Effects of nacelle shape on drag and weight of a supersonic cruising aircraft

NASA CR-144893 EFFECTS OF NACELLE SHAPE ON DRAG AND WEIGHT OF A SUPERSONIC CRUISING AIRCRAFT Prepared by: ELLWOOD BONNER RONALD Y. MAIRS RAY M. TYSON October 1975 PREPARED UNDER CONTRACT NAS 1-13906 BY LOS ANGELES AIRCRAFT DIVISION ROCKWELL INTERNATIONAL CORPORATION LOS ANGELES, CALIFORNIA FOR LANGLEY RESEARCH CENTER NATIONAL AERONAUTICS AND SPACE ADMINISTRATION U.S. GOVERNMENT AGENCIES AND THEIR CONTRACTORS ONLY FOREWORD This document presents results of the study "Parametric Analysis of the Effects of Nacelle Shape on Drag and Weight of Supersonic Cruising Aircraft."

The NASA Technical Representative was Russell B. Sorrells, III. In addition to the authors noted, significant contributions to this study and report were made by Bruce E. Moore, aerodynamics; David Chaloff, mass properties; Henry K. Chin and Louis C. Young, propulsion; and Lester D. Hendrix, configuration design.

TABLE OF CONTENTS Page SUMMARY ' 1 INTRODUCTION 1 LIST OF SYMBOLS 2 STUDY PROCEDURE 3 Approach 3 Baseline Airplane Definition 4 Basepoint Airplane 4 Design Mission 8 Alternate Mission 10 FAR 36 Takeoff 10 Balanced Field Takeoff 10 Thrust-to-Drag Ratio 12 Baseline Airplane 12 Propulsion 16 Mass Properties 25 Aerodynamics 27 Parametric Drag Analysis 34 Weight Sensitivity Analysis 85 Incremental Nacelle Drag 88 Propulsion System Weight Trades 88 Engine Specific Fuel Consumption Trades 88 Takeoff Thrust 88 DISCUSSION OF RESULTS . 88 Nacelle Drag 88 Weight .Sensitivity 94 RECOMMENDATIONS . 98 REFERENCES 98 ill LIST OF ILLUSTRATIONS Figure Title Page 1 Baseline Airplane 5 2 Vehicle Sizing and Performance Evaluation 7 Program 3 Reference Mission 9 4 Alternate Mission 11 5 Balanced Field Length Definition 12 6 Gross Weight Versus Thrust-to-Weight $ Wing Loading 13 7 Balanced Field Length Versus Thrust-to-Weight 14 and Wing Loading 8 Thrust/Drag Ratio Versus Thrust-to-Weight £ Wing 15 Loading 9 Baseline Airplane Climb Path 22 10 Mach 2.4 Inlet Configuration 23 11 Basepoint Nacelle 24 12 Comparison of Reference and Basepoint Nacelles 29 13 Nacelle Cross- Area Variation 30 14 Basepoint Cross- Area Variation 31 15 Basepoint Incremental Nacelle Drag Versus Mach 33 Number 16 Basepoint Wave Drag Sizing Data 35 17 C Versus Mach Number 36 T L K 18 c Versus Mach Number 37 n TC 19 «K' Factor Versus Mach Number 38 20 Baseline Cross-Sectional Area Variation 39 21 Effect of Nacelle Size on Nacelle Location 42 22 Nacelle Parametric Cross-Sectional Area Extremes 46 23 Nominal Baseline Nacelle Shape Perturbation 47 Extremes, £/dc = 6 24 Effect of Geometry Perturbations on Nacelle 49 Incremental Wave Drag at M = 1.2, AC = 1.86 Q sq.m. (20 sq.ft.), A/d = 5.5, A /A = 1.25 c MAX C 25 Effect of Geometry Perturbations on Nacelle 50 Incremental Wave Drag at M= 1.2, A = 1.86 sq.m. (20 sq.ft.), A/d = 5.5, A^/A =1.5 IV LIST OF ILLUSTRATIONS (Continued) Figure Page 26 Effect of Geometry Perturbations on Nacelle • 51 Incremental Wave Drag atM -1.2, A =1.86 sq.m. (20 sq.ft.), Vd = 5.5, A7A = 2.0 27 Effect of Geometry Perturbations on Nacelle 52 Incremental Wave Drag at MQ = 1.2, A =1.86 sq.m. (20 sq.ft.), 4/d = 7.0, Aft = 1.25 28 Effect of Geometry Perturbations on Nacelle 53 Incremental Wave Drag at M = 1.2, A = 1.86 c sq.m. (20 sq.ft.), 4/d = °7.0, A ^ f y . = 1.5 c 29 Effect of Geometry Perturbations on Nacelle 54 Incremental Wave Drag atM =1.2, A = 1;86 sq.m. (20 sq.ft.), J2/d = °7.0, A^fy, = 2.0 c 30 Effect of Geometry Perturbations on Nacelle 55 Incremental Wave Drag at M = 2.32, A =1.86 sq.m. (20 sq.ft.), j8/d = §.5, A/A = 1.25 31 Effect of Geometry Perturbations on Nacelle 56 Incremental Wave Drag at M = 2 . 32 , A = 1 . 86 sq.m. (20 sq.ft.), jg/d = °5.5, A^/A^ = 1.5 c 32 Effect of Geometry Perturbations on Nacelle 57 Incremental Wave Drag at M = 2.32, A =1.86 sq.m. (20 sq.ft.), j&/d = °5.5, A^/A^ = 2.0 c 33 Effect of Geometry Perturbations on Nacelle 58 Incremental Wave Drag at M = 2.32, A =1.86 sq.m. (20 sq.ft.), _0/d = °7.0, A/A = 1.25 34 Effect of Geometry Perturbations on Nacelle 59 Incremental Wave Drag at M = 2.32, A =1.86 sq.m. (20 sq.ft.), i/d = °7.0, A^/A^ = 1.5 c 35 Effect of Geometry Perturbations on Nacelle 60 Incremental Wave Drag at M = 2.32, A =1.86 .sq.m. (20 sq.ft.), j2/d = °7.0, A^/A^ = 2.0 c 36 Effect of Geometry Perturbations on Nacelle 61 Incremental Wave Drag atM =1.2, A =2.79 sq.m. (30 sq.ft.), j?/d = °5.5, A ^ f y . = 1.25 c v LIST OF ILLUSTRATIONS (Continued) Figure Page 37 Effect of Geometry Perturbations on Nacelle 62 Incremental Wave Drag atM =1.2, A = 2. 79 sq.m. (30 sq.ft.), jfl/d, = °5.5, A^ft^ = 1.5 38 Effect of Geometry Perturbations on Nacelle 63 Incremental Wave Drag at M = 1.2, A = 2.79 sq.m. (30 sq.ft.), j0/d = °5.5, A^/A = 2.0 c 39 Effect of Geometry Perturbations on Nacelle 64 Incremental Wave Drag atM =1.2, A =2.79 sq.m. (30 sq.ft.), j2/d = ° 7 . 0 , A ^ f y . = 1.25 c 40 Effect of Geometry Perturbations on Nacelle 65 Incremental Wave Drag at M = 1.2, A = 2.79 sq.m. (30 sq.ft.), jg/d = 9.0, A/ = 1.5 41 Effect of Geometry Perturbations on Nacelle 66 Incremental Wave Drag at M = 1 . 2 , A = 2 . 79 sq.m. (30 sq.ft.), A/d = 9.0, \ ^ / \ = 2.0 c 42 Effect of Geometry Perturbations on Nacelle 67 Incremental Wave Drag at M = 2.32, A = 2.79 sq.m. (30 sq.ft.), j2/d = °5.5, A^/A^ = 1.25 c 43 Effect of Geometry Perturbations on Nacelle 68 Incremental Wave Drag at M = 2.32, A = 2.79 sq.m. (30 sq.ft.), j2/d = §.5, A ^ / A j : = 1.5 c 44 Effect of Geometry Perturbations on Nacelle 69 Incremental Wave Drag at M = 2.32, A =2.79 sq.m. (30 sq.ft.), l/d = °5.5, A / A . = 2.0 45 Effect of Geometry Perturbations on Nacelle 70 Incremental Wave Drag at M = 2.32, A = 2.79 sq.m. (30 sq.ft.), j2/d = °7.0, A^/X. = 1.25 c 46 Effect of Geometry Perturbations on Nacelle 71 Incremental Wave Drag at M = 2.32, A = 2.79 sq.m. (30 sq.ft.) , j2/d = °7.0, A^/X^ = 1.5 c 47 Effect of Geometry Perturbations on Nacelle 72 Incremental Wave Drag at M = 2.32, A = 2.79 sq.m. (30 sq.ft.), j2/d = °7.0, A / A , = 2.0 VI LIST OF ILLUSTRATIONS (Continued) Figure Page 48 Effect of Geometry Perturbations on Nacelle 73 Incremental Wave Drag atM =1.2, A =3.72 sq.m. (40 sq.ft.), J>/d = °5.5, A^fy. = 1.25 c 49 Effect of Geometry Perturbations on Nacelle 74 Incremental Wave Drag atM =1.2, A =3.72 sq.m. (40 sq.ft.), jg/d = °5.5, A^fy. = 1.5 c 50 Effect of Geometry Perturbations on Nacelle 75 Incremental Wave Drag atM =1.2, A =3.72 sq.m. (40 sq.ft.), j2/d = °5.5, A. = 2.0 51 Effect of Geometry Perturbations on Nacelle 76 Incremental Wave Drag atM =1.2, A =3.72 sq.m. (40 sq.ft.), j2/d = °7.0, A^fy. = 1.25 c 52 Effect of Geometry Perturbations on Nacelle 77 Incremental Wave Drag at M = 1.2, A =3.72 sq.m. (40 sq.ft.), j2/d = °7.0, A^fy. = 1.5 c 53 Effect of Geometry Perturbations on Nacelle 78 Incremental Wave Drag atM =1.2, A =3.72 sq.m. (40 sq.ft.), j2/d = 7.0, A^/X^ = 2.0 c 54 Effect of Geometry Perturbations on Nacelle 79 Incremental Wave Drag at M = 2.32, A =3.72 sq.m. (40 sq.ft.), 4/d = §.5, A^^/K^ = 1.25 c 55 Effect of Geometry Perturbations on Nacelle 80 Incremental Wave Drag at M = 2.32, A = 3.72 sq.m. (40 sq.ft.), jg/d = °5.5, A^/A^ = 1.5 c 56 Effect of Geometry Perturbations on Nacelle 81 Incremental Wave Drag at M = 2.32, A =3.72 sq.m. (40 sq.ft.), j2/d = °5.5, A/A = 2.0 57 Effect of Geometry Perturbations on Nacelle 82 Incremental Wave Drag at M = 2.32, A = 3.72 sq.m. (40 sq.ft.), j2/d = 7.0, A.. /A = 1.25 AY C JVI/VA. C 58 Effect of Geometry Perturbations on Nacelle 83 Incremental Wave Drag at M = 2.32, A = 3.72 sq.m. (40 sq.ft.), l/d = 7.0, A^/A^ -1.5 c VI1 LIST OF ILLUSTRATIONS (Concluded) Figure Page 59 Effect of Geometry Perturbations on Nacelle . 84 Incremental Wave Drag at M = 2.32, A =3.72 sq.m. (40 sq.ft.), j2/d = °7.0, A^/A^ = 2.0 c 60 Typical Nacelle Incremental Wave Drag Variation 86 With Nacelle Size at M 2.32 o 61 Typical Nacelle Incremental Wave Drag Variations 87 With Mach Number 62 Nacelle Drag Sensitivity Trade 89 63 Typical Incremental Nacelle Drag Variation 90 64 Propulsion Weight and SFC Sensitivity Trades 91 65 Takeoff Thrust Sensitivity Trade 92 66 Typical Nacelle Incremental Drag Variation 93 With Boattail Area at Mach 2.32 67 Typical Nacelle Incremental Drag Variation 95 Versus Axial Position of Maximum Area at Mach 2.32 68 Typical Nacelle Installation Drag Variation 97 With Nacelle Size at Mach 2.32 Vlll LIST OF TABLES Table Title Page I Airplane Characteristics 17 II Baseline Design Mission Summary-International Units 18 III Baseline Design Mission Summary-English Units 19 IV Baseline Alternate Mission Summary-International Units 20 V Baseline Alternate Mission Summary-English Units 21 VI Vehicle Weight Summary 26 VII Basepoint Engine Weight 27 VIII Basepoint Nacelle Weight 27 IX Basepoint Configuration Estimated Skin Friction and 28 Wave Drag Characteristics X Baseline Configuration Surface Area and Length 41 Summary XI Baseline Configuration Estimated Skin Friction 41 and Wave Drag Characteristics XII Nacelle Parameter Values 44 XIII Summary of Parametric Friction Drag Results 45 XIV Summary of Parametric Wave Drag Results 48 XV Sensitivity Comparison 90 IX EFFECTS OF NACELLE SHAPE ON DRAG AND WEIGHT OF A SUPERSONIC CRUISING AIRCRAFT By Ellwood Bonner, Ronald Y. Mairs, and Ray M. Tyson Los Angeles Aircraft Division, Rockwell International SUMMARY The objectives of this study were to develop the quantitive relation- ship of cruise drag and nacelle shape for a representative advanced super- sonic transport configuration and to provide system sensitivity data which could be used to assess the overall value of propulsion system variations.

The NASA arrow-wing configuration was used as a baseline airframe. Only those changes that were necessary to install the Pratt and Whitney Aircraft VSCE 502B engine were made to the aircraft.

Nacelle shape parameters were systematically varied, and the effects of these variations on wave and friction drag were determined. The effects of changes in vehicle drag, propulsion weight, and specific fuel consumption on vehicle takeoff gross weight were computed.

Generally, it was found that nacelles shaped such that the maximum cross-sectional area occurred at or near the nozzle exit resulted in the lowest wave drag. In fact, nacelle shapes were found that produce favorable interference effects (drag reduction) of such magnitude as to nearly offset the friction drag of the nacelle. It should be emphasized that these results are valid only for vehicles of this general configuration and nacelle loca- tion. Different vehicle configurations or nacelle locations could result in different "best" shapes.

It is recommended that the drag and sensitivity data generated in this program be used in the analysis of future propulsion system trade studies.

Further, it is recommended that vectorable, two-dimensional nozzles be studied, and that canting downward of non-vectorable, axisymmetric nozzles be examined.

INTRODUCTION The National Aeronautics and Space Administration is conducting a con- tinuing program of advanced supersonic technology studies with the objective of developing an adequate technology base to support development of future supersonic cruising aircraft. It is recognized in this program that one of the more sensitive problems in the synthesis of a successful supersonic cruising aircraft is that of airframe/engine integration. This process must investigate and properly manage the interactions between the technical disci- plines of external aerodynamics, internal aerodynamics, engine cycle design, acoustics, mass properties, and structural design; and, it must be responsive to the practical considerations of fabrication, maintenance, and operation.

The results of a recently-completed study, reference 1, of the effects of nacelle size and nacelle shape on the drag, weight, and wing camber plane warping of a supersonic transport illustrated the sensitivity of these para- meters to relatively small changes in nacelle shape. The resultant shape of a nacelle is dependent on the geometry of the engine (inlet area, mounting provisions, accessory location, nozzle area, etc.) since this establishes certain control points in the design of the nacelle. It is important, therefore, that the engine designer be aware of this sensitivity to engine geometry, and be provided with some guidelines for favorable geometry rela- tionships. It is probable that some engine geometry control can be achieved by the designer with no penalty in engine performance, although on a total system basis some engine performance degradation could be accepted in trade for reduced drag.

Although considerable effort has been expended on the problem of air- frame/engine integration, it has been mostly in the nature of point designs.

The referenced study produced results for two specific nacelle shapes which resulted from installation of a dry turbojet engine and a duct heating turbo- fan engine. A comparison of these results shows the superiority of one over the other, but gives no information directly applicable to other installa- tions having differing nacelle shapes. Because of the economy that can be affected in configuration synthesis by having good engine geometry character- istics for use in the first iteration of the engine/airframe integration process, an effort has been made to supply the engine designer with guide- lines for favorable geometry that he can apply early in the engine develop- mental process. Parametric data have been generated on the effects of variations of nacelle shape on cruise drag for a range of shapes that reasonably cover engine designs applicable to supersonic cruising aircraft.

In considering possible trades of reduced drag through design changes in the engine envelope for some penalty in engine weight and specific fuel consumption (SFC), it is necessary to have visibility of the net impact of all three effects on the total airplane system in order to make a compara- tive evaluation. Therefore, sensitivity data were developed for the effects of changes in drag, propulsion system weight, takeoff thrust, and SFC on the takeoff gross weight as a figure of merit.

The program was of seven months duration, including review and sub- mittal of the final report, and .was organized around four tasks. These were: 1 - Baseline Airplane Definition; 2 - Parametric Drag Analysis; 3 - Weight Sensitivity Analyses; and 4 - Reporting.

LIST OF SYMBOLS A Area, sq. m. (sq. ft. or sq. in.)

BLB Boundary layer bleed BLC Boundary layer control B.P. Basepoint C Coefficient or Chord, m. (ft. or in.)

d Diameter, m. (ft. or in.)

D Drag, (Ib.)

dB Decibel K Drag-due-to-lift factor Si Length, m. (ft. or in.)

L Lift, n. (Ib.)

M Mach number S Area, sq. m. (sq. ft. or sq. in.)

SFC Specific fuel consumption, kg./hr./n. (lb./hr./lb.)

T Thrust, n. (Ib.)

TOGW Takeoff gross weight, kg. (Ib.)

V Velocity, m./sec. (ft./sec.)

W Weight, kg. (Ib.)

X Nacelle station, m. (ft. or in.)

A Increment Subscripts AMAX Maximum cross-sectional area c Capture D Drag F Friction i Inlet throat K Indicates lift coefficient at minimum drag L Lift LO Liftoff MAX Maximum n Nozzle exit P Maximum protrusion REF Reference R Root W Wave 0 Freestream 1 Critical engine failure STUDY PROCEDURE Approach The general approach used in this study was to: 1) define a baseline airplane and determine its performance, 2) parametrically vary nacelle shape and determine nacelle drag increments due to these variations, 3) determine vehicle takeoff gross weight sensitivity to changes in drag, weight, take- off thrust, and SFC, and 4) report results. In this report, descriptions of the airplane configurations used are as follows: 1) the reference airplane is the NASA modified SCAT 15F arrow-wing, supersonic transport (defined in reference 2) 2) the baseppint vehicle is the reference modified only as required to install the Pratt and Whitney Aircraft (P§WA) VSCE 502B engine 3) the baseline airplane is the basepoint resized to the design requirements on a standard-plus-8 C day The structure design and operational empty weight of the reference airplane were assumed to meet all design criteria. Weight and aerodynamic character- istics of the study airplanes were derived by increments from the reference configuration.

Baseline Airplane Definition Basepoint airplane.- The "basepoint" airplane for this study is based on the NASA modified SCAT 15F arrow wing reference configuration as described in reference 2. The propulsion system of this airplane has been replaced with P§WA variable stream control engines (VSCE 502B) having 408 kg./sec.

(900 Ib./sec.) airflow each and with axisymmetric variable geometry inlets designed for Mach 2.4 cruise conditions. The resulting basepoint vehicle is shown in figure 1. This airplane has a gross weight of 336 973 kg.

(742 890 lb.), a range of 7 471 km. (4 034 n.mi.), and a balanced field length of 3 017 m. (9 898 ft.).

All performance and sizing calculations were made using the Rockwell Vehicle Sizing and Performance Evaluation Program (VSPEP). This computer program is a design tool capable of scaling a known basepoint vehicle according to specified values of several different design parameters.

These include vehicle gross weight (or fuel weight), thrust-to-weight ratio (or engine size), wing-loading (or wing area), and payload or fixed equipment weight and volume. Performance may be determined at specified gross weight, or alternatively, a search routine permits automatic sizing of the vehicle gross weight such that a specified radius or range of the design mission is satisfied. Vehicle performance is calculated internally from a set of sub- routines programmed according to a detailed performance analysis model.

The subroutines are general in nature and permit calculation of a wide variety of mission profiles. Several mission profiles may be calculated simultaneously. Takeoff and landing distances and maneuvering capability may also be determined. Figure 2 illustrates the evaluation process.

Typical mission legs which may be calculated include warmup, taxi, takeoff, climb, descent, cruise, and loiter operations. Climb and descent performance are determined by numerical integration of the equations of motion along a specified flight schedule. Internally generated schedules are also available, including minimum time and minimum fuel flight paths as defined by the energy method. Constraints on the allowable flight regime are included. Cruises and loiters may be determined at fixed or optimum speeds and altitudes. Numerical searches are used to determine optimum speeds and altitudes at the beginning and end of each of these legs.

Data input to the VSPEP for the AST basepoint vehicle include: . Weights broken down by major component, along with scaling information on the wing, tails, fuselage, and engines.

• Drags broken down by major component and by type (e.g., friction drag, wave drag, drag due to lift, base drag).

• Installed propulsion data, including thrust and fuel flow as functions of speed, altitude, and power setting.

• Dimensional data such as lengths, areas, and volumes for major components and the total vehicle.

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4. Balanced field takeoff distance 5. Thrust-to-drag ratio at Mach 2.32, 18 300 m. (60 000 ft.)

6. Thrust-to-drag ratio at Mach 1.2 during the climb leg A description of each of these performance items is given in the following paragraphs. Because engine data were provided for a standard- plus-8°C (14.4°F) day, all airplane performance characteristics were computed for that atmospheric condition.

Design mission.-A profile of the design mission is shown in figure 3.

This mission consists mainly of a Mach 2.32 cruise. Fuel reserves as recommended in reference 3 are calculated for an alternate airport located 460 km. (250 n.mi.) from the destination airport.

The design mission consists of: 1. Warmup and takeoff - 10 minutes at idle power plus 1 minute at maximum power.

2. Climb - Maximum power climb and accelerate to cruise altitude and Mach number.

3. Cruise - Cruise at Mach 2.32 at altitude for best cruise range.

4. Descent - Descend and decelerate to Mach 0.5 and 457 m.

(1 500 ft.) using idle power.

5. Approach and land - Descend to Mach 0.3 at sea level using idle power.

6. Taxi - 5 minutes at idle power.

7. Reserve allowance - 5 percent of total fuel used in all previous legs.

8. Reserve climb - Climb to subsonic cruise conditions.

9. Reserve cruise - Subsonic cruise at Mach number and altitude for best range.

10. Reserve descent - Descend and decelerate to holding altitude and Mach number using idle power.

11. Reserve hold - Loiter for 30 minutes at 3 048 m. (10 000 ft.)

at the Mach number for best endurance.

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Alternate mission.-A profile of the alternate mission is shown in figure 4. The first half of the alternate is identical to the first half of the design mission. At the point corresponding to the midpoint of the design mission, a failure is assumed to occur in the most critical engine.

At this point, the airplane descends and continues to cruise subsonically with one engine windmilling. The fuel reserve remaining at the end of this mission is equal to the reserve fuel as calculated for the design mission.

The alternate mission consists of: 1. Warmup and takeoff - Same as design mission.

2. Climb - Same as design mission.

3. Cruise - Same as design mission.

4. Descent - Descend and decelerate to subsonic cruise condi- tions using idle power, following failure of most critical engine.

5. Cruise - Subsonic cruise at Mach number and altitude for best range with one engine inoperative.

6. Descend and land - Descend to sea level using idle power 7. Reserve - Allow total reserve fuel equal to that calculated for design mission legs 7 through 12 .

FAR 36 takeoff.-Takeoff distance is calculated over a 10.7 m. (35 ft.)

obstacle using takeoff thrust which has been throttled so that FAR 36 noise requirements are not exceeded. It is assumed that a maximum usable lift coefficient of 0.555 is available for climbout.

Balanced field takeoff.-The balanced field length is defined such that segments B+C = D*E for takeoff over a 10.7 m.(35 ft) obstacle as shown in figure 5. The speed at which the engine failure occurs (i.e., V^) is varied until this definition is satisfied to a reasonable tolerance. The balanced field length is then taken as the larger of the total takeoff distance or the accelerate-stop distance. For the stop distance calculation, the re- maining engines are cut to idle power.

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The thrust -to -drag ratio is also calculated for the point in the climb- accelerate leg at which Mach 1.2 is reached. In this case the altitude and vehicle weight are the actual values during the climb at which the Mach number reaches 1.2.

Baseline Airplane. -The "baseline" airplane for this study is a resized version of the above-described "basepoint." Resizing was accomplished by exercising the VSPEP for a matrix of thrust-to-weight and wing loading values, and allowing the program to search for the gross weight, in each case, that satisfies the design mission range requirement of 7 408 km.

(4 000 n.mi.). Plots of the results are shown in figures 6 through 8. The parameters shown include vehicle gross weight as well as those performance items for which requirements must be met.

Requirement lines are crossplotted onto the airplane gross weight plot in figure 6. This allows a "baseline" airplane to be chosen which is defined as the minimum gross weight vehicle that meets or exceeds the following performance requirements: Range - 7 408 km. (4 000 n.mi.)

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Figure 8 .-Thrust/drag ratio versus thrust-to-weight and wing loading Design mission range 7 408 km. (4 000 n.mi.) with 292 passengers Balanced field length 3 200 m. (10 500 ft.)

Minimum T/D during climb or 1.2 cruise Since the thrust-to-drag requirements are exceeded for all cases, only the balanced field length requirements are included in figure 6.

The resulting "baseline" airplane has a gross weight of 323 046 kg.

(712 188 lb.); a thrust-to-weight ratio of 3.16 n./kg. (0.323 Ib./lb.) based on installed, static takeoff thrust; and a wing loading of 354 kg./sq.m.

(72.5 Ib./sq. ft.) based on gross wing area.

Further airplane design and performance characteristics for both the "basepoint" and "baseline" airplanes are shown in table I. Design and alternate mission summaries are shown in tables II through V for the base- line airplane. In these tables, the first leg of the alternate mission includes the first four legs of the design mission while leg 5 of the al- ternate includes reserves for legs 9 through 15 of the design mission.

The path followed during the climb-accelerate leg is a minimum fuel path cal- culated internally by the VSPEP program. This path as calculated for the baseline airplane is shown in figure 9.

Propulsion.-Because many of the current and recently-completed supersonic cruising aircraft studies have used axisymmetric inlets, a mixed-compression, axisymmetric inlet was defined for use in the basepoint aircraft for this study. This inlet is illustrated in figure 10, and the nacelle is shown in figure 11. The inlet diameter is 1.93 meters (76 inches), and capture area is 2.926 square meters ( 4 536 square inches). The three primary modes of inlet operation are illustrated. Position 1 shows the supersonic cruise geometry. Position 2 shows the centerbody translated forward to the maximum throat area (Aj/Ac of 0.65); the centerbody is translated forward as a single unit. In Position 3, the basic centerbody is held in the transonic position, but the fore and aft conical segments are translated aft to create a center- body auxiliary inlet. The auxiliary inlet opening in the centerbody is ten percent of capture area. The station cuts on figure 10 show approximate flow passages in the centerbody. In particular, the auxiliary ducting must be divided into four passages around the boundary layer control (BLC) bleed air passages near the centerbody maximum diameter. Inlet pressure recoveries and spillage, bypass, and BLC drags were estimated, and their effects were included in installed propulsion performance.

The engine performance data available for the VSCE 502B engine included the effects of an inlet recovery schedule, nozzle external drags (base plus boattail), 0.45 kilogram-per-second (1.0 pound-per-second) high-pressure compressor air bleed, and 149 kilowatts (200 horsepower) power extraction.

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At takeoff, the engine power setting was scheduled so that the aircraft meets FAR 36-traded noise levels. The SAE exhaust jet noise prediction method was used with the modification to overall sound pressure level recommended by Bushell (reference 4). This modification has no effect on perceived noise level at static condition, but it results in approximately four decibels (dB) higher noise than the standard SAE method at Mach 0.3. A 1.5dB reduction in sideline noise has been assumed due to sideline shield- ing while the airplane is on the ground. An eight-decibel reduction in noise level due to the coannular nozzle effect has been assumed for all flight conditions and power settings. Information from P$WA indicates that coannu- lar configurations reduce noise by seven to nine decibels when the difference between core velocity and bypass velocity is 152 meters per second (500 feet per second) or more, with the core stream having the lower velocity.

Mass properties.-The basepoint vehicle weight summary is given in table V E T h e NASA reference vehicle weight summary (reference 2) from which the basepoint was derived is also shown. The differences between the weights of the two vehicles are in the engines and nacelles. The basepoint vehicle has VSCE 502B 408 kg./sec. (900 Ib./sec.) airflow engines in lieu of the 363 kg./sec. (800 Ib./sec.) engines in the NASA reference vehicle.

The VSCE 502B bare engine weight including nozzle and thrust reverser was supplied by P§WA. Weight increments of 22.7 kg. (50 Ib.) for residual fluids and 22.7 kg. (50 Ib.) for miscellaneous engine/airframe interfacing provisions were added to the bare weight to obtain an installed weight. Table VII shows the installed engine weight summary.

The basepoint nacelle weight estimate is based on the nacelle drawing, figure 11. For the weight evaluation the nacelle was divided into three sections: forward of the engine front face (inlet cowl), aft of the front face (engine cowl) and inlet spike. The engine cowl weight was estimated at 34.2 kg./sq.m. (7 Ib./sq. ft.) of wetted area. This weight includes all the nacelle structure that supports and surrounds the engine and was derived from prior Rockwell International studies of a similar type. The inlet cowl and spike weights were calculated using statistical weight estimating equa- tions obtained from the Technical Report SEG-TR-67-1, Preliminary Design Methodology for Air-Induction Systems (reference 5). Engine mount weights were calculated statistically at 1.5 percent of the engine weight. The mount weights are included with the nacelle weight.

The weight summary of the basepoint nacelle is presented in table VIII.

TABLE VI.-VEHICLE WEIGHT SUMMARY NASA REFERENCE VEHICLE BASEPOINT ITEM LB LB Kg Kg 37 805 83 347 37 805 83 347 Wing 2 391 2 391 5 271 5 271 Horizontal Tail 4 735 2 148 2 148 4 735 Vertical Tail Fuselage 54 314 24 636 54 314 24 636 13 138 13 138 Landing Gear 28 965 28 965 8 625 19 015 7 410 16 336 Nacelle (88 743 (195 647 (87 528 (192 968 Structure Total 27 139 59 832 Engines J24 494 \ 54 000 10 601 Thrust Reversers 4 809 1. 780 807 1 '780 807 Miscellaneous Systems 2 622 5 781 2 622 5 781 Fuel System- Tanks and Plumbing (35 377 ( 77 994 (27 923 (61 561 Propulsion Total 4 527 9 981 4 527 9 981 Surface Controls 1 542 3 400 1 542 3 400 Instruments 2 540 5 600 2 540 5 600 Hydraulics 2 291 5 050 2 291 5 050 Electrical 1 220 2 690 1 220 2 690 Avionics 25 111 11 390 25 111 Furnishings and Equipment 11 390 3 720 8 200 3 720 8 200 Air Conditioning 210 95 210 Ant i- icing (27 325 ( 60 242 (27 325 (60 242 Systems and Equipment Total : 142 776 314 771 Weight Empty 151 445 333 883 675 306 675 Crew and Baggage -Flight, 306 744 1 640 744 1640 -Cabin, 2 335 1 059 2 335 Unusable Fuel 1 059 361 795 361 795 Engine Oil 8 852 4 015 8 852 Passenger Service 4 015 Cargo Containers 1 343 2 960 1 343 2 960 351 140 150 604 332. 028 159 273 Operating Weight- 21 854 48 180 21 854 48 180 Passengers, (292) 5 828 12 848 Passenger Baggage 5 828 12 848 186 955 412 168 178 286 393 056 Zero Fuel Weight : 158 681 158 680 349 832 349 834 Mission Fuel 336 967 742 890 345 635 762 000 Design Gross Weight TABLE VII.-BASEPOINT ENGINE WEIGHT WEIGHT/VEHICLE ITEM Kg LB ENGINES (INCLUDING NOZZLE 24 312 53 600 § THRUST REVERSER) (4) 91 200 RESIDUAL FLUIDS 91 200 MISCELLANEOUS PROVISIONS 24 494 54 000 ENGINES AS INSTALLED TABLE VIII.-BASEPOINT NACELLE WEIGHT WEIGHT/VEHICLE ITEM Kg LB NACELLES 2 782 6 132 ENGINE COWL 2 864 1 299 INLET COWL 6 528 2 961 SPIKE ENGINE MDUNTS 7 410 16 336 TOTAL NACELLE Aerodynamics.-Friction drag estimates were made for a fully turbulent, hy- draulically smooth condition using the incompressible Von-Karman-Schoenherr method (reference 6) in conjunction with the adiabatic compressibility correction of Sommer and Short (reference 7). Component characteristic lengths (e.g. the distance from the inlet lip to the exhaust nozzle exit, the exposed mean aerodynamic chord of planar surfaces, etc.) and the altitude along the mission climb profile were used to evaluate length Reynolds numbers.

Flat plate values were increased by three percent to account for form losses.

The wave drag due to thickness was estimated as a function of Mach.

number using supersonic area rule theory (reference 8 and 9) in conjunction with a transparent wing simulation, an inlet mass flow ratio of one, and the nozzle exit area held fixed at its supersonic cruise position. The effect of inlet spillage and nozzle position is included in the installed thrust. All results reported here are based on the use of a 50-Mach-plane (AX = .02 L (0) ), 13-roll-angle (A© = 15°) analysis. Basepoint configuration results for increased solution mesh density did not indicate any appreciable change.

Supersonic cruise trimned drag-due-to-lift characteristics are assumed to be equal to the reference configuration of reference 2 and consequently independent of wing and engine size and nacelle shape. A different design wing twist and camber is required for each case to realize this performance.

The lifting efficiency may be conservative for some of the more favorably- shaped nacelles of the parametric drag study in that any increased benefit that may be realized from favorable nacelle thickness/wing lift interference over and above that of the reference configuration is neglected. Conversely, for the less favorably-shaped nacelles, the analysis may be somewhat opti- mistic. At off design conditions, the above assumption is necessary because the required analysis is beyond the scope of the contract effort.

A comparison of the VSCE 502B (408 kg./sec., 900 Ib./sec. airflow) nacelle of figure 11 to that of the reference configuration non-afterburning single spool turbojet with variable geometry turbine (363 kg./sec., 800 lb./ sec. airflow) of reference 2 is presented on figure 12. The basepoint nacelle is 1.95 meters (6.4 feet) shorter and has a 0.14 meter (0.46 feet) smaller maximum diameter. The relative cross-sectional shape of the two nacelles is presented on figure 13. The basepoint total configuration normal cross- sectional area distribution is shown on figure 14.

Estimated total and nacelle incremental skin friction and wave drag characteristics (relative to nacelles off) for the basepoint configuration are presented on table IX. The wave drag results are for the case in which the nozzle exit planes are the same as the reference configuration. A slightly higher drag results (AC = 0.00006 at Mach 2.7) if the inlet planes DW are matched.

The friction, wave, and total drag increments of the basepoint nacelle are compared to those of the reference nacelle on figure 15. The basepoint TABLE IX.-BASEPOINT CONFIGURATION ESTIMATED SKIN FRICTION AND WAVE DRAG CHARACTERISTICS S^^ = 929 sq.m. (10 000 sq.ft.)

M AL TITUDE AIR CRAFT NAC ELLE o m. ft.

_ — - — .4 457 1 500 0.0061 0.00065 — — — — .8 6 400 21 000 0.00572 0.00062 1.2 10 455 34 300 0.00545 0.00365 0.00060 -0.00017 1.4 11 521 37 800 0.00522 -0.00018 0.00316 0.00058 1.8 13 594 44 600 0.00490 0.00254 0.00055 -0.00019 2.32 16 764 55 000 0.00450 0.00222 0.00050 -0.00018 2.7 18 288 60 000 0.00418 0.00217 0.00046 -0.00014 o Tfr O * o o • oo

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configuration has a slightly smaller installation drag in spite of 12.5 percent greater airflow. It was subsequently determined that a further re- duction of 0.5 count could be realized by meridial contour optimization.

The aerodynamic characteristics used in resizing the basepoint wing and engine size to produce the baseline configuration used for all para- metric nacelle drag studies were established as follows: Fully turbulent friction levels were adjusted for difference in surface area and length Reynolds number of the wing and nacelle. The wave drag variation of the basepoint configuration as a function of wing and engine size were parametrically evaluated for input to the sizing program. The results are presented on figure 16. The effect of engine size was essentially nil at this scale for the nacelle shape under consideration.

The trimmed drag due to lift characteristics were assumed to be inde- pendent of wing size and equal to the reference configuration. The specific levels used are presented on figures 17 through 19 and were taken directly from reference 2.

Sizing of the basepoint configuration produced the study baseline (table I) which had a 12 percent smaller wing size and a 3.5 percent smaller engine size. The associated normal cross-sectional area distribution is presented on figure 20. A summary of the component surface areas and reference lengths is presented on table X, and table XI presents baseline drags.

Parametric Drag Analysis The parametric nacelle wave drag analysis utilized the baseline con- figuration described in the previous section. The installation of the propulsion system followed several general ground rules in order to preserve the basic arrangement concepts and provide consistent comparisons concerning the effect of nacelle size variations. They are: 1. Nacelle overhang of the wing trailing edge and vertical nacelle-wing separation was limited to the reference con- figuration values for structural reasons.

2. The longitudinal and lateral separation distance between the inboard and outboard nacelles was preserved in order to maintain inlet flow quality.

3. The reference configuration philosophy of locating the nacelle volume in a region of decreasing wing thickness was maintained.

The impact of these considerations in the propulsion system location for the largest and smallest nacelles of the study are presented on figure 21. The reference configuration is superimposed on these results for comparison purposes. The outboard nacelle is moved inboard and. forward as required along the mid-chord (approximate maximum thickness) line of the wing CO vO cr o o o Csl s O cr CNJ CT> u I -~7 00 V •H oo I/) • o DO nj -S I O PH 0) (U

I

•H I-L, 0.09 til 2.8 Mach number ~M_ Figure 17. - C versus Mach number T L K 0.0032 0.0028 0.0024 0.0020 0.0016 K 0.0012 0.0008 ._ 0.0004 0.0000 1.2 1.6 Mach number Figure 18. - C versus itech number n U K o -p

SP

f-l Q 0.2 2.4 2.8 1.2 1.6 Mach number-] Figure 19. - 'K' factor versus Mach number TABLE X.-BASELINE CONFIGURATION SURFACE AREA AND LENGTH SUMMARY Length Swet COMPONENT sq.m. (sq.ft.) m. (ft.)

96 (315) 786 (8 450) Fuselage 1 505 (16 987) 7.65-39.4 (25.1-129.)

Wing 276. (5 088) 12,7 (35.1) Nacelles (4) 4.9 (16.2) 20.1 (219) Center Line Vertical 7.9 (25.9) 91. (992) Wing Verticals 5.8 ( 1 8 . 9 ) 89.5 (921) Horizontal TABLE XI.-BASELINE CONFIGURATION ESTIMATED SKIN FRICTION AND WAVE DRAG CHARACTERISTICS = 929 sq.m. (10 000 sq. ft.)

AL TITUDE

AIM :RAFT NACE LLE

A

M m.

AC (ft.)

D o

%

%

\ F

0.00568 0.00065 0.4 457 (1 500)

—

0.00537 0.8 6 400 (21 000) 0.00061

—

10 455 (34 300) 1.2 0.00508 0.00339 0.00058 -0.00009 11 521 (37 800) 1.4 0.00489 0.00305 0.00056 0.00003 (44 600) 1.8 13 594 0.00455 0.00237 0.00052 -0.00011 2.32 (55 000) 16 764 0.00420 0.00209 -0.00012 0.00049 2.7 18 288 (60 000) 0.00392 0.00200 0.00045 -0.00011 LO CD o r^ •H ti O o O g 0) N •H tn (1) X o rt M-i O +J O 0) (Nl LO LO I •H U-,

_ , -

!

1 ^ "1 ~ 1 ~

LO S LO 0 0 until its trailing edge overhang does not exceed 3 meters (10 feet). The inboard nacelle is shifted laterally by the same amount holding the longi- tudinal distance between the inboard and outboard nacelle inlet planes the same as the reference configuration.

The nacelle parametric variables considered in the present analysis were the ratio of nozzle area to capture area A /Ac, the ratio of maximum n cross-sectional area to capture area AMAX/AC» the relative axial position of maximum area X^A^/A, the ratio of nacelle length to capture diameter, £/d , and the nacelle absolute capture area A^. A summary of the number c of variations and variable range analyzed is presented on table XII. For purposes of computation, the nacelles were assumed to be axisymmetric and the inlet, maximum area, and nozzle planes to be connected by straight lines, The parametric nacelle friction drag analysis is based on the use of fully turbulent flat plate levels in conjunction with the expression for surface areas (for four nacelles): X AMAX The largest deviation between the exact and approximate expression occurs for X^j^/£ approaching 0.4 and A /A approaching 2.0 with the former re- n c sulting in 10 percent greater area. It will be subsequently found that these differences are negligible in terms of the total installation drag for such cases. A summary of the parametric nacelle friction analyses based on the approximate surface area is presented on table XIII; the table shows friction drag increments relative to nacelles off. As would be expected, progressive penalties are associated with the increasing nacelle diameter and length.

Nacelle normalized cross-sectional area parametric extremes of the present study are presented in figure 22. Maximum-to-capture area ratio of 1 to 2 at 40, 60, and 80 percent of the nacelle length are shown for nozzle-to-capture area ratios of 1 and 2. In order to establish a feel for the corresponding physical variations, the results of figure 22 have been applied to a representative nacelle size ( A , - . = 2.6 sq.m., 28 sq.ft., i/d c = 6). The results are presented on figure 23 for the same range of variables used for the normalized area distribution. Note that the vertical scale has been expanded by a factor of 2 for clarity in presentation.

An index of the parametric wave drag results is presented on table XIV.

Each figure cited presents the wave drag increment relative to nacelles off as a function of the relative axial position of maximum thickness TABLE XII.-NACELLE PARAMETER VALUES PARAMETER VALUES 1.2, 2.32 Mach number 1.0, 1.25, 1.5, 2.0 A V c 1.0, 1.25, 1.5, 2.0 A /A MAX c 0.4, 0.6, 0.8, <1.0 * X / il AMAX ' 1.86, 2.79, 3.72 sq.m.

(20, 30, 40 sq.ft.)

5.5 and 7.0 * Maximum value considered corresponds to a boattail angle of ten degrees.

TABLE XIII.-SUMMARY OF PARAMETRIC FRICTION DRAG RESULTS i AC A £ MAX A = sq.m. Mo - 1.2 = 2.32 A C

(sq. ft.) C V

1.86 5.5 1.0 0.00033 0.00029 (20.0) 1.25 0.00034 0.00031 0.00036 0.00032 1.5 2.0 0.00039 0.00035 7.0 1.0 0.00040 0.00035 1.25 0.00042 0.00037 1.5 0.00044 0.00040 2.0 0.00048 0.00043 2.79 5.5 1.0 0.00047 0.00042 ( 3 0 . 0 ) 1.25 0.00050 0.00044 1.5 0.00053 0.00047 2.0 0.00057 0.00051 1.0 0.00058 7.0 0.00051 1.25 0.00061 0.00053 1.5 0.00065 0.00056 2.0 0.00070 0.00061 3.72 5.5 1.0 0.00062 0.00055 ( 4 0 . 0 ) 1.25 0.00066 0.00058 1.5 0.00069 0.00061 2.0 0.00075 0.00066 1.0 0.00076 0.00067 7.0 1.25 0.00080 0.00071 1.5 0.00085 0.00075 2.0 0.00092 0.00081 NOTE: Altitude is 10 455 m. (34 300 ft.) for M = 1.2 case and is Q 18 288 m. (60 000 ft.) for M = 2.32 case.

Q A/A 1.0 0.0 Figure 22.-Nacelle parametric cross-sectional area extremes o * \o o o • o o oo (SI o o • * to 00 OO (U

g

•H

tJ

<u p.

E I X 0) O o C-l

s

a) tn o as • ,0 oo to (SJ o • I o •H o P-, O o • • • CM O o o • o -ui - -m ~ TABLE XIV.-SUMMARY OF PARAMETRIC WAVE DRAG RESULTS ure * Fig A - sq. m. I ^MAX Mo 1.2 2.32 C (sq. ft.)

A c 1.86 5.5 1.25 24 30 (20.0) 1.5 25 31 2.0 26 32 7.0 1.25 27 33 1.5 28 34 2.0 29 35 2.79 5.5 1.25 36 42 (30.0) 1.5 37 43 2.0 38 44 7.0 1.25 39. 45 1.5 40 46 2.0 41 47 1.25 48 3.72 5.5 54 ( 4 0 . 0 ) 1.5 49 55 2.0 50 56 7.0 1.25 51 57 1.5 52 58 2.0 53 59 * Each figure cited implicitly contains the result for a cylindrical nacelle; this corresponds to the abscissa (i.e. ACn = 0).

0.005 0.004 0.003 0.002 AC r 0.001 O.OOQ -0.001 -0.002 1.0 X /£ AMAX Figure 24. -Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q 5.5 A = 1.86 sq. m. ( 20 sq. ft.)

C l 2S - ' 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 X /£ AMAX Figure 25.- Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q A - 1.86 sq. m. ( 20 sq. ft.) l/d = 5.5 C c 0.005 0.004 0.003 0.002 0.001 0.000 -0.001 -0.002 1.0 X /£l AMAX Figure 26.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q A^ « 1.86 sq. m. ( 20 sq. ft.) «,/<*<- = 5.5 0.005 0.004 0.003 0.002 AC D, 'W 0.001 0.000 -0.001 -0.002 1.0

o.o

Figure 27.-Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 A - 1.86 sq. m. ( 20 sq. ft.) £/d = -° C c L25 - 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 X /£ AMAX Figure 28.-Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q A - 1.86 sq. m. ( 20 sq. ft.) £/d = 7.0 C c

o.oos

0.004 0.003 0.002 AC T 0.001 0.000 -0.001 -0.002 0.8 1.0 Figure 29. -Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q A »= 1.86 sq. m. ( 20 sq. ft.) = 7.0 C 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 Figure 30.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A,. = 1.86 sq. m. (20 sq. ft.) i/d = 5.5 c *•* 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 Figure 31. - Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 = 5.5 A = 1.86 sq. m. (20sq. ft.)

c A A l 5 MAX/ c ' ' 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 AMAX/S, Figure 32.- Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q A = 1.86 sq. m. (20 sq. ft.)

C 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 A AMAX/£ Figure 33. - Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q A = 1.86 sq. m. (20sq. ft.) 1/& = 7.0 c C 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 Figure 34.- Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q = 7.0 A = 1.86 sq. m. ( 20 sq. ft.)

c 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 A AMAX/£ Figure 35.- Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q A = 1.86 sq. m. (20sq. ft.) Vd = 7.0 c c

o.oos

0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 X /£ AMAX Figure 36.-Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q A - 2.79 sq. m. ( 30 sq. ft.) A/d = 5.5 c c = 1.25 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 0.8 1.0 X /Jl AMAX Figure 37.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1. 2 Q A * 2.79 sq. m. ( 30 sq. ft.) i/d = 5.5 C c 0.005 0.004 0.003 0.002 AC °W 0.001 0.000 -0.001 -0.002 0.2 0.0 Figure 38.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1. 2 Q A - 2.79 sq. m. ( 30 sq. ft.) £/d = 5.5 C c A /A - 2 .0 MAX c 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 Figure 39.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1. 2 Q A - 2.79 sq. m. ( 30 sq. ft.) Si/d =7.0 C c

o.oos

0.004 0.003 0.002 AC, 0.001 0.000 -0.001 -0.002 Figure 40.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q A * 2.79 sq. m. ( 30 sq. ft.) SL/d = 7.0 c 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 X //£ AMAX Figure 41.-Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q A - 2.79 sq. m. ( 30 sq. ft.) £/d =7.0 c c 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 _ -0.0004 Effect of geometry perturbations on nacelle Figure 42.- incremental wave drag at M = 2.32 Q A = 2.79 sq. m. (30sq. ft.) -£/d = 5. 5 c A 25 W c ' i' 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 _ -0.0004 (LT 0.4 0.6 0.8 1.0 AMAX/i Figure 43. - Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A = 2.79 sq. m. (30 sq. ft.) i/d = 5.5 c c 0.0016 0.0012 0.0008 0.0004 AC r 0.0000 -0.0004 -0.0008 X /Jl AMAX Figure 44.- Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q A = 2.79 sq. m. (30 sq. ft.) i/d = 5. 5 c c 0.0008 ..!..;; I ; !; 0.0006 :i;;i

•1-4

0.0004 A V c 0.0002 ... i.

1.0 0.0000 1.25 -0.0002 -0.0004 1.0 0.2 0.4 0.6 0.8 0 0 X, Effect of geometry perturbations on nacelle Figure 45.- incremental wave drag at M = 2.32 Q = 7.0 A = 2.79 sq. m. ( 30 sq. ft.)

C = 1.25 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 Effect of geometry perturbations on nacelle Figure 46.- incremental wave drag at M =2.32 Q = 7.0 A = 2.79 sq. m. (30 sq. ft.)

c

= 1.5

0.0016 0.0012 0.0008 0.0004 AC T 0.0000 -0.0004 -0.0008 1.0 Figure 47.- Effect of geometry perturbations on nacelle incremental wave drag at M =2.32 Q A =2.79sq. m. (30 sq. ft.) l/d = 7.0 c c 0.005 0.004 0.003 0.002 AC r 0.001 O.OOQ -0.001 -0.002 Figure 48.-Effect of geometry perturbations on nacelle incremental wave drag at M * 1. 2 Q A - 3.72 sq. m. ( 40 sq. ft.) £/d = 5.5 c c 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 0.6 0.8 1.0 X /£ AMAX Figure 49. -Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q - 3.72 sq. m. ( 40 sq. ft.) Vd = 5.5 c Q 0.005 0.004 0.003 0.002 AC T 0.001 0.000 -0.001 -0.002 Figure 50. - Effect of geometry perturbations on nacelle incremental wave drag at M = 1.2 Q A * 3.72 sq. m. ( 40 sq. ft.) H/d = 5.5 C c = 2.0 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 0.6 0.8 1.0 X /£ AMAX Figure 51.-Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q A - 3.72 sq. m. ( 40 sq. ft.) i/d =7.0 C c = l 2S ' 0.005 0.004 0.003 0.002 AC |D

W

0.001 0.000 -0.001 -0.002 Figure 52. -Effect of geometry perturbations on nacelle incremental wave drag at M =1.2 Q 7.0 "= 3.72 sq. m. ( 40 sq. ft.)

C 0.005 0.004 0.003 0.002 AC r 0.001 0.000 -0.001 -0.002 X /£ AMAX Figure 53.-Effect of geometry perturbations on nacelle incremental wave drag at M = 1. 2 Q A - 3.72 sq. m. ( 40 sq. ft.) SL/d = 7.0 C c 0.0008 0.0006 0.0004 AC T 0.0002 0.0000 -0.0002 -0.0004 X L AMAX/£ Figure 54.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A = 3.72 sq. m. ( 40 sq. ft.) A/d = 5.5 c c 0.0008 0.0006 0.0004 AC T 0.0002 0.0000 -0.0002 -0.0004 Effect of geometry perturbations on nacelle Figure 55.- incremental wave drag at M = 2.32 Q A = 3.72 sq. m. (40 sq. ft.) = 5.5 C =l S ' 0.0016 0.0012 0.0008 0.0004 AC 0.0000 -0.0004 -0.0008 X /S< AMAX Figure 56.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 A =3.72sq. m. (40 sq. ft.) i/d = 5.5 c c 0.0008 0.0006 0.0004 AC T 0.0002 0.0000 -0.0002 -0.0004 Figure 57.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A = 3.72 sq. m. (40 sq. ft.) i/d^ =7.0 C 0.0008 0.0006 0.0004 0.0002 0.0000 -0.0002 -0.0004 Figure 58.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A = 3.72 sq. m. (40 sq. ft.) i/d = 7 .0 c c i6= i!£ i 1.0 !; 0.0016 ::;;!;;; 0.0012 1.25 0.0008 1.5 .—,_.

0.0004 AC r 2.0

1-4—

: I.

0.0000 -0.0004 ~f-- -0.0008 0.1 0.0 0.2 0.4 0.6 1.0 X //£ AMAX Figure 59.- Effect of geometry perturbations on nacelle incremental wave drag at M = 2.32 Q A =3.72sq. m. (40 sq. ft.) i/d = 7,0 c c = 2.0 and the ratio of nozzle area to capture area An/Aj.. In all cases, the max- imum value of the former variable considered corresponds to a boattail angle of ten degrees. The results can be easily re-interpreted as total nacelle installation drag by a simple vertical shift of origin in accordance with table XI.

A discussion of the results is presented for a representative case of a medium size nacelle (AC = 2.79 sq.m., 30 sq. ft.) for a fineness ratio of 5.5 at Mach 2.32 cruise condition. Examination of figures 42 through 44 indicates the incremental nacelle wave drag is a strong function of the ratio of maximum-to-capture cross-sectional area, A_/A , boattail area, and to a somewhat lesser extent relative axial position of maximum cross-sectional area, ^AMAY/ " Th° results for a circular cylinder correspond to the axis = ie act t at (i.e. ^ T ) W ^* ^ ^ ^ nacelle shapes of less wave drag exist is a consequence of favorable total system thickness interferences associated with the location of growing nacelle cross-sectional area in a region of decreasing wing thickness.

The previously-cited nacelle geometric variable behavior and sensitivity are unchanged by Mach number, nacelle capture area, or nacelle fineness ratio.

The incremental wave drag results are, in general, weak functions of the latter two variables for efficient installations (see figure 6 0 ) .

Detailed nacelle wave drag variations with freestream Mach number are defined for a range of levels covering high-positive, zero, and negative installation increments. These characteristics correspond to nacelles with large maximum cross-sectional area relative to the capture and nozzle area, cylindrical, and near-truncated conical shapes, respectively. Figure 61 illustrates the Mach number difference for these extremes for the fineness ratio 5.5 medium size nacelle. Examination of the results indicate that weak to moderate Mach number variations are associated with small nacelle install- ation drags. Conversely, strong compressibility variations are exhibited for inefficient installations. The large benefit at transonic speeds is somewhat illusory as the thrust must be progressively penalized for nozzle contraction with decreasing Mach numbers.

Weight Sensitivity Analysis To evaluate the effects of propulsion system variations on the airframe/ engine system, the sensitivities of the airplane takeoff gross weight to varia- tions of propulsion system parameters were determined. These sensitivity data were obtained by conducting design trades on the baseline airplane for varia- tions of the following items: • Incremental Nacelle Drag • Propulsion System Weight • Engine Specific Fuel Consumption • Takeoff Thrust 0.0010 5.5 7.0 0.0008 0.0006

J

0.0004 X, = 2.0, ™x-0.6, >= 1.25 0.0002 AC 0.0000 -0.0002 X A, VWAX = 1.0, * = 2.1 -0.0004 -0.0006 (20) (24) (28) (32) (36) ( 4 0 ) i I J I 8.0 10.0 12.0 6.0 A ~sq.m. (sq.ft.)

Figure 60.-Typical nacelle incremental wave drag variation with nacelle size at Mach 2.32 0.003 0.002 0.001 0.000 -0.001 - -0.002 -0.003 Figure 61. -Typical nacelle incremental wave drag variations with Mach number A = 2.79 sq.m. (30 sq.ft.) i/d =5.5 c c In each case the parameter of interest was varied independently and the airplane resized to the design mission range of 7 408 km. (4 000 n.mi.)

while maintaining thrust-to-weight and wing loading values equal to those for the baseline vehicle.

Incremental nacelle drag. -Several variations of nacelle drag were investi- gated!These were chosen as representative of the combined wave and friction drag variations as found in the nacelle shape analysis to allow use of the trade data for any nacelle geometry analyzed in this program.

The results of this trade are shown in figure 62 which shows relative TOGW versus nacelle drag at Mach 2.32 for several variations of the drag increment at 1.2 Mach. A typical curve of drag increment versus Mach number is depicted in figure 63.

Propulsion system weight trades. -Airplane takeoff gross weight was calculated for several propulsion system weight increments. Incremental propulsion weight, in this case, is defined as a percent of the sum of the engine, engine accessories, and nacelle weights. For the baseline airplane, having an engine airflow of 394 kg./sec. (869 Ib./sec.), these total 31 522 kg.

(69 439 lb.). The results of this trade are shown in figure 64 which plots relative TOGW versus propulsion weight increment.

Engine specific fuel consumption trades.-Four separate trades were performed with SFC increments applied independently to the following mission segments: (1) maximum power climb legs only (2) supersonic cruise legs only (3) subsonic cruise and loiter legs only (4) the entire mission.

The results of this trade are presented in figure 64 as relative takeoff gross weight versus percent change in SFC.

Takeoff thrust.-Figure 65 presents relative TOGW versus percent change in takeoff thrust. In this trade it is assumed that the thrust available during both ground roll and climbout portions of the takeoff vary without any change in propulsion characteristics at other flight conditions.

DISCUSSION OF RESULTS Nacelle Drag Following is a discussion of the parametric nacelle installation drag results at Mach 2.32 cruise conditions. The presentation of cuts through the basic data of figures 24 through 59 emphasize the low drag variable domain as it is necessarily the most interesting. In order to keep the results in proper perspective, it should be remembered that a typical super- sonic cruise condition for the baseline configuration has a total drag : ; . . .

: : ; : : : : : ! . : • • ' !

:: t ; - ; : | : : : ! • I. - ' \ \ : i ! : ! ! • \ ; : : | . - Na< ;e] .le drag @ 1.2 Mach i . !

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1 iange 7 408 km. (4 000 n.mi.)

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I tef. TOGW 323 048 kg. (712 188 Ib.)

0,9 0.001 0.002 Nacelle drag @ 2.32 Mach Figure 62.-Nacelle drag sensitivity trade 0.004 0.003 ; 0.002 i:!ii*: 0.001 0.000 1.2 1.

0. 0.4 0.8 Mach number Figure 63•-Typical incremental nacelle drag variation 1.20 Supersonic cruise 1.10

I

1.00] *-> rt : 0.90 0.80: - 4 0 4 SFC increment - percent 1.08, Baseline propulsion weight 31 522 kg. (69 439 Ib.)

T/W (0.323 Ib./lb.)

3.16 n./kg.

354. kg./sq.m. (72.5 Ib./sq. ft.)

: W/S (4 000 n.mi.)

^ Range 7 408 km.

0.92 (712 188 Ib.)

i Ref. TOGW 323 048 kg.

- 8 - 4 0 4 8 -12 Propulsion weight increment - percent Figure 64.-Propulsion weight and SFC sensitivity trades T/W 3.16 n./kg. (0.323 Ib./lb.)

W/S 354. kg./sq.m. (72.5 Ib. /sq.ft.)

Range 7 408 km. (4 000 n.mi.)

Ref. TOGW 323 048 kg. (712 188 Ib.)

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0.0000 1.0 1.2 1.4 1.6 1.8 2.0 •^VlAX^c Figure 66. -Typical nacelle incremental drag variation with boattail area at M = 2.32 Q coefficient of approximately 0.0105 at a C, of 0.1 Typical impact on installation drag of nacelle relative maximum cross- sectional area and nozzle area for a fixed axial location of maximum area is presented on figure 66. Clearly, both variables have a strong influence on the overall installation efficiency through the wave drag-due-to-thickness component.

The sensitivity of the nacelle drag to the axial position of maximum cross-sectional area is presented on figure 67. It is generally less im- portant than either the maximum or nozzle area parameters although its pro- per selection can significantly aid in realizing a propulsion system installation goal of zero total drag.

The influence of the absolute size of the nacelle on installation drag is presented on figure 68. As would be expected, inclusion of friction drag results in more pronounced size dependence than observed previously for the wave drag component (figure 60). This range of nacelle sizes could repre- sent a wide range of engine thrust and, therefore, vehicle thrust-to-weight ratios.

The parametric nacelle analysis has identified a family of quasi-conical AMAX n ^4AX i —o— approaching 1, -r— approaching -r approaching 21 which have

(

A *• c c a potential installation drag (wave plus friction) of zero at the supersonic cruise condition when modest allowances (ACL of approximately -0.0002) are made for optimum meridial shaping refinements and the benefit of wing lift/ nacelle thickness interference is accounted for. These propulsion system integrations effectively remove 200 to 400 sq.m. (2 000 - 4 000 sq.ft.), depending on nacelle size, of wetted area by use of favorable pressure drag interference.

Weight Sensitivity Results of the weight sensitivity trades show that airplane gross weight is highly sensitive to both drag and engine specific fuel consumption (SFC) at the supersonic cruise condition of 2.32 Mach number just as expected. On the other hand changes in drag and engine SFC at other flight conditions and even propulsion system weight changes have a relatively minor effect on air- plane gross weight. However, the baseline airplane has a relatively large thrust margin throughout the transonic acceleration. If that margin were reduced, gross weight would become more sensitive to drag variations at Mach 1.2 Variations in engine thrust at the takeoff condition have a significant impact on gross weight although not so great as that for drag and SFC at supersonic cruise conditions. Small thrust variations at conditions other than takeoff have only a minor impact on gross weight. Table XV compares the changes in parameters required to achieve a 4 540 kg. (10 000 Ib.) re- duction in takeoff gross weight. It can be seen from table XV that a one- drag count change at supersonic cruise results in a vehicle takeoff gross weight change approximately equal to that of a one percent SFC change at supersonic cruise.

0.0018 0.0016 0.0014 0.0012 0.0010 AC 0.0008 0.0006 0.0004 0.0002 0.0000 Figure 67. -Typical nacelle incremental drag variation vs. axial position of maximum area at M 2.32 Q A = 2.79 sq.m. (30 sq.ft.) /- = 5.5 = -° -Lt TABLE XV.-SENSITIVITY COMPARISON Parameter Change required to achieve a 1.41 reduction in vehicle takeoff gross weight -0.00013 Supersonic cruise drag coeffi- cient (C e 1.2 Mo = baseline) D ASFC, % -1.3 Supersonic cruise -0.9 Total mission -4.3 APropulsion weight, I ATakeoff thrust. % +3.3 0.0018 0.0016 —F-— 0.0014 0.0012 0.0010 1C D 0.0008 0.0006 -- 0.0004 0.0002 0.0000 (32) (36) (40> (20) (24) (28) I 12.0 6.0 8.0 10.0 A ~ sq.m. (sq.ft.)

Figure 68.-Typical nacelle installation drag variation with nacelle size at Mach 2.32 RECOMMENDATIONS It is recommended that the nacelle drag and vehicle sensitivity data generated in this program be used in future preliminary analyses of propulsion system trades.

It is further recommended that airframe/engine integration studies con- tinue. In particular, investigations should include the following: 1) Close coordination between the engine companies and the airframe manufacturers is essential to achieve maximum benefit from all disciplines and technology advances.

2) Consideration should be given to vectorable, non-axisymmetric nozzles. These nozzles offer the potential advantages of reduced afterbody drag and vectored thrust with lift enhance- ment to reduce wing size (currently being set by takeoff re- quirements) .

3) A study should be made to determine the angle at which exhaust nozzles should be canted. Theoretically, directing the nozzles downward relative to the freestream can produce sufficient lift (without excessively penalizing horizontal thrust) to signi- ficantly reduce wing size. However, this introduces a pitching moment; whether a canard is required remains to be determined.

REFERENCES 1. Bonner, E., et al, Influence of Propulsion System Size, Shape, and Location on Supersonic Aircraft Design, Rockwell International, NASA CR 132544~;December 1974.

2. NASA CR-132374, Advanced Supersonic Technology Concept Study Reference Characteristics, Hampton Technical Center LTV Aerospace Corporation, 21 December, 1973.

3. Lockheed-California Company Report LR 26133, An Airline's View of Reserve Fuel Requirements for the Supersonic Transport, 19 September 1973.

4. Bushell, K. W., Measurement and Prediction of Jet Noise in Flight, AIM Paper 75-461.

5. Crosthwait, E. L., Kennon, Jr., I.G., and Roland, H.L., et al, Preliminary Design Methodology for Air-Induction Systems, General Dynamics, Fort Worth Division, Technical Report SEG-TR-67-1, January 1967.

6. Schoenherr, K. E., Resistance of Flat Plates Moving Through a Fluid, Transactions of Society of Naval Architects and Marine Engineers, Vol 40, pp 279-313, 1932.

7. Sommer, S., and Short, B., Free Flight Measurements of Turbulent Boundary Skin Friction in the Presence of Severe Aerodynamic Heating at Mach" Numbers from 2.8 to 7.U, NACA TN 5591, 1955.

REFERENCES (Concluded) . .

8. Lomax, H., The Wave Drag of Arbitrary Configurations in Linearized Flow as Determined by Areas and Forces in Oblique Planes, NA.CA RM A55A18, 1955 9. Bonner, E., Theoretical Prediction of Supersonic Pressure Drag, Rockwell International Report MA-66-862, 1966.