DESIGN POINT M=2.7,CL=.10 ARBITRARY REGION (REF) 1.Basic wing solution(no fuselage) 2.CPleve limit =.7 vacuum 3.Gradient limit set by maximum obtained with optimum 3 load solution 6 Flagged symbols are for solutions having same maximum wing upper surface pressure gradient .5 3 loads 11 loads .1 0 -.004 0 .004 .008 .012 .016 .020 .024 Cmo FIGURE 3.7-2.-EFFECT OF PRESSURE CONSTRAINTS 22
DESIGN POINT J M = 2"7_ CL = .10 ARBITRARY REGION (REF) /-27 Basic wing solution (no fuselage) __" ,// ,. 2. CPlevel limit = .7 vacuum 3. Gradient limit set by maximum _/- %,&/ i obtained with optimum 3 load _ I _ I solution CD CL 2 I -.004 .6 .5 .4 .3 .2 Flagged symbols are for solutions having same maximum wing upper surface pressure gradient 3 loads I I I J I .004 .008 .012 .016 .020 Cm o 11 loads .024 FIGURE 3. 7-2. -EFFECT OF PRESSURE CONSTRAINTS 22
TABLE 1.-DESCRIPTION OF WING LOADING TERMS Loading Number Definition 1. Uniform 3 Proportional to x,the distance from the leading edge 3. Proportional to y,the distance from the wing centerline 4. Proportional to y2 5 Proportional to x2 6. Proportional to x(c-x).where c is local chord 7. Proportional to x2 (1.5c-x) 8. Proportional to 2(1+5)-0.5 9 Proportional to n2 (n-1)2 10. Elliptical spanwise,proportional to(1-) 11. Proportional to x,the distance from the leading edge of an arbitrarily defined region 12. A camber-induced loading proportional to the body buoyancy loading 13. A camber-induced loading proportional to the body upwash loading 14. A camber-induced loading proportional to the nacelle buoyancy loading 15. The body buoyancy loading 16. The body upwash loading 17. The nacelle buoyancy loading 23
Loading Number TABLE I.-DESCRIPTION OF WING LOADING TERMS Definition 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. Uniform Proportional to x, tile distance from the leading edge Proportional to y, the distance from the wing centerline Proportional to y2 Proportional to x 2 Proportional to x(c - x), where c is local chord Proportional to x 2 (1.5 c -x) Proportional to 2 (1 + 5 x)-0.5 C Proportional to _2 (rt - 1)2 Elliptical spanwise, proportional to _ (1 - r_) Proportional to x, the distance from the leading edge of an arbitrarily defined region A camber-induced loading proportional to the body buoyancy loading A camber-induced loading proportional to the body upwash loading A camber-induced loading proportional to the nacelle buoyancy loading The body buoyancy loading The body upwash loading The nacelle buoyancy loading 23
● As an independent effect,the configuration-dependent loading acts upon the wing in the optimization process,but cannot be varied (loadings 15-17). ● As a loading definition (12-14),a configuration-dependent loading produced by wing camber may be introduced in addition to its independent effect.The optimization then could cancel the lift of the independent effect with this camber-generated loading,if that were the optimum solution. In general,a configuration-dependent loading may not be used as the source of a variable loading without also using it as an independent loading. An exception is the nacelle pressure field loading,which may be used as a camber surface loading independent of the configuration dependent loading.The reason for this treatment is to allow input of a special loading as the nacelle pressure field,for use either in the optimization process or as a single loading.This feature is discussed on page 30. Fuselage in wing design solution.The fuselage may be included in the wing design solution by input of fuselage geometry and specifying a side-of-fuse- lage semispan station.The resulting solution is then split into two parts: the wing part (outboard of side-of-fuselage station)with loading definitions as described previously,and the "fuselage"part (inboard of side-of-fuselage station).Loadings inboard of the side-of-fuselage station are of the "carry-over"type,and are dependent functions of the loadings outboard of the side-of-fuse lage. Drag of the "fuselage"part is calculated by applying the carry-over loadings to the fuselage camberline.The outboard,or wing,part is handled as for the wing alone case,with integrations beginning at the side-of-fuselage.The wing-fuselage solution thus reflects the interdependence of wing and fuselage contributions to the wing design optimization. There are several considerations of importance in the wing/fuselage solution: ● Wing paneling (internal definition of wing geometry)may require a slight shift in the input side-of-fuselage station. This is accomplished automatically in the program and an explanatory note is printed. ● The fuselage attitude and wing camberline at the side-of-fuselage must approximately align for the drag integrations to be valid. Experience has shown that it is necessary for Z constraints to be applied to the wing camberline at the side-of-fuselage for this to occur.(Z constraints are discussed in more detail on page 36.) For convenience,the fuselage attitude (relative to the basic geometry definition)in both the lift analysis program (to generate 24
As an independent effect, the configuration-dependent loading acts upon the wing in the optimization process, but cannot be varied (loadings 15-17). As a loading definition (12-14), a configuration-dependent loading produced by wing camber may be introduced in addition to its independent effect. The optimization then could cancel the lift of the independent effect with this camber-generated loading, if that were the optimum solution. In general, a configuration-dependent loading may not be used as the source of a variable loading without also using it as an independent loading. An exception is the nacelle pressure field loading, which may be used as a camber surface loading independent of the configuration dependent loading. The reason for this treatment is to allow input of a special loading as the nacelle pressure field, for use either in the optimization process or as a single loading. This feature is discussed on page 30. Fuselage in win 9 design solution. - The fuselage may be included in the wing design solution by input of fuselage geometry and specifying a side-of-fuselage semispan station. The resulting solution is then split into two parts: the wing part (outboard of side-of-fuselage station) with loading definitions as described previously, and the "fuselage" part (inboard of side-of-fuselage station). Loadings inboard of the side-of-fuselage station are of the "carry-over" type, and are dependent functions of the loadings outboard of the side-of-fuselage. Drag of the "fuselage" part is calculated by applying the carry-over loadings to the fuselage camberline. The outboard, or wing, part is handled as for the wing alone case, with integrations beginning at the side-of-fuselage. The wing-fuselage solution thus reflects the interdependence of wing and fuselage contributions to the wing design optimization. There are several considerations of importance in the wing/fuselage solution: Wing paneling (internal definition of wing geometry) may require a slight shift in the input side-of-fuselage station. This is accomplished automatically in the program and an explanatory note is printed. The fuselage attitude and wing camberline at the side-of-fuselage must approximately align for the drag integrations to be valid. Experience has shown that it is necessary for Z constraints to be applied to the wing camberline at the side-of-fuselage for this to occur. (Z constraints are discussed in more detail on page 36.) For convenience, the fuselage attitude (relative to the basic geometry definition) in both the lift analysis program (to generate 24
upwash loading)and the wing design program can be changed without revising the basic geometry.(In the lift analysis module,this is done by a special application of the pressure limiting option.Set FLIMIT=1.0,an appropriate value of vacuum fraction VACFR,and define the fuselage angle of attack in TLALP). Inclusion of the fuselage in the wing design solution is illustrated in figure 3.7-3.In the example case shown,the wing camberline at the side-of-fuselage was constrained at the four locations indicated.A bucket plot is not produced when Z constraints are used;however,the effect of adding pressure constraints to the solution is illustrated by the symbols on the drag-due- to-lift versus Cmo plot. Use of configuration-dependent loadings.An example of the inclusion of a configuration-dependent loading is illustrated in figure 3.7-4,showing "reflexing"of the wing due to nacelle influences.The wing trailing edge is bent upward locally,or reflexed,to take advantage of positive pressure coefficients from the nacelle pressure field. The loadings due to the fuselage include both lift caused by the fuselage upwash field and also lift due to asymmetric distribution of fuselage volume above and below the wing (non mid-wing arrangement). As a special case,the asymmetric fuselage buoyancy loading (number 15),can be used even if its net lift is zero.This feature permits the inclusion of fuselage thickness pressures in the pressure limiting case for a mid-wing arrangement.However,if the fuselage buoyancy lift is zero,the wing camber loading proportional to the fuselage buoyancy loading (number 12)cannot be used,since it would cause the optimization solution to fail. An option in the wing design program permits card input of the configuration dependent loadings. As calculated by the near-field and lift analysis programs, these loadings are written on a file in the wing pressure summary program (see page 54).This option is provided to allow editing of the configuration dependent loadings and/or input of special loadings in their place. Optimization of the wing design considering influence of the fuselage upwash field is performed iteratively,using both the wing design and lift analysis modules.A fuse lage shape and incidence is first assumed,the upwash field and corresponding loading is calculated by the analysis program,and the design solution is performed.Because the resulting wing shape probably differs from the shape used in the initial upwash solution,the upwash loading used is approximate. It may be desirable to then rerun the wing design solution and/or alter the fuselage angle of attack.A representative program executive card sequence would be: 25
upwashloading) and the wing design program can be changedwithout revising the basic geometry. (In the lift analysis module, this is done by a special application of the pressure limiting option. Set FLIMIT=I.O, an appropriate value of vacuumfraction VACFR,and define the fuselage angle of attack in TLALP). Inclusion of the fuselage in the wing design solution is illustrated in figure 3.7-3. In the examplecase shown, the wing camberline at the side-of-fuselage was constrained at the four locations indicated. A bucket plot is not produced whenZ constraints are used; however, the effect of adding pressure constraints to the solution is illustrated by the symbols on the drag-dueto-lift versus Cmoplot. Use of configuration-dependent loadings. - An example of the inclusion of a configuration-dependent loading is illustrated in figure 3.7-4, showing "reflexing" of the wing due to nacelle influences. The wing trailing edge is bent upward locally, or reflexed, to take advantage of positive pressure coefficients from the nacelle pressure field. The loadings due to the fuselage include both lift caused by the fuselage upwash field and also lift due to asymmetric distribution of fuselage volume above and below the wing (non mid-wing arrangement). As a special case, the asymmetric fuselage buoyancy loading (number 15), can be used even if its net lift is zero. This feature permits the inclusion of fuselage thickness pressures in the pressure limiting case for a mid-wing arrangement. However, if the fuselage buoyancy lift is zero, the wing camber loading proportional to the fuselage buoyancy loading (number 12) cannot be used, since it would cause the optimization solution to fail. An option in the wing design program permits card input of the configuration dependent loadings. As calculated by the near-field and lift analysis programs, these loadings are written on a file in the wing pressure summary program (see page 54). This option is provided to allow editing of the configuration dependent loadings and/or input of special loadings in their place. Optimization of the wing design considering influence of the fuselage upwash field is performed iteratively, using both the wing design and lift analysis modules. A fuselage shape and incidence is first assumed, the upwash field and corresponding loading is calculated by the analysis program, and the design solution is performed. Because the resulting wing shape probably differs from the shape used in the initial upwash solution, the upwash loading used is approximate. It may be desirable to then rerun the wing design solution and/or alter the fuselage angle of attack. A representative program executive card sequence would be: 25
Event Executive Card Define fuselage GEOM Calculate upwash loading ANLZ (WHUP=1.0) Wing design solution WDEZ Recalculate upwash ANLZ (WHUP=1.0,TIFZC=3.0) loading with new camber surface Wing design WDEZ When the wing camber surface is finalized,it may be transferred into the basic geometry by the executive control card WGUP.(With interactive graphics attached,the design wing shape may also be viewed and edited between design and analysis solutions. Nacelle pressure field options.Special options are available for use with the nacelle pressure field (NPF)loading.The NPF is given special treatment because of the character of the NPF and because it is the most general loading that can be input in table lookup format.(This general input format makes the NPF useful as a means of inputting a desired loading which may not be the nacelle pressure field.) The typical nacelle pressure field,as calculated by the near-field or analysis programs,is characterized by pressure discontinuities associated with shock waves in the flow (e.g.,figure 3.7-4).When the NPF is then used to calculate the corresponding camber surface,undesirable discontinuities are required in the surface shape.An option is provided to alleviate this difficulty,which preserves the general nature of the NPF but substitutes an approx imate (smoothed)NPF for the camber surface calculation.This option is controlled by the input code FNPSMO.A typical improvement in camber surface smoothness is shown in figure 3.7-5. Regardless of the code FNPSMO,the configuration-dependent loading of NPF (loading 17),if used,uses the NPF without smooth ing. The nacelle pressure field,like the other configuration-dependent pressure tables,may be input into the wing design program to define a special loading to be used either alone or with other loadings.In this usage,the configuration-dependent loading version of the NPF (loading 17)cannot also be used since there is only one definition of the NPF (except for the smooth form noted above,which would not be of interest).In this case,the configuration- dependent version of the NPF could be combined with one of the other configu- ration-dependent loadings and input in card form (see page 80). 26
Event Define fuselage Calculate upwash loading Wing design solution Recalculate upwash loading with new camber surface Wing design Executive Card GEOM ANLZ (WHUP=I.0) WDEZ ANLZ (WHUP=I.0, TIFZC=3.0) WDEZ When the wing camber surface is finalized, it may be transferred into the basic geometry by the executive control card WGUP. (With interactive graphics attached, the design wing shape may also be viewed and edited between design and analysis solutions.) Nacelle pressure field options. - Special options are available for use with the nacelle pressure field (NPF) loading. The NPF is given special treatment because of the character of the NPF and because it is the most general loading that can be input in table lookup format. (This general input format makes the NPF useful as a means of inputting a desired loading which may not be the nacelle pressure field.) The typical nacelle pressure field, as calculated by the near-field or analysis programs, is characterized by pressure discontinuities associated with shock waves in the flow (e.g., figure 3.7-4). When the NPF is then used to calculate the corresponding camber surface, undesirable discontinuities are required in the surface shape. An option is provided to alleviate this difficulty, which preserves the general nature of the NPF but substitutes an approximate (smoothed) NPF for the camber surface calculation. This option is controlled by the input code FNPSMO. A typical improvement in camber surface smoothness is shown in figure 3.7-5. Regardless of the code FNPSMO, the configuration-dependent loading of NPF (loading 17), if used, uses the NPF without smoothing. The nacelle pressure field, like the other configuration-dependent pressure tables, may be input into the wing design program to define a special loading to be used either alone or with other loadings. In this usage, the configuration-dependent loading version of the NPF (loading 17) cannot also be used since there is only one definition of the NPF (except for the smooth form noted above, which would not be of interest). In this case, the configurationdependent version of the NPF could be combined with one of the other configuration-dependent loadings and input in card form (see page 80). 26