow was obtained by subtracting the calculated core flow from the total flow deter mined from the inlet total pressure rakes.

Both methods indicated overall gross thrust coefficient improvements in the order of 0 0025 as shown in Figures 12 and 13. The first method is de pendent on accurate physical area measurements of both Long Fixed and Short Core Nozzles and accurate flow coefficient characteristics both in shape and in absolute level from the model test data. Any slight characteristic change between model and full scale hardware is to be reflected as a Root Sum Square (RSS) error in the overall gross thrust coefficient (CT) calculation. Calcu lating the overall gross thrust coefficient with the second method eliminates the above-mentioned potential errors but transfers the potential RSS error to the repeatability of the test instrumentation since none of the hardware changes made influence the low pressure turbine effective area. Because of this, the second method can result in a more consistent gross thrust coeffici ent curve shape but contain the same amount of data scatter as in the first method. In this test, the becond method did result in more Lonsistent curve shape characteristics as indicated in Figure 13.

It can be noted from Figure 13 that the data show an unexplained shift in calculated fan nozzle flow coefficient. This indicates data inaccuracies and, to some extent, appears to validate the 0.3% improvement in overall gross thrust coefficient. However, thrust at fan speed, thrust at engine pressure ratio, and sfc at thrust show conclusive evidence of having demonstrated approximately 0.3% improvement in gross thrust with the Short Core Nozzle.

Figure 14 shows the low pressure turbine pressure ratio comparison be tween the Long Fixed and Short Core Nozzles. This figure indicates the rela tive effective core nozzle throat area for the two nozzles. Note the cross over point of the curve occurs at a low pressure turbine discharge to engine oo00 0 9 short Core Motte 0 Lov V1zed Core Xz.

Is. ao-le pre**..-C Rtio PTln19 Coefficient, Figure 12. Test Cell Data, Overall Gross Thrust Areas and Flow Obtained Using Measured Core Flow Coefficients.

o 97 - short No90le o Short ore tle 11 12 13 1A 15 16 1 7 0 97 ran Nozzle pressure Pti OP rsF.

o 9 0, 0 F Mottle pressure Ctio o Data, Fan Nozzle Flow Coefficient and Figure 13. Test Cell Overall Gross Thrust Coefficient, Core Flow Obtained from Low Pressure Turbine Flow Function.

t4 "o) 38 Equal Effective Nozzle Areas c 37 Ok 36 O Short Core Nozzle C Long Fixea Core Nozzle E 3 5 0) to 14 15 16 17 18 Low Pressure Turbine Discharge Pressure/Engine Inlet Pressure, PT5/PT2 Figure 14. Test Cell Data - Low Pressure Turbine Pressure Ratio.

1.65. This corresponds to a thrust of inlet pressure ratio of approximately Ib). This region of equal effective throat approximately 226,860 N (51,000 areas (a condition of equal ideal thrust) should reflect the improvement due to the Short Core Nozzle in a measured thrust difference and also in an sfc in Figures 15a and b improvement of equal magnitude This improvement is shown established from the test data which indicates that at a corrected fan speed Core Nozzle improvement is approxi corresponding to PT5/PT2 = 1 65, the Short the engine mately 0.33% in thrust of 0.37% in sfc. The thermodynamics of above and below this crossover point of equal effective nozzle areas is in fluenced by the fact that the core nozzle effective areas are different, short core nozzle improvement in param thereby making it difficult to see the eters other than overall gross thrust coefficient.

sea The test results from the standpoint of pretest predictions for level conditions are consistent with A thrust versus fan speed, A thrust engine pressure ratio, and A sfe versus thrust curves shown in Figures versus 17, and 18. The figures show differences based on the Long Fixed Core 16, Nozzle test data since the cycle deck used reflects an average engine and not specific characteristics of the particular engine tested Figure 19 indicates a small amount of adjustment is needed to the cycle to exactly match the effective area characteristic of the two nozzles. The cycle was modified to area characteristics and resulted in model the measured nozzle effective corrected gross approximately the same overall performance improvement in thrust at corrected fan speed and sfc at corrected gross thrust as the model used for pretest predictions.

Power management changes for converting to the Short Core Nozzle are not necessary since the thrust at fan speed increases as shown in Figure 16. A small amount of exhaust gas temperature (- 3 C) margin can be realized, how ever, if the improvement in thrust at corrected fan speed for the Short Core Nozzle is used to lower the power management.

The following summarizes the results of the performance test 1. The full scale engine back-to-back Long Fixed Core versus Short Core Nozzle testing indicates an improvement in overall thrust coeffi cient of approximately 0 3% Test data, in the region of equal effective exhaust nozzle areas, show improvements in gross thrust at engine pressure ratio and at fan speed, and improvements in sfe at thrust for the Short Core Nozzle configuration. The close agree ment between full scale and model test data at sea level verifies the 0.35% overall gross thrust improvement with the Short Core Noz zle as determined from scale model tests.

2. Based on the agreement of full scale with model test results at sea level, the model test results simulating altitude operation can be used for estimating "on-wing" Short Core Nozzle improvements.

3. The Short Core Nozzle does not require a power management change to meet minimum engine thrust at fan speed.

K Equal Effective Nozzle Area

_ ..

0.

o4

'00 0.33

0 Avg. % 'G&52

Q 235 240 220 225 230 - RN Gro0s Thrust, FQG62 corrected I I I 0 I 52 53 x 10 50 51 lb Thrust, F M62 - Corrected Gross with Test Cell Data, Thrust Improvement Figure 15a.

Nozzle.

Core Short Effective Nozzle Areas 05 -Equal U -04 0.37 A sfc Avg.

_0 -02 235 240 225 230 215 220 kN Gross Thrust, FGM/62 Corrected I I I I I I : 53 x 10 49 50 51 52 lb Corrected Gross Thrust, FGM/& Improvement with Short Test Cell Data, SFC Figure 15b.

Core Nozzle.

55X (i45 3 235 -0- - Short Core Nozzle A 5Z ------ Long Fixed Core Nozzle 230 Cycle Deck Pretest Prediction for Short Core Nozzle ,o 215

48 /

7 210 46 220E o45 ~20 -1951

/

170] I i/ 831 32 33 34 35 36 37 38 39 40 x 10 CotXOeted Pan Speed, NIK, rp, Figure 16. Test Cell Data, Corrected Gross Thrust vs. Corrected Pan Speed.

10345 x so a,.z 220 48 "215 48 .205 47 0 U) o a 44 / -- 0D---- Short Core Nozzle 190 " --- O-]---- Long Fixed Core Nozzle

/Cycle Deck Pretest

Prediction for Short 38 170 I I p 50 55 60 65 70 Engine Pressure Ratio - PTdo/ Figure 17. Test Cell Data, Corrected Gross Thrust vs. Engine Pressure 32c Ratio.

---

0,001 (0.01) H/ 0 Short Core Nozzle core Nozzie E] Long Fixed Pretest Cycle Deck H for Short ~Prediction Nozzle o Core 220 230 240 250 170 180 190 200 210 F GM/16 - N Corrected Gross Thrust, 54 56 x 103 48 50 52 42 44 46 36 38 40 Corrected Gross Thrust, F GM /S ibs Specific Fuel Consumption vs. Corrected Figure 18. Test Cell Data, Gross Thrust.

3 9 Ll H a) 0 Short Core Nozzle Z Long Fixed Core Nozzle - .---.-.- Cycle Deck Pretest 0Prediction for Short Core Nozzle 14 15 16 17 18 Low Pressure Turbine Discharge Pressure/Engine Inlet Pressure, PT /PT 5 2 Figure 19. Test Cell Data, Low Pressure Turbine Pressure Ratios.

TEST 6.0 ACOUSTIC 6.1 TEST CONFIGURATIONS A series of static back-to-back noise tests was conducted on a CF6-50 engine with the Short Core Nozzle and the Core Reverser Nozzle.

The production CF6-50 engine was fitted with a reference acoustic inlet and bellmouth lip and production fan and core exhaust duct acoustic treatment.

A description of the acoustic treatment is shown in Table I below: Table I. Acoustic Treatment.

Location Treatment Type Treatment Area 2 2 (SDOF) 5.57 m (60 ft ) Inlet Single Degree of Freedom Fan 2 2 Multiple Degree of Freedom (MDOF) 5.85 m (63 ft ) Fan Casing 2 2 (SDOF) 4.65 m (50 ft ) Duct Single Degree of Freedom Fan Exhaust 2 2 4.37 m (47 ft ) Degree of Freedom (MDOF) Multiple 2 2 2.32 m (25 ft ) Core Nozzle "Tophat" (SDOF) Short 2 2 1.95 m (21 ft ) Core Reverser Nozzle "Tophat" (SDOF) The engine was configured with the advanced fan blades and a smooth micro balloon shroud with a fan tip clearance of 1.9 mm (0.075 in.).

All performance rakes were excluded from the fan inlet, fan exhaust, and core exhaust ducts for these tests. A comparison of the flow lines and the acoustic treatment between the Core Reverser and Short Core Nozzle is made in Figure 20.

6.2 TEST FACILITY The static back-to-back noise tests were performed at the General Elec tric Peebles Test Operation Site 4D at Peebles, Ohio. The site is paved with concrete extending a minimum of 6.1 m (20 ft) beyond the microphone positions.

The acoustic field is free of obstructions for 45.7 m (150 ft) minimum dis tance beyond the far field microphone locations. The engine was mounted to a thrust frame supported by an open-trussed cantilever structure as shown in Figure 21. The engine centerline was located 4 m (13 ft) above the concrete.

2 2 Acoustic Treatment Area - m (ft ) Core Reverser Nozzle - 1.95 (21) Short Core Nozzle - 2,32 (25) Core Reverser Nozzle Short Core Nozzle Figure 20.

Compaaison of Short Core Nozzle to Core Reverser Nozzle.

il iii iiii Sound Field.

F'igure 21. CF6-50 Acoustic U,0 Portable Microphone Locations

Engine Cj Height of 4 m (13 ft) 30

at

4 )r

(150 ft)

45.7 m

-IL 150

TO CONTROL 51.8 m+ ROOM (170 ft) +130+

+12

PORTABLE .125 ENVIRONMENTAL 1 5 DATA WEATHER CONCRETE TEST PAD Layout.

CF6-50 Sound Field Figure 22.

cartridge was replaced and, prior to each test, a 124 dB The microphone of Standards, was applied to pistonphone, traceable to the National Bureau compared to the most recent The microphone sensitivity was each microphone.

assure compliance within ±1.5 dB; any sys laboratory calibration data to replaced. The microphone outputs were then tem falling outside this band was lev attenuators in order to record the same voltage normalized Using variable the source input. At the conclusion of each test, el with the pistonphone and the voltage level was recorded as a verification pistonphone was feapplied of microphone system integrity.

occasions throughout the test series, 2-minute recordings of On several These record were made with "facility on" and "facility off".

ambient noise settings used during the sound measurements to assure ings were made at gain acoustic data.

acceptable signal-to-noise ratios for the Turbine Sound Separation Probe A water-cooled sound separation probe was used to record dynamic pres transducers at the core exhaust nozzle. Two Kulite pressure sure fluctuations spaced 12.7 cm (5.0 in.) apart in a line parallel to the engine axis were on a 0.95 cm (0.375 in.) diameter tube. Water flowed through flush-mounted thereby greatly extending the temperature the tube, cooling the transducers, transducers to determine operating range. A laboratory calibration of the prior to probe assembly.

pressure response was performed two transducers was positioned 90 degrees to The probe tip containing actuator to permit a radial traverse.

the probe stem which was installed in an 1.27 cm (0.5 The downstream transducer was positioned inside the core nozzle to the concrete pad in.) from the exit plane. A box beam support attached the ground at the 9 o'clock engine posi held the probe stem horizontal to aft looking forward. A shield was positioned around the actuator in tion, the fan exhaust flow.

order to reduce buffeting due to when testing the The probe aft transducer was immersed to four positions 16.0 cm relative to the outer wall).

Short Core Nozzle (2.0, 6.4, 10.8 and (30 ips) for 1 minute at each immersion.

Data were recorded at 76.2 cm/sec speed changes. The A 2-minute stabilization time was allowed between engine fan speeds of 2207, 2400, 2598 and data were recorded at average corrected to and following the test, a 3.45 N/cm (5 psi) static 2841 rpm. Prior pressure was applied to the rear face of each Kulite diaphragm to verify probe calibration Atmospheric Test Condition Instrumentation pressure was recorded for each test point. Wind speed, direc Barometric all measured using a Portable Envi tion, air temperature and dew point were of these stations were located approxi ronmental Data Station (PEDS). Two degrees from the inlet on a 51.2 m (168 ft) arc (Figure 22). The mately 45 ft) height. Wind speed and direction sensors were positioned at a 4 m (13 charts. The were recorded continuously on strip measured on one of the PEDS wind.direction instrumentation.

incorporated wind speed and V cosO second PEDS These signals were also recorded continuously on strip charts. Ambient tem perature and dew point temperature were measured by aspirator resistance tem perature devices. The dew point measurement was made with a hygrometer which sampled air from the 4 m (13 ft) location. These measurements were all re corded on the DMS computer system.

6.4 TEST PROCEDURE Atmospheric Test Condition Limits Atmospheric condition limits were set prior to the test. Any data re corded outside these limits (listed below) were discarded: Relative humidity 20% < RH < 95% Temperature 2640 K (-90 C) < T < 3050 K (320 C) Headwind <4.1 m/sec (including gusts) Crosswind <2.6 m/sec (including gusts) Tailwind 0 m/sec (including gusts) Gusts <1.5 m/sec Engine Test Conditions The two engine configurations were run to obtain data for comparisons at the same corrected thrust over a range of conditions that encompass the ap proach, cutback, and takeoff power ranges for aircraft powered by the CF6-50 engine. The nominal test conditions consisted of 19 fan speeds covering the range of 2090 to 3905 rpm equivalent to the thrust range of 53,632 to 234,421 N (12,057 to 52,700 Ib).

For each configuration, the 19 test conditions were repeated twice in the same order for a total of three readings at each power setting. A shutdown of at least 30 minutes occurred between each test series. At each power setting, the engine was stabilized for at least 2 minutes prior to recording acoustic data.

The engine performance data were corrected to standard sea level pressure, zero humidity, zero wind day using the measured atmospheric data for ambient temperature and pressure, absolute humidity, wind velocity, and wind direction.

Acoustic Data Recording Acoustic data were recorded on magnetic tape using a 28-channel FM tape recorder system operated in the interrange instrumentation group (IRIG) wide band Group I mode at a tape speed of 76.2 cm/sec (30 ips). The recorder was

set up for 40% carrier -deviation CU 40%) at full scale record level. Signal

amplification was provided by a a.c/d.c preamp module. During testing, the tape recorder input and output were monitored to assure that adequate ampli fication was used and to assure proper operation of the recorder. Data were recorded for at least 2 minutes at each speed point.

Acoustic Data Reduction Off-line data reduction was performed using an automated 1/3 octave re duction system. The recorded data were played back on a 28-track system oper ating in the IRIG wide band Group 1 mode.

In the automatic operating mode, control of the system was provided by means of a minicomputer and operator-provided information. The data to be sampled were located by means of a time code reader, indexing from the time code signal recorded on the data tape. This tape-shuttling was continued for each data channel with sampling performed over the same time increment until all channels of a particular reading were processed. The system then advanced to the next data point, based on the operator-supplied time reference, and re peated the shuttling process. After the processing information (including reading identification, reading time, gain changes, etc.) was set up by the operator, the system ran without further operator assistance until a magnetic tape change was required.

All 1/3 octave analyses were performed using a 1/3 octave analyzer. The frequency range of the data reduction process was 50 Hz through 10 kHz. A normal integration time of 32 seconds was used to provide adequate sampling of the low frequency portion of the data signal.

The data sampling for the spectrum analysis was done within the 52-second time interval for which the average, performance and ambient weather conditions were determined.

Each data channel output was passed through an interface to the mini computer where data were corrected for both the frequency response of the acquisition and reduction system (as determined from the pink noise calibra tLon) and for the microphone head response. The minicomputer was interfaced to a main frame computer to generate a file containing the 1/3 octave band data for further processing. The 1/3 octave band data were also punched on paper tape as a backup for the communication interface system.

The noise data at each test point were processed using a digital computer program to normalize the data to a 298 K (770 F)/70% relative humidity stan dard day using the atmospheric data (ambient pressure and temperature and relative humidity) and perform data extrapolations to various sideline dis tances. Overall sound pressure level (SPL), perceived noise level (PL) and tone corrected perceived noise level (PNLT) were computed for each angle at the sideline distances. The sound power level for each 1/3 octave band and -overall sound power level was computed for each test point. These results were used for subsequent analysis and data comparisons.

6.5 TEST RESULTS Measurement Accuracy The transducers utilized in the acoustic test are of two types: 1.27 cm (1/2 in.) microphones and Kulite miniature pressure transducers. The accuracy of any sound measurement is dependent on the accuracy of the acoustic data re cording and reduction system which is dependent on the tolerance of each inde pendent component in the system. A list of each component and the accuracy (3 a tolerances) obtained from the respective manufacturers is presented below: Component 3o Tolerance (± dB) (Microphone Cartridge Calibration 0.2 f < 10 kHz Cathode Follower Amplifier 0.2 Far Field Pistonphone 0.2 Microphone Noise Generator 0.5 System Power Supply 0.09 Variable Gain Amplifier 0.2 Tape Recorder 0.5 Common to (Tape Deck 0.5 Both 31/3 OB Analyzer 0.25 Systems Computer 0.0 0.0 Minicomputer Kulite (Kulite Pressure Transducer Cali- System jbration 1.5 Voltage Source, d.c. 0.2 Since these variances are independent of each other, the estLmated vari ance of the sound level can be computed as the RMS of the variances due to each component separately. The accuracy of the far field microphone system is then ±1.0 dB and that of the Kulite system is ±1.7 dB.

The accuracy of the system does not define data reproducibility which is dependent on many other factors.

The intrinsic variation of acoustic data due to meterological conditions, source variations, random instrument error, etc., defines data sample variance about the "true" absolute level of the noise source under evaluation. Hence, the instrumentation accuracy defines the tolerance (systematic error) on a "true" noise level determined from test sample statistics (random error).

Noise level differences obtained from static back-to-back tests remove any data bias introduced as a result of instrumentation systematic error.

The remaining random error can be collapsed using sample statistics to de termine the statistical significance of noise level differences determined from the test data. The conclusions drawn from such an analysis are inde pendent of instrumentation system absolute accuracy.

the Core Rever was postulated that the change from Prior to the test, it noise through the Short Core Nozzle could modify engine ser Nozzle to the cycle, change in principal mechanisms- change in the thermodynamic following change in jet noise.

turbine (LPT) noise, and suppression of low pressure testing and daLa analysis to from the subject back-to-back static Results configurations are discussed differences between the engine evaluate small herein.

Cycle Effect on Noise as determined during versus corrected fan speed data, Corrected thrust agreement was maintained compared in Figure 23. God this test series, are Thermodynamic test runs and between engine configurations.

between repeat apparently do not (for acoustic cycle differences between the configurations relationship. Therefore, affect the fan speed/thrust purposes) significantly to cycle differences.

effect can be attributed no acoustic Far Field Perceived Noise to reference sidelines The 45.7 m (150' ft) arc data were extrapolated was used to cor 305 m (1000 ft). Spherical divergence of 122 m (400 ft) and Practice ARP866A (Reference 2) rect for distance, and Aerospace Recommended These reference distances was used to correct for atmospheric absorption.

altitudes for ap because they are typical of FAA certification were chosen noise levels were computed at and takeoff, respectively. Perceived proach of the effect of the Short Core these distances to allow a general assessment on the far field acoustic data.

Nozzle relative to the Core Reverser Nozzle between the engine configurations, all To facilitate data comparisons at the and comparative plots wer6 made. PNL directavities data were averaged II are exhibited takeoff and approach power settings listed in Table typical should reflect in Figures 24 through 27. The aft quadrant data comparisons the engine changes that are related to differences between any noise level only configuration change was the core nozzle. Dif configurations, since the inlet noise levels, which are dominated by fan ferences in front-quadrant Settings Selected For Data Presentation.

Table II. Nominal Engine Power Nominal Fan Speed Sideline Distance Nominal Thrust Condition m) (ft) kN (lb) rpm 3850 Takeoff 305 228 51,200 Cutback 305 1000 158 35,600 3275 High Approach 122 400 88 19,800 2610 Low Approach 122 400 59 13,300 2190 x 103 45 52200-* N C11

Z'40- 9)80

4j 0 CF6-50 with Core Reverser Nozzle' 43 =160 - SCF6-50 with Short Core Nozzle C o 30 -bU AM j 25 o!

$4 U0100 15.6 I I I 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 Corrected Fan Speed, NIK rpm Figure 23. Corrected Thrust/Corrected Fan Speed Relationship.

C"i C o 3-

Z

~ § 3 Figure~~ ~ ~ C24-. wecie osLvlDrciit Coepevrser ozzlet oeozl and ~ ~ ~ Core Reverser Nozzle Co50wt a 0 I00f)Sdln n 2 and~ Co-rO ~ withShorsCor Nozzleat35m(,0f)Sienend28Y 51,2006-5 with Shortte Corus Nozzlfe) (5,0zb orce hut(aef) r4l 0E M-0 Nozzle 0 CF6-50 with Core Reverser c with Short Core Nozzle

C3 CF6-50

Z 8S C

"b 20 40 60 Sb IbO 10 140 ISO' lbo

Angle to Inlet, degrees 25. Perceived Noise Level Directivity Comparison of Short Core Nozzle Figure Nozzle at 305 m (1,000 ft) Sideline and 158 kN and Core Reverser (35,600 lb) Corrected Thrust (Cutback).

C M - 00 OI z O CF6-50 with Core Reverser Nozzle El CF6-50 wth Short Core Nozzle 0-

20 40 60 sbo too

120 140 16 IO

Angle to Inlet, degrees Figure 26. Perceived Noise Level Directivity Comparison of Short Core Nozzle and Core Reverser Nozzle at 122 m (400 ft) Sideline and 88 kN (19,800 lb) Corrected Thrust (High Approach).

C

e El

- ' r CF6-50 wth Short Core Nozzle

20 40 60 so IOU 20 140 160 ISo

Angle to Inlet, degrees Figure 27 Perceived Noise Level Directivity Comparison of Short Core Nozzle and Core Reverser Nozzle at 122 m (400 Et) Sideline and 59 kN (13,300 lb) Corrected Thrust (Low Approach).

noise, are assumed to have been influenced by changes in meterological condi tions, possibly atmospheric turbulence, even though all tests were run within stringent wind limits. The PNL directivity behavior at the 305 m sideline distance for takeoff power settings is similar between the two configuraLons.

This result implies that the Short Core Nozzle does not significantly affect jet noise sound levels nor directivity patterns.

PNL directivity behavior at the 122 m sideline distance at approach power settings represents a superposition of the three primary noise components of the engine: jet noise, fan and low pressure turbine (LPT) noise. Of these sources, the Short Core Nozzle should influence only the noise signatures of the LPT and the core jet. Differences in the fan noise characteristics be tween the two configurations should be independent of the core nozzle since the engine thermodynamic cycles are essentially identical. The small differ ences in the aft quadrant that occur between the configurations for approach power settings at the 122 m sideline distance (Figures 26 and 27) may be due to a difference in LPT noise levels. The apparent differences observed here are discussed further under component noise analysis.

PNL data at typical peak forward and peak aft angles determined from the directivity plots are shown versus thrust in Figures 28 through 31 for 50 and degrees at both the 305 and 122 m sideline distances. No significant trends are discernible for the front quadrant angle. The PNL versus thrust data for the Short Core Nozzle at 115 degrees diverges from the Core Reverser Nozzle data at low power settings but is similar to the Core Reverser Nozzle data at high power settings.

This behavior is the only difference that can be attributed to the Short Core Nozzle when the engine configuration data are compared on a PNL basis. It is possible that the LPT directivity pattern has been influenced by the Short Core Nozzle producing the observed behavior. How ever, this result could only affect aircraft flyover noise at the very low power approach conditions.

Far Field Noise Spectra Averaged 1/3 octave spectra at angles of peak PNL (50 and 150 degrees) for the power settings listed in Table II are presented at the 305 and 122 m sideline distances in Figures 32 through 39. Noise signature component dif ferences are discussed in a later section while salient features common to each are discussed below.

The post distinct feature common to all the noise spectra is the "null" that falls between the 160 to 315 Hz 1/3 octave bands. This is an inter ference phenomenon produced by the sound field. The minimum level associated with the "null" occurs in the frequency band where the path length difference between the direct path from the source to the microphone and the reflected path from the concrete surface of the pad is one-half wave length. For the 45.7 m arc microphones at 4 11 centerline height, the minimum level occurs at approximately 250 Hz. Reinforcement of the engine noise occurs when the path length difference is one wave length with the maximum level occurring at approximately 500 Hz. Higher order mimina and maxima effects are washed z a) > O CF6-50 with Core Reverser Nozzle . [] CF6-50 with Short Core Nozzle Zr Thrust -kN 50 F75 wit 125 15R 175 20 Nzz5 -I I I I I . I i' I, 0 3 so 35 4 45 50' 55 x 1 1is2a 2 Corrected Gross Thrust, pounds Figure 28. Perceived Noise Level Comparison of Short Core Nozzle and Core Reverser Nozzle as a Function of Thrust for 50 Degrees at 305 m (5000 ft) Sideline.

cn3 P4 -3 Oz O CF6-50 with Core Reverser Nozzle > [] CF6-50 with Short Core Nozzle Thrust -kN 75 100 125 150 175 200

"1 is 20

25 3 b '35 40 45

"bbc1

Thrust, pounds Figure 29. Perceived Noise Level Comparison of Short Core Nozzle and Core Reverser Nozzle as a Function of Thrust for 115 Degrees at 305 m (1000 ft) Sideline.

NASA

FORMAL

REPORT

FFNo 665 Aug 65 1 le 0 El P-4 Z CD 0O CF6-50 with Core Reverser Nozzle -0 CF6-50 with Short Core Nozzle C.o C Thrust - kN 100 125 150 175 200 225 50 75 I I' - 0 ". I I , ' , 45 5sx 10 (011 20 25 30 35 40 Thrust, pounds Short Core Nozzle ane Core Noise Level Comparison of Figure 30. Perceived m for 50 Degrees at 122 as a Function of Thrust Reverser Nozzle (400 ft) Sideline.

C), M C 0,0 A0 0O CF6-50 with Core Reverser Nozzle 0- Thrust -kN 50 75 100 125 150 175 200 2:25

_- I

I I I

1 ',

i

1 3

25 3b 3b 40 45 so 5 x 1O0 '1 s20 Thrust, pounds Figure 31. Perceived Noise Level Comparison of Short Core Nozzle and Core a Function of Thrust for 115 Degrees at 122 m Reverser Nozzle as (400 ft) Sideline.

' with Core Reverser Nozzle 0 CF6-50 z Nozzle CF6-50 with Short Core I

V

o r. - N a a) R N B o

a)

0N Vl ~ l l l ~lfl Sl l l ~ l l l l

501C 0 00500{50 § 310 00

FrCuny,) Reverser1 Nozl 0a 0 eg 0ee3 5 m(000 ft)S5dl0 n an50 80800N (51,200 it) Corrected Thrust (Takeoff).

0-0 M E31 El 0 C6-0wBhShr Cr Nzl , El 0 CFG-50 with Core Reverser Nozzle

E5 CF6-50 with Short Core Nozzle

50 1Z5 $is 800 2000 8000 £2500 31500 86006 Frequency, Hz Figure 33. 1/3 Octave Spectrum Comparison of Short Core Nozzle and Core Reverser Nozzle at 115 Degrees 305 m (1000 ft) Sideline and 228 ks (51,200 lb) Corrected Thrust (Takeoff).

0.

C)- El tj MD EDB = El rEl c) w m O CF6-50 with Core Reverser Nozzle "[] CF6-50 with Short Core Nozzle M 0 125 315 800 2000 6000 12500 31600 80000 Frequency, Hz Figure 34. 1/3 Octave Spectrum Comparison of Short Core Nozzle and Core ft) Sideline and 158 kN Reverser Nozzle at 50 Degrees 305 m (1000 Corrected Thrust (Cutback).

(35,600 lb) C :3 Ela C-4 $4 H CO with Core Reverser Nozzle O0 CF6-50 C CF6-5O with Short Core Nozzle 80000 K0 5000 12500 800 2O 000 50 125 Frequency, Hz Nozzle and Core of Short Core Spectrum Comparison 1/3 octave Figure 35.

305 m (1i00 ft) Sdelne and Reverser Nozzle at 115 Degrees (Cutback).

158 kN (35,600 lb) Corrected Thrust m

SsS

S

o

10) E3l diC Reverser Nozzle CF6-50 wth Core SO Nozzle with Short Core CF6-50 r 60000 12500 31600

Boo 2060 . . . B bo

50 126 M1 Frequency, Hz CAC Nozzle and Core of Short Core Spectrum Comparison Figure 36. 1/3 Octave ft) Sideline and 88 at 50 Degrees 122 m (400 Reverser Nozzle Thrust (High Approach).

kN (19,800 lb) Corrected 0- W W

>B

I0 0 CF6-50 with Core Reverser Nozzle o CF6-50 with Short Core Nozzle 50 125 $15 800 2000 5000 12500 31500 8606ad Frequency, Hz Core Nozzle and Core Figure 37. 1/3 octave Spectrum Comparison of Short ft) Sideline and Reverser Nozzle at 115 Degrees 122 m (400 (19,800 lb) Corrected Thrust (High Approach).

88 kN El E z0 0 2 0 El 0o 0 CF6-50 with Core Reverser Nozzle

E] CF6-50

with Short Core Nozzle Frequency, Hz Figure 38. 1/3 Octave Spectrum Comparison of Short Core Nozzle and Core Reverser Nozzle at 50 Degrees 122 m (400 ft) Sideline and 59 kN (13,300 lb) Corrected Thrust (Low Approach).

Qo QD E Z >M C3N O CF6-50 with Core Reverser Nozzle [] CF6-50 with Short Core Nozzle C:) so 125 315 BOB 2000 Boca 12500 315od $606 Frequency, Hz of Short Core Nozzle and Core Figure 39. 1/3 Octave Spectrum Comparison Reverser Nozzle at 115 Degrees 122 m (400 ft) Sideline and (Low Approach).

59 kN (13,300 lb) Corrected Thrust out due to atmospheric turbulence, the extended nature of the source, broad band noise characteristics of the source, the intrinsic tLime variation of the source noise levels, and the band width used for data analysis. Interference effects above 1000 Hz are negligible when compared to the data scatter. No attempt was made to correct the data for interference effects. This is justi fied by the reproducibility of the interference effect and the direct compari sons made between the back-to-back static test results. Data corrections for this effect would have been required if the tests were performed under radi cally different environmental conditions.

The fan blade passing frequency (BPF) fundamental and harmonics are pre dominant features of the far field noise spectra at both sideline distances for all engine power settings. These tones are easily identified in the spectra. The bands that are affected are summarized in Table III for the takeoff and approach power settings listed in Table II.

Table III. Source IdentLificatLion of the Far Field Tones.

Source Identification, kHz Nominal Fan Fn____ P Fan LPT Condition Speed, Ni (rpm) -()F1- 2FI 3F 4FI LPT3 Takeoff 3850 (2)2.5 5.0 8.0 10.0 8.0 Cutback 3275 2.0 4.0 6.3 8.0 6.3 High Approach 2610 1.5 3.15 5.0 6.3 5.0 Low Approach 2190 1.25 2.5 4.0 5.0 4.0 Notes.

() Fl = Fan BPF fundamental 2FI Fan BPF second harmonic 3F1 = Fan BPF third harmonic 4FI = Fan BPF fourth harmonic LPT3 = LPT third stage BPF fundamental (2) 1/3 octave band containing the specified tone, kHz The LPT third stage fundamental BPF also contributes to the far field noise levels. The bands affected by the LPT are also listed in Table III.

The LPT third stage BPF tone is in close proximity to the third harmonic BPF of the fan, always fallng in the same band. Hence, the two tones cannot be differentiated in the 1/3 octave spectra.

The far field noise spectral comparisons tend to reinforce the observa tion made from the PNL data comparisons that the Short Core Nozzle has affect ed the directivity of the LPT third stage tone,. The far field data at 115 degrees for low power settings show the largest differences for both the PNL and spectral comparisons. The directivity of the 4.0 kHz band data for the low approach power setting at ,the 122 m sideline is shown in Figure 40 for both engine configurations. The relative effects of the core nozzles are shown by the data at the aft angles. The forward angle data appear to be controlled by the third harmonic of the fan BPF.

The spectrum level differences for the forward angle data between the configurations are not significant. The 4.0 kHz band is dominated by the fan BPF third harmonic for angles less than,80 degrees. The tone level of the fan BPF third harmonic in the forward quadrant is dependent on a rotor-turbu lence interaction for static testing without a turbulence control screen.

Consequently, large differences between the 1/3 octave bands containing the tone may occur if atmospheric conditions (turbulence) differ between test series. However, the spectrum level of the 4.0 kHz band for the aft angles is not controlled by the fan tone.

This band is controlled by the LPT third stage and fan broad band noise which is discussed in the Low Pressure Turbine Noise paragraph which follows. The 5 kHz band width spectra shown, n Figures 41 through 50 show that the LPT third stage "haystack" and fan broad band noise are controlling the band level.

Jet Noise Before'the test series was performed, it was postulated that jet noise could be affected by the change in core nozzle geometry even with no change in the magnitude of the fan and core jet velocities. Shortening of the core noz zle tends to reduce turbulence and could reduce jet noise slightly. The in crease in boattail angle was expected to have no significant effect on jet noise level or directivity. Data analyses to evaluate the above effects are discussed here.

I One-third octave band sound power level (PWL) data for the engine con figurations are compared in Figure 51 at two typical takeoff power settings.

The noise signature of the engine is seen to be dominated by jet noise (fre quencies below -1000 Hz). As shown in Figure 51, there was no significant change in the sound power level between the engine configurations.

Noise spectra at a 305 m sideline distance for takeoff power are compared in Figures 52 through 59 for 130 through 160 degrees angular locations. Again, there were no significant differences observed between the two engine configu rations.

m a E Z N MU Lfl - Nz

0 CF6-50 with Core Reverser Nozzle

o 0] CF6-50 with Short Core Nozzle

,% zb 4b so eb faoao i'o so leo

Angle to Inlet, degrees Figure 40. 4.0 kHz 1/3 Octave Band flirectivity Comparison of Short Core Nozzle and Core Reverser Nozzle 122 m (400 f) and 59 kN (13,300 lb) Corrected Thrust (Low Approach).

90- I IA I s a 0 10 0 Frequency ka r Figure 41. Short Core Nozzle Far 45.7 m Field Spectra (150 at 100 ft) Arc Degrees, 88 kN (19,800 lb) Corrected Thrust (High Approach).

I 80

o C,, - 80 10 0 40 60 0 20 Frequency - kHz 105 Degrees, Field Spectra at Nozzle Far Short Core 42.

Figure Thrust lb) Corrected Arc 88 kN (19,800 45.7 m (150 ft) Approach).

(High 80 100 40 60 0 20 Frequency - kilz at 110 Degrees, Nozzle Far Field Spectra Figure 43. Short Core Thrust (19,800 lb) Corrected 45.7 m (150 ft) Arc 88 kN (High Approach).

go - 90 (ig A a %,., I.

, d) 0 0 2 0 40 60 80 100' Frequency - kHz Figure 44. Short Core Nozzle Far Field Spectra at 115 Degrees, 45.7 m (150 ft) Arc 88 kN (19,800 ib) Corrected Thrust (High Approach).

5D 5 0 ., .

14 800 I ] 8 * 1 (0Jg Aprach.

9 SC 9O0 o 90 C g CC S80 Ac Cd C70 ,'I J, 60.1",Ni 0 20 40 60 80 10 Frequency -klz 020 4 0 6 0 30 1O00 - kl Frequency Figure 46. Core Reverser Nozzle Far Field Spectra at 100 Degrees, 45.7 m (150 ft) Arc 88 kN (19,800 ft) Corrected Thrust 60. (High Approach).

o

j r 0 a01

70qunc k~ U.

Degrees, Far Field Spectra at 105 Core Reverser Nozzle Fgure 47.

Thrust 88 kN (19,800 it) Corrected 45.7 m (150 ft) Arc (High Approach).

"go t . . , c70 ' jiji f~' IAd va60 6 0 8 0 100 0 2.0 4 0 Frequency - kllz Far Field Spectra at 110 Degrees, 48. Core Reverser Nozzle Figure ib) Corrected Thrust ft) Arc 88 kN (19,800 45.7 m (150 (High Approach).

' = o It 7)

il - I

5 210 1 4'0 6.'D B'0 10 0( Frequenc -lfz 115 Degrees, Far Field Spectra at 49. Core Reverser Nozzle Figure Thrust Arc 88 kN (19,800 Ib) Corrected 45.7 m (150 ft) (High Approach).

8o0 50a.

41 I,I

so

S60 0 2.0 46 0 a 0 0 Frequency- kt Figure 50. Core Reverser Nozzle Far Fi]eld Spectra at 1i20 Degrees, 45.7 m (150 ft) Arc 88 kN (19,800 lb) Corrected Thrust (igh Approach).

7o Short Core Nozzle W aCore Reverser Nozzle FN/ = 227 kN (51,000 2 ib) o i6o cu P M

a

4) 140o 0) , 170 P-4 r FN/6 = 218 kN (49,000 ib) 0 160.

'50 :j

W

cd 150'J U E- 14 0)11 iO00 10000 Frequency in Hertz Figure 51.

Power Level Comparison of Short Core Nozzle and Core Reverser Nozzle on CF6-50 Engine at Takeoff Power.

BAND NO 29 30 31 32 33 24 35 SA 37 38 39 40 41 4 1 17' 18 19 20 21 22 23 24 25 26 27 2Q i00 t,.i -. - - - - -.-.- - - Short Core Nozzle - ----------- -G-Core Reverser Nozzle 0 a)' IN go 04 a a @ 0 - -r - - 5 1 OFPORoA~y7 ) FREQUENCY IN HERTZ Nozzle and Figure 52, 1/3 Octave Spectrum Comparison of Short Core Core Reverser Nozzle at 130 Degrees 305 m (1,000 ft) (Takeoff).

~Sideline and 217 kN (48,800 ib) Corrected Thrust

aMNLPAGE

I

BAND NO 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 1I00 'a Short Core Nozzle; o 0 Reverser Nozzle 0 -Core 0; OfO 8o SO , 70 6oo 0l 4o I-I + - F U I ZH

o50

Figure 53. 1/3 Octave Spectrum Comparson of Short Core Nozzle and Core Reverser Nozzle at 140 Degrees 305 m (1,000 ft) Sideline and 217 kN (48,800 lb) Corrected Thrust (Takeoff).

BAND NO 36 37 38 39 M0 41 12 33 I4 35 27 2S 2V S0 31 22 23 2. 25 26 17 18 19 20 21 1C0 16 =t -Short Core Nozzle Core Reverser Nozzle 9 F o 0 IIt 0- "- 0 .: 80 -~ - U.

- - - - 4 - - - - - - - - ,t LZ - - o x iI '033

7-b

-4

'I ~~ ~ ~ ~ ~ idln an 27 215-(48,tO62 800 l)-25 Corrcte Thru3-35-st-0-300 (Taeof) 10 00 FREQUNCY I HERT O NO 31 32 33 34 35 36 37 38 39 - 41 4C 16 47 18 19 20 21 22 23 24 25 26 27 22 29 30 Short Core, Nozzle =- _ -- Core Reverser Nozzle o 90 o g Oz o8 0; 05!

o - ) 60 I'-H o5 i00 10000 FREQUENCY IN HERTZ 0 Figure 55. 1/3 Octave Spectrum Comparison of Short Core N(2zzle and Core Reverser Nozzle at 160 Degrees 305 m (1,600 ft) Sideline and 217 kN (48,800 lb) Corrected Thrust (Takeoff).

BAND NO 16 17 18 19 20 21 22 22 24 25 26 27 28 29 30 31 32 33 34 35 3 37 3 9 41 d :m - - - L.

.

o 100

-Short

Cord Nozzle - O-Core Reverser Nozzle z 52 £- - - - - -

9 90

_ 0 0 PQm 80 I&i II 0 50 - 40-..0--6.--0--10--25_1.;0 2, , 25.:11-40 -60-8G -10 -125| -25-31.--&0- 0. 0 -- (0--3 W-6 0 5_ 2w 11"ET5q5 10000 FREQUENCY IN HERTZ Figure 56. 1/3 Octave Spectrum Comparison of Short Core Nozzle and Core Reverser Nozzle at 130 Degrees 305 ma (1,000 ft) Sideline and 228 kN (51,200 lb) Corrected Thrust (Takeoff).

DND NO 16 17 1 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 3S 34 37 38 39 40 41 42 S -~1 - -. . . . . . . . . . . .

i00 - Short Core Nozzle

z9:Core Reverser

Nozzle 90 =t70 S80 S70 4-) M Cc 0 50 630 80 0 1 2 5 A)O- 50-6 - M'125 T -200-250 315-40-9 - -1 o 0--3---00- - -- |2 5 2,, 100 i000 10000 FREQUENCY IN HERTZ Figure 57. 1/3 Octave Spectrum Comparison of Short Core Nozzle and Core Reverser Nozzle at 140 Degrees 305 m (1,000 ft) Sideline and 228 kN (51,200 lb) Corrected Thrust (Takeoff).

BAND NO 26 27 28 29 30 31 32 3334 35 SA37 3a 39 sO_ 41_ i oo ( 16 17 18 19 20 21 22 2324 25 I0

- - oShort Core Nozzle

- Core Reverser Nozzle - - - - - - d U U - -- - ------ --- ------- xU -F 10000 10 1000 L IN HERTZ IFREQUENCY and Comparison of Short Core Nozzle Figure 58. 1/3 Octave Spectrum .

305 m (1,000 ft) Nozzle at 150 Degrees Core Reverser Corrected Thrust (Takeoff).

and 228 kN (51,200 ib) ~Sideline e: BAND NO 36 37 38 39 40 41 42 29 30 31 32 33 34 35 =1 100 22 23 94 25 26 27 28 16 17 12 19 20 21 Core Nozzle - - _--Short Core Reverser Nozzle o $4 PQ ci > 0 1-i -- - - ' 40---- - 10000 IN HERTZ FREQUENCY and Core Nozzle Comparison of Short Spectrum 59. 1/3 Octave Figure (1,000 ft) Degrees 305 m Nozzle at 160 Core Reverser (Takeoff).

Corrected Thrust (51,200 ib) and 228 kN Sideline Low Pressure Turbine Noise LPT noise suppression resulting from the acoustic treatment in the core nozzle was expected to be similar between the Short Core Nozzle and Core Reverser Nozzle configurations based on previous rectangular duct testing at temperatures and flow velocities which simulated the engine cycle. The rec tangular duct testing, however, did not account for the complex shape of the real core duct. The design of the acoustic treatment proper was not changed between the core nozzles. Unfortunately, test scheduling difficulties, due to adverse weather conditions, did not permit comparative tests using the sound separation probe to evaluate treatment suppression performance between the nozzles.

Results from a 5 Hz band width spectrum analysis of the core nozzle data and from a narrow band analysis of the 45.7 m far field data at 100 through 120 degrees for a typical high approach power setting (88 kN) are shown in Figures 60 and 61. The Short Core Nozzle core probe data (immersion No. 1), as shown in Figure 60, clearly shows the fundamental BPF tones associated with each turbine stage, and the dominance of the third stage tone. By comparisons, the far field data show no discernible turbine tones.

The phenomenon of "haystacking" as observed in the far field narrow band spectra for the third stage tone has received considerable attention in the literature discussing turbine noise generation. Only the salient features of the mechanism will be discussed here. A detailed discussion of the phenomenon is given in Reference 3.

The transformation of the third stage turbine BPF tone into a "haystack" observed in the far field spectra is an apparent result of scattering of the acoustic wave by the core and fan jet turbulence. The observed spectral broadening and amplitude modulation is hypothesized to be primarily produced by propagation through the jet interface and shear layers. In addition to spectral broadening, the turbulence scattering centers act as the source of the scattered wave. The far field data will exhibit a Doppler shift dependent upon the relative velocity of the turbulent boundary layers at each far field microphone location. For conditions where the energy of the scattered wave is comparable to or less than that of the incident tone, the far field data will exhibit a haystack and an observable tone. The tone will not exhibit a fre quency shift even though the peak frequency of the haystack exhibits a Doppler shift.

The far field properties of the third stage turbine BPF, as shown in Fig ures 41 to 50, are essentially as described above with both engine configura tions exhibiting similar behavior. The fourth stage BPF fundamental tone is evident at 105 degrees for the Short Core Nozzle (Figure 42). Evidently, either there is an unknown mechanism that governs the far field propagation of this tone which differs from the scattering process described above, or the haystack amplitude is below the broad band noise in this frequency region.

A composite of pressure levels which represent the envelope of pressure amplitudes measured at the eight transducer positions (see Section 6.3) is Z

x

L Io >' 4 4 4 0=60

I 8

20 40 60 80 Frequency -kH Figure 60. Short Core Nozzle Core Probe Spectrum at Immersion No.

88 kN (19,800 lb) Corrected Thrust (High Approach).

-4 cii ot 2400 NIK a > 2598 NIK 2841 NIK 6.0 8.b 4.0 2.0 k z Frequency, Pressure Measured Sound Short Core Nozzle Probe Figure 61.

20 Hz Band Width.

of Eight Spectra, Level Envelope shown in Figure 61 for four corrected fan speeds spanning the range of ap proach power settings. A 20 Hz band width tracking filter was used to reduce the data. The blade passing frequencies of the low pressure turbine (128, 126, 112 and 86 blades, respectively, for the first through fourth stages) are seen ad-distinct tones. The third stage fundamental is always a predomi nant feature of the core probe spectrum. Second harmonics of the tones as well as several sum and difference tones produced by tone modulation through blade rows are also observed. Broad band levels remained fairly uniform across the nozzle, but tone levels varied significantly, even between adjacent positions.

The power levels of the strongest tones were calculated from the loga rithimic average of all the pressure measurements and are shown in Figure 62 for the four values of corrected fan speed. The duct cross section was as sumed to be composed of eight equal area regions with uniform sound pressure level distributions over each area. The power level was calculated as shown below.

PWL = SPL + 10 Log A + 20 Log (1.0 + 0.707M) + 10 Log (P /ps //T ) + 10 dB o where. PW = space average PWL, dB re I0- 3 watts 5 2 = space average SPL, dB re 2.0 x 10- N/m SPL the nozzle (m ) A = area of M = duct Mach number Po/Ps = ratio of ambient to duct static pressure T/To = ratio of duct total temperature to ambient temperature The spectral broadening of both the third stage low pressure turbine BFP fundamental and the fan BPF third harmonic tones in the far field results in an overlap of the broad band noise from each component. Correlation of the far field data with the source is inconclusive due to the nonlinear scattering process, preventing the determination of the spectral content due to each component in the frequency regions of interest. Hence, direct evaluation of treatment performance between the two confLgurations could not be performed.

Any conclusions regarding treatment effectiveness drawn from comparing the small spectrum level differences in this frequency region should be regarded as questionable.

Community Noise Impact Assessment Community annoyance as a result of noise produced from aircraft operation is inherently subjective. Annoyance produced by aircraft noise is dependent on sound level, spectrum content, tone content, time duration, and an individ C C) -.t w -±0 Z H Z 9- a. a.C4 a4 04 V.

0.41 - -3 1 ( N 140 A 4-) ! 120 en a) 110 _ 2841 NI1K 2-98 111K o P4 '130 110 --- Nozzle.

Level for the Short Core 62. Duct Measured Power Figure ual's perception of the noise. In order to assess community noise annoyance as related to a statistical population sampling that accounts for all of the above variables, the effective perceived noise level (EPNL) was evolved. Cur rent FAA regulations governing aircraft noise are based on EPNL. The regula tions specify maximum EPNL as a function of gross takeoff weight for sideline, takeoff (flyover), and approach.

To evaluate the impact of the Short Core Nozzle on the CF6-50 engine com munity noise levels, simulated EPNL values were analytically obtained using the static noise data from this test. It was assumed that t e static PNL data were measured flight levels recorded at typical aircraft velocities. No corrections were applied to estimate flight effects on engine noise. Simu lated EPNL values were calculated at power settings typical of the DC-10-30, B-747, and A300B aircraft for approach and takeoff operating conditions.

These power conditions and associated altitudes and velocities ire presented in Table IV.

An analysis of the simulated EPNL values was performed to determine sample statistics for the population mean and variance for each engine con figuration (CF6-50 with the Short Core Exhaust and the CF6-50 with the Core Reverser Nozzle). To assess the community noise impact of the Short Core Nozzle engine configuration, the differences between the average EPNL for each engine configuration at the same thrust were determined and the resulting delta evaluated against the hypothesis that the engine configurations are acoustically equivalent.

A least square fit of the EPNL data as a function of thrust was deter mined for each engine configuration using an orthogonal polynomial regression model. The EPNL data were then interpolated to the nominal power settings shown in Table V to remove data bias produced from thrust set-point scatter.

Sample statistics for each power setting were calculated to obtain the sample mean, unbiased variance for the sample, and the pooled variance for the entire engine configuration data sample. The differences between the sample means (delta) were used to determine the change in community noise obtained with the Short Core Nozzle. The statistical parameters obtained from this analysis that were used to construct confidence bands for the deltas between sample means are summarized in Table VI.

The hypothesis that the engine configurations are acoustically equivalent was evaluated by comparing the magnitude of the deltas to the 90% confidence band for the difference between sample means obtained from the pooled variance, centered about zero. Results of this evaluation are shown in Figure 63. The Short Core Nozzle noise levels are "statistically equivalent" to those of the Core Reverser Nozzle if the delta value lies within the 90% confidence band, "statistically less than" if below the lower band limit, and "statistically greater than" if above the upper band limit.

As shown in Figure 63, the deltas between the subject engine configura tions are always within or below the 90% band limits. Based on the above analysis, these results imply that community noise levels for the CF6-50 with the Short Core Nozzle are equivalent to or less than the CF6-50 with the Core Reverser Nozzle engine configuration.

Table IV Typical Flight Operating Conditions for CF6-50 Engine.

Thrust Range (Max/Mn) Altitude (Max/Min) Flight Velocity (Max/Min) Condition kN lbs m ft m/sec knots Takeoff 230/200 52,000/46,000 610/305 2000/1000 103/93 200/180 Cutback 170/150 38,000/34,000 610/305 2000/1000 103/93 200/180 High Approach 100/65 23,000/15,000 122/113 400/370 85/77 165/150 Low Approach 70/50 16,000/12,000 122/113 400/370 85/77 165/150 ~-4 Thrust Set Point For Data Groupings.

Table V. Nominal Nominal Flight Velocity Thrust Condition Nominal Altitude lbs m ft m/sec knots N 51,248 227,962 48,741 Takeoff 305 1000 103 216,810 204,489 45,971 168,921 37,975 305 1000 103 200 158,477 35,627 Cutback 33,382 148,491 22,860 101,686 94,654 21,279 Approach 122 400 85 165 88,079 19,801 High 81,958 18,425 76,291 17,151 72,341 16,263 Low Approach 122 400 85 165 68,053 15,299 59,094 13,285 53,632 12,057 Table VI. StatistLcal Parameters.

Engine 90% Confidence Band Condition Configuration 62 v N t(v) Limits on AEPNL CF6-50 0.0280 19 Takeoff w/Short Core Nozzle and Cutback CF6-50 0.0595 19 3 w/Core Reverser Nozzle Pooled 0.0438 38 1.684 ±0.29 CF6-50 0.1473 25 3 w/Short Core Nozzle Approach CF6-50 0.1572 26 3 w/Core Reverser Nozzle Pooled 0.1523 51 1.678 ±0.54 Notes 02 Pooled unbiased estimate of sample variance Up = Pooled unbiased estimate of configuration sample variance v= Degrees of freedom of sample N = Number of data samples comprising sample mean at each nominal thrust t(v) = "c" statistic for v degrees of freedom, 95th percentile for H AEPNL = 0, ±t(v) ap 90% Confidence Band limits O " He = Statistical hypothesis N = Number of data samples comprising sample mean of Core Reverser Nozzle at each nominal thrust N2 = Number of data samples comprising sample mean of Short Cord . .. Nozzle at each nominal thrust N N 0 -0.8 .900 Confidence Interval Low Approach High Approach on LEPNL = 0.0

s- .0.6

Takeoff Cutback "....

.....

(1) +0.4 ............................. o ii ....

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

0.:...............

iiiii.... .... .... iii !!!i !!!i!!i~~iliili~ll ................

.. : :i

H . .::::.:.:::.:.:.:.:.:.:.: . . . . . . .

. ..

PO -0.2

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

-0.4 o.......................................x Q) 1z -0.8 Thrust -R 220.0 240.0 140.0 160.0 180.0 200.0 60.0 80.0 100.0 120.0 P4 10.0 15.0 20.0 25.0 30.0 35.0 40.0 45.0 50.0 55.0 Thrust - lb With the CF6-50 Engine Figure 63. Change in Community Noise Obtained Short Core Nozzle.: With the Based upon tatLIc back-to-back testing, the CF6-50 engine with a Short Core Nozzle produces almost identlcal community noise levels as the CF6-50 engine with a Core Reverser Nozzle when operated at identical corrected thrust and flight conditions. Results presented within this report document that the dominant noise components of the CF6-50 engine (fan, LPT, and core jet) were not significantly impacted by the Short Core Nozzle.

Small changes in the directivity pattern of the third stage LPT fundamen tal BPF tone were observed for the Short Core Nozzle when compared to the Core Reverser Nozzle. The acoustic treatment suppression performance of the Short Core Nozzle may have produced this change but no definitive conclusions could be drawn as a result of the LPT third stage tone modulation and scattering by engine exhaust turbulence. It is predicted that this result will not affect aircraft flyover noise characteristics with the possible exception of aircraft operating at very low power approach flight conditions.

Even at this condi tion, however, the overall system EPNL should not increase.

7.0 ENDURANCE TEST 7 1 TEST CONFIGURATION Endurance testing of the Short Core Nozzle was conducted on a CF6-50C configuration engine. Standard flight-type fan reverser doors and prototype core cowling were installed on the engine The Short Core Nozzle outer cowl and centerbody shown in Figures 64 through 66 were installed.

7.2 TEST FACILITY In endurance testing of engine cowling, it has proven more representative to perform these tests in an outdoor test site rather than an enclosed test cell to avoid pressure/noise perturbations from the cell walls. For reasons of site availability, it was decided to utilize the Edwards Flight Test Center outdoor test facility for endurance testing. The test site used was General Electric's Test Site MB4 which is shown in Figure 67.

7.3 TEST PROCEDURE Following the mechanical checkout, break-in run, and a short engine performance analysis test, the engine was placed on "C" cycle endurance A "C" cycle is a 15 minute cycle that simulates the transient movements made during a typical airline flight Figure 68 gives a graphical presen tation of the cycle.

The mission mix of exhaust gas temperatures was selected to simulate the distribution of takeoff temperatures seen in air line service. If the required exhaust gas temperature could not be reached without exceeding fan speed limits, the engine was shut down and bleed pipes were installed The mixture of cycle was.

Takeoff EGT, 0 C (0 F) Number of Cycles Below 878 (1613) 879 (1614) 914 (1678) 350 915 (1679) - 942 (1728) 943 (1729) - 950 (1742) 951 (1743) - 960 (1760) Total 1000 7.4 TEST RESULTS A detailed visual inspection upon completion of endurance testing of the Short Core Nozzle indicated that'there was no sign of distress after the com pletion of the 1000 endurance cycles.

A formal dye-penetrant inspection cor roborated these findings.

Figure 64. Disassembled Short Core Nozzle - View A.

OPst0

a

AFT CENTERBODY iFORWARD CENTERBODY ~SOUND TREATMENT Figure 65. Disassembled Short Core Nozzle - View B.

C942

AFT CENTERBODY FORWARD CENTERBODY O0TER i L. CORE COWL SUPPORT V• WEAR STRIPS Figure 66. Disassembled Short Core Nozzle - View C.

A I

Figure 67. CF6-50 Engine at General Electric Test Site MB4, Edwards AFB.

The visual inspection was made with suspect areas being examined at loX with the following results or visual damage. Two nutplates "frozen" * Aft Centerbody - No dents and not free to slide in their holders * Forward Centerbody - No distress evident including weld line of skin been properly to forward flange. The radial access holes had not positioned at three locations and had to be elongated for wrench hole elogation was required.

clearance at the factory. Up to 30% * Exhaust Nozzle - No distress evident including the circumferential of rewelded weld of the liner skin to the forward flange Evidence joint with minor undercutting and overlapping due to areas in this fitup initial * Outer Cowl - An indentation approximately 25 mm (1 in.) in diameter location, aft and 5 mm (3/16 in ) deep is visible at the 9 o'clock looking forward, approximately 200 mm (8 in.) aft of the front cowl flange which appears to be handling damage An indication of door wear on outer cowl wear pads is evident. Leading edge of wear pad at 6:30 position, aft looking forward is well polished locally worn halfway into the forward retaining countersunk rivet and carbide) is in place but diameter Wear coating material (tungsten locally worn. No other distress was discernible to the "U" channel corner radii * Channel - Particular attention was paid distress.

s-eeFigure 6) No evidence of cracking or other The unit was then inspected to the requirements of the quality control of the instructions which had been previously prepared for the inspection for flight test and Short Core Nozzle prior to shipment to Airbus Industrie after completion of a short instrumented endurance test at Peebles The results of the zyglo inspection on the endurance test nozzle revealed no zyglo discrepancies.

8.0 ECONOMIC ASSESSMENT The Short Cord Nozzle concept was evaluated by Boeing and Douglas under Task I of this program (Reference 1). Boeing studied the concept for the 747-200 aircraft for 1% ceuise sfc improvement and Douglas evaluated the concept for 2% sfc improvement, 1% for internal performance improvement and 1% for reduced interference drag.

The engine ground test demonstrated a gross thrust coefficient improve ment of approximately 0.3% which is equivalent to a cruise sfc improvement of 0.9% Preliminary assessments of flight testing conducted by Airbus Industrie on the A300B airplane and by Douglas on the DC-l0-30 aircraft outside this program support the sfc improvement The 0.9% reduction in cruise sfc due to the internal thrust coefficient improvement results in the block fuel savings shown in Table VII for the minimum fuel consumption mission This is based, on the data presented in Reference I Table VII. Short Core Nozzle Block Fuel Savings for Internal Thrust Coefficient Improvement.

(Minimum Fuel Analysis) Aircraft Range AtFuel km kg % DC-10-30 805 -58 -0.5 2735 -211 -0.8 -599 -1 0 B-747-200 770 -37 -0.4 3460 -49 -0.1 6195 -98 -0.1 For the Boeing 747-200 aircraft, a block fuel savings of 0 4% was pro jected for the 770 km flight and 0 1% for the longer flights. The benefit in reduced nacelle weight and improved internal performance is accounted for along with the increased external nacelle drag on block fuel savings.

The effect of the increased external nacelle drag has a greater impact on the long-range flights than on the short-range flights Thus, a smaller savings is shown for the longer flights The estimated annual fuel savings per aircraft for the above block fuel savings are shown in Table VIII.

Table VIII Estimated Annual Fuel Savings Per Aircraft for Internal Thrust Coefficient Improvement (Minimum Fuel Analysis) Range Fuel Aircraft km i/AC/Year DC-10-30 805 145,900 2735 258,500 6275 526,900 109,900 B-747-200 770 42,400 6195 49,900 The economic assessment for the medium fuel price of 14 5j/1 (55U/gal) for the DC-10-30 and 11 89J/1 (451/gal) for the B-747-200 is summarized in IX for a 0.9% sfc reduction Table Table IX. Economic Assessment of Short Core Exhaust Concept for 0 9% SFC Reduction Due to Internal Thrust Coefficient Improvement (Medium Range, Medium Fuel Price, Minimum Fuel Analysis) Payback ROI Aircraft (Years) (%) DC-10-30 0 02 4106 B-747-200 13 1 2 9 0 SUMMARY OF RESULTS The Short Core Nozzle has been evaluated in three full scale engine ground tests. The main results of these tests are summarized below Performance Test The CF6-50 engine back-to-back static performance test verified within data accuracy the scale model test results and indicated a thrust coefficient improvement of approximately 0.3% This improvement results in a cruise sfc reduction of 0.9% at M 0 85, 10,668 m (35,000 ft) altitude for a thrust level of 37,800 N (8,500 ib).

At equal effective exhaust nozzle area, the Short Core Nozzle showed improvements in gross thrust at engine pressure ratio and fan speed over the Long Fixed Core Nozzle. Therefore, the Short Core Nozzle does not require a power management change to meet minimum thrust at fan speed Acoustic Test The CF6-50 engine back-to-back static acoustic test demonstrated that the Short Core Nozzle produces almost identical community noise levels as the Core Reverser Nozzle at identical thrust levels.

Dominant noise components of the CF6-50 engine, such as fan, low pressure turbine, and core jet were not signficantly impacted by the Short Core Nozzle.

Endurance Test The CF6-50 engine static endurance test demonstrated the life capability of the Short Core Nozzle hardware in 1000 flight cycles without any indica distress.

of tion Flight tests conducted outside the program indicate that a cruise sfc reduction of at least 0 9% is attainable on the Airbus Industrie A300B and the DC-10-30 aircraft with the installation of the Short Core Nozzle. The estimated annual fuel savings per aircraft for the DC-10-30, assuming only a 0.9% improvement, amounts to 146,000 to 527,000 liters, depending on range.

APPENDIX A ASSURANCE QUALITY INTRODUCTION The quality program applied to this contract is a documented system the design, manufacture, repair, overhaul, and modification throughout for gas turbine aircraft engines. The quality system has been con cycle military specifications MIL-Q-9858A, MIL-I-45208, structed to comply with Aviation Regulations FAR-145 and applicable and MIL-C-45662 and Federal portions of FAR-21.

and its implementation are defined by a complete The quality system which set of procedures which has been coordinated with the DOD and FAA and concurrence In addition, the quality system as described in has their the quality program for this contract is consistent with the requirements by NASA-Lewis Research Center. The following is a brief synopsis of established the requirements established by this system.

QUALITY SYSTEMS system is documented by operating procedures which coordinate The quality Manu the quality-related activities in the functional areas of Engineering, Programs The quality system facturing- Materials, Purchasing, and Engine all product lines are controlled by is a single-standard system wherein The actions and activities associated with deter the common quality system.

are recorded, and documentation is available for review.

mination of quality Inherent in the system is the assurance of conformance to the quality requirements This includes material verification and the performance of inspections and tests. In addition, the system provides change required changes are incorporated into control requirements which assure that design and quality documentation, and into the products.

manufacturing, procurement Measuring devices used for product acceptance and instrumentation used to results of readings during inspection control, record, monitor, or indicate periodically are reveri and test are initially inspected and calibrated and at a prescribed frequency. Such calrbration is per fied or recalibrated by technicians against standards which are traceable to the National formed Bureau of Standards. The gages are identified by a control number and are on calibration. The calibration func a recall schedule for reverification and of the location of each gage and the date it requires tion maintains a record recalibration. Instructions implement the provisions of MIL-C-45662 and the appropriate FAR requirements.

IRECEDING PAGE BL M L 103 Work sent to outside vendors is subject to quality plans which provide for control and appraisal to assure conformance to the technical requirements.

Purchase orders issued to vendors contain a, technical description of the work to be performed and instructions relative to quality requirements Engine parts are inspected to documented quality plans which define the characteristics to be inspected, the gages and tools to be used, the condi tions under which the inspection is to be performed, the sampling plan, laboratory and special process testing, and the identification and record requirements Work instructions are issued for compliance by operators, inspectors, testers, and mechanics Component part manufacture provides for laboratory overview of all special and critical processes, including qualification and certification of personnel, equipment and processes.

When work is performed in accordance with work instructions, the opera tor/inspector records that the work has been performed. This is accom plished by the operator/inspector stamping or signing the operation sequence sheet to signify that the operation has been performed.

Various designs of stamps are used to indicate the inspection status of work in-process and finished items. Performance or acceptance of special processes is indicated by distinctive stamps assigned specifically to per sonnel performing the process or inspection Administration of the stamp system and the issuance of stamps are functions of the Quality Operation The stamps are applied to the paperwork identifying or denoting the items requiring control. When stamping of hardware occurs, only laboratory approved ink is used to assure against damage.

The type and location of other part marking are specified by the design engineer on the drawing to assure effects do not compromise design require ments and part quality.

Control of part handling, storage, and delivery is maintained through the entire cycle Engines and assemblies are stored in special dollies and trans portation carts. Finished assembled parts are stored so as to preclude damage and contamination, openings are covered, lines capped and protective covers applied as required.

Nonconforming hardware is controlled by a system of material review at the component source Both a Quality representative and an Engineering representative provide the accept (use-as-is or repair) decLsions Nonconfor mances are documented, including the disposition and, corrective action if applicable to prevent recurrence The system provides for storage, retention for specified periods, and re trieval of nonconformance documentation Documentation for components is filed in the area where the component is manufactured/inspected VA._Jf, 11 MI>.;J A buildup record and test log are maintained for the assembly, inspec tion and test of each manor component or engine. Component and engine testing is performed according to documented test instructions, test plans and instru mentation plans. Test and instrumentation plans are submitted to NASA for approval to the testing.

Records essential to the economical and effective operation of the qual ity program are maintained, reviewed, and used as a basis for action. These records include inspection and test results, nonconforming material findings, laboratory analysis, and receiving inspection.

Maintainability, reliability, and safety are items considered in the basic design concept and are covered in Section 4.0.

APPENDIX B LIST OF SYMBOLS BPF Blade Passing Frequency, kHz CF Flow Coefficient CF1 5 Fan Nozzle Flow Coefficient CT Gross Thrust Coefficient DMS Data Management System EGT Exhaust Gas Temperature, 0 C C F) EPNL Effective Perceived Noise Level, EPNdB FGM Measured Gross Thrust, N (ib) FiCore Ideal Thrust of Core Nozzle, N (lb) FiFan Ideal Thrust of Fan Nozzle, N (ib) Fn, FN Net Thrust, N (b) Hum Humidity, Grains LPT Low Pressure Turbine M Mach Number N Number of Samples in Data Grouping N Fan Speed (rpm) I NIK Corrected Fan Speed (rpm) N2 Core Speed (rpm) ) N/cm (lb/in Pressure, Barometric PBar PEDS Portable Environmental Data Station PNL Perceived Noise Level, PNdB PNLT Tone Corrected Perceived Noise Level, PNdB 2 2 (lb/in ) Ambient Pressure, N/cm PO 2 2 PS Static Pressure, N/cm (lb/in ) ) N/cm (lb/in Pressure, Nozzle Total Fan Fan PT PWL Sound Power Level dB re 10-13 Watts RMS Root Mean Square sfc Specific Fuel Consumption, kg/hr N (ibm/hr lbf) - 5 Sound Pressure Level dB re 2 0 x 10 N/m SPL T Total Temperature, 0 C, (0 F) To Ambient Temperature * C ( F) t(v) "t" Statistic for (V) Degrees of Freedom, 95th Percentile V Velocity, m/sec (ft/sec) o Polar Angle Referenced to Engine Centerline, Clockwise from Inlet o Unbiased Estimate of Sample Standard Deviation (y Unbiased Estimate of Sample Variance 62 Standard Pressure Correction, PT2/10 133 (PT2/1 696) V Degrees of Freedom of Sample REFERENCES I Fasching, W.A., "CF6 Jet Eng-ine Performance Improvement Program, Task I - Feasibility Analysis", NASA Report CR-159450.

2. "Aerospace Recommended Practice ARP866A", Society of Automotive Engi 15, 1979 Revised March neering 3. Kazin, S.B and Matta, R K., "Turbine Noise Generation, Reduction and Pre diction", AIAA Paper No.

75-449, March 1975.

COMPANY/PERSONNEL Nielsen Engineering & Research 510 Clyde Avenue Mountain View, CA 94043 Dr. William Roberts TRW Equipment, TRW Inc.

23555 Euclid Avenue Cleveland, OH 44117 C. S. Kortovich Richard A. Whitaker Electronic Supply division ESB/OCNG MS 34 Hanscom AFB, MA 01731 Major Robert P. Couch Computer Avionics Corporation 120 Charcot Avenue San Jose, CA 95131 Milton E. Gregory General Electric Company AFPRO (Det.

28) Cincinnati, OH 45215 R. Glindmeyer/Mail Drop N-i Chief, Plans Requirements & Eval. Branch AUG 08 197.

NASA/LeRC Approved Distribution List COMPANY/PERSONNEL COMPANYIPERSONNEL National Aeronautics & Space Robert E. Jones/60-6 Administration Washington, DC 20546 Lawrence P. Ludwig/23-2 Dr James J. Krarner/R Harold E. Rohlik/77-2 I C(sqi!S) Colladay/RJP-2 Raymond S.

Dr.

Tita'T. Serafini/49-1 Deutsch/RT-6 C.

George Kenneth E. Skeels/500-305 William S. Aiken/RD-5 (2-capes.)

Lewis LibraryIS-3 W. Johnson/RJG-4 Harry Report Control/5-5 Nysmith/RP-4 Robert C.

NASA Langley Research Center Richard A. Rudey/RTP-6 Hampton, VA 23665 S Robert W. Leonard, Dr./158 Stafford W. Wilbur/RJP-4 Ray V.,Hood/158 L. Winblade/RHG-5 Roger NASA Ames Research Centdr NASA Lewis Research Center Moffett Field, CA 94035 21000 Brookpark Road Louis*J. Williams/237-9 Cleveland, Ohio 44135 Dr. John M. Klineberg/3-3 NASA Dryden Flight Research PO Box 273 Warner L. Stewart/3-5 Edwards, CA 93523 Dr./2089 A. Albers, James Donald L.

Nored/301-2 William 1. K0o/34820 (!copies) Ziemianski/301-4 A.

Joseph Frank V. Olinger/2093 (20copies) McAulay/301-4 E.

John NASA Scientific and Technical Edward M. Szanca/301-4 (20_coples) Information Facility PO Box Milton A. Beheim/86-1 Balt/Wash Intl Airport, MD 21240 (3Qcqqie) Department Accessioning W. HaH/49-1 Robert Federal Aviation Administration Melvin J. Hartmann/5-3 DOT/FAA/NAFEC ANA-410, Bldgl 211 Richard A. Rudey/60-4 Atlantic City, NJ 08405 Frings Gary W Schroeder/500-207 Robert Department of Transporation William J. Anderson/23-2 21000 Second St., SW Washington, DC 20591 Salvatore J. Grisaffe/49-3 Harold True/ARD 550 Marvin H. Hirschberg/49-1 William T. Westfield/ARD 500 Robert S. Zuckerman/ARD 550 COMPANY/PERSONNEL COMPANY/PERSONNEL Civil Aeronautics Board Naval Air Propulsion Center Washington, DC 20428 L 1440 Parkway Avenue J. E. Constantz/B-68 Trenton,,NJ 08628 Walter L. Pasela (PE 63) Wright-Patterson Air Force Base Dayton, OH 45433 Naval Weapons Center Everett.E. Bailey/AFAPL/TBD Code 3271 China Lalke, CA 93555 R. C. Cochran/ASD/SDUB J. A. O'Nalley Arnold Engineering and Development Center R. M. Cox/AFFDL/DA S Arnold AFS, TX 39389 Dr.

James G. Nitchell/AEDC/XRFX Lt. Col. D. S. Dickson/ASD/YZI R. Roepke/AEDC/XRFX Lt. John Edens/ASD/ENFPA Air Transport Association Lt. Col. Reynald E. Fitzsmmons/AFAPL/TBD 1709 New York Avenue, NW Washington, DC 20056 Keith R. Hamilton/AFAPL/TBC E. L. Thomas C. M.

High/ASD/YZE General Electric Company One Neumann Way Capt. Charles M. Hutcheson/ASD/YZET Evendale, OH 45215 Qg_ggie) Al Schexnayder/H4 ",t. C. L. Klinger/ASD/YZET Ray Wulf/F117 Lt. Col. James L.

Pettigrew/ASD/YZEA General Electric Company Perry Shellaberger/ASD/ENFPA 5300 Riverside Drive Cleveland, CH 44135 E. C. Simpson/AFAPL/TB Meade Rudasill Le.-E; Whonic/ASD/YZN United Technologies Corp.

Pratt & Whitney Aircraft Offutt Air Force Base 400 Main Street Headquarters East Hartford, CT 05108 Omaha, NE 68113 William 0. Gaffin (L9_EoPies) Col. J. Streett/SAC/LGME G. Phillip Sallee Oklahoma City Air Logistics Center Tinker AFB, OK 73145 United Technologies Corp.

Capt. P. Davis/OC-ALC/MM Pratt & Whitney Aircraft 20800 Center Ridge Road, Rm. 105 Capt. Steven Erickson/OC-ALC/MA USAF Rocky River, OH 44116 George C. Falkenstein E. Reynolds, Engine Test Branch (MAET) COMPANY/PERSONNEL COMPANY/PERSONNEL United Technologies Corp. Braniff International Hamilton Standard Division Braniff Tower Bradley Field PO Box 35001 Windsor Locks, CN 06096 Dallas, TX 75235 K. Liebing Hank Nelson Louis Urban/MS3-2-36 Continental Air Lines, Inc.

Los Angeles Inter. Airport The Boeing Company Los Angeles, CA 90009 PO Box 3707 Frank Forster Seattle, WA 98124 Don Nordstrom Delta Air Lines, Inc.

Hartsfield-Atlanta Int'l Airport William B. Anderson Atlanta, GA 30320 James Goodrum Kenneth H. Dickenson 3N-33 Eastern Air Lines, Inc.

Paul G. Kafka Miami International Airport Miami, FL 33148 Richard L.

Martin MS 73-07 2 copies M. Dow, Bldg. 21 John L. White Arthur Fishbein, Bldg. 21 McDonnell Douglas P. M. Johnstone 3855 Lakewood Blvd.

Long Beach, CA 90846 The Flying Tiger Line, Inc.

Ronald-Kawai MC 36-41 7401 World Way West Los Angeles Inter. Airport F. L. Junkermann 36-41 Los Angeles, CA 90009 James 14. Dimin 35-31 Klotsche Max Bruno Lewandowski Technical Library 36-84 J. R. Thurman Lockheed-California Co.

PO Box 551 National Airlines, Inc.

Burbank, CA 91520 PO Box 592055 John L. Benson Airport Mail Facility Miami, FL 33159 Tom Laughlin, Jr. R. A. Starner Allegheny Airlines, Inc. Northwest Airlines, Inc.

Greater Pittsburgh Int'l Airport Minn.-St. Paul Int'l Airport Pittsburgh, PA 15231 , St. Paul, MN 55111 William G. Peppler Al Radosta American Airlines, Inc.

Pan American World Airways N. Mingo Road JFK International Airport Tulsa, OK 74151 Jamaica, NY 10430 Keith Grayson John G. Borger Bob B. Cooper Lewis H. Allen Andersen Niels B.

Ray G. Fenner Robert E. Clinton, Jr.

Anqus MacLarty COMPANY/PERSONNEL COMPANY/PERSONNEL Aerojet Manufacturing Company Piedmont Airlines - Engineering Smith Reynolds Airport Vice President Winston-Salem, NC 27102 601 S. Placentia Fullerton, CA 92634 H. M. Cartwright John Kortenhoeven Paul M. Rehder Air Research Manufacturing Company Inc. 402 South 36th Street.

Seaboard World Airlines, Seaboard World Bldg. PO Box 5217 Phoenix, AZ 85010 John F. Kennedy Int'l Airport Karl R. Fledderjohn Jamaica, NY 11430 Ralph J. Barba Dept. 93-200/503-3S Dr. M. Steele Jere T. Farrah Dept. 93-010/503-4B Trans World Airlines Kansas City Inter. Airport F. Weber PO Box 20126 Dept. 93-2001503-3S Kansas City, MO 64195 MCI AVCO Lycoming Division Ken Izumikawa 2-280 550 South Main Street CN 06497 D. L. Kruse 2-280 MCI Stratford, A. Bright Walter D. Sherwood W. L. Christensen United Airlines, Inc.

Sn Francisco Inter. Airport General Motors Corporation Detroit Diesel Allison Division San Francisco, CA 94128 Box 894 John Curry PO IN 46206 Indianapolis, P. Hardy R. A. Sulkoske MS V19 G. A. Williams MS T8 James Uhl Westarn Air Lines, Inc. The Aerospace Corporation 6060 Avion Dr. Box 92,005 PO Box 92957 Postal Center Los Angeles, CA 90009 World Way CA 90009 Ronald R. Covey Los Angeles, Wal tgktHol tz The Aerospace Corporation 2350 East El Segundo Blvd.

Co6per Airmotive, Inc.

El Segundo, CA 90245 4312 Putman Street Dallas, TX 75235 W. Roessler B. Carter Advanced Technology, Inc.

7923 Jones Branch Drive Terry Harrison McLean, VA 22101 Pacific Airmotive Corporation Bernard C. Doyle, Jr.

2940 N. Hollywood Way Burbank, CA 91503 Delco Electronics Avionics Sales Office Oddvar 0. Bendikson Avenue 7929 S. Howell 53207 113 Joseph R. Gast Milwaukee, WI J. Sheldrick