* * File : ShortS23.dat * * Purpose : This is an input file for the Digital Datcom program. * * Author : Anders Gidenstam * * Based on a file by: * Bill Galbraith * Holy Cows, Inc. * billg (at) holycows.net * * Model-Specific notes: * o The origin is at the nose of the aircraft. * * * General Notes : * 1. In general, we aren't concerned with anything faster than the * subsonic speed regime. DATCOM can handle transonic, supersonic, * and hypersonic, so check the book for more information. * * 2. DATCOM pg 47 states : * * "In general, the eight flap types defined using SYMFLP * (variable FTYPE) are assumed to be located on the most * aft lifting surface, either horizontal tail or wing if * a horizontal tail is not defined." * * We are going to assume that this means that we will have to * perform a wing-body-vertical_tail analysis in order to get * the flap effects, then add the horizontal tail in order to * get it's effects and trim points. The first case is used * ONLY to get the flap effects. * * 3. The page references below are from The USAF Stability and * Control Digital DATCOM, volume 1, Users Manual, April 1979 * * ************************ * List of Command Card ************************ * * Note : Command cards MUST start in column 1. They may appear in * any order, with the exception of NEXT CASE. * * NAMELIST The contents of each applicatble namelist is dumped for the * case in the input system of units. (Very useful to see * input parameter values). * SAVE Preserve the input data for the case to be used in the following * cases. This would be useful if multiple or comparison cases * are built. * DIM FT Sets the system units of measure. Options are FT, IN, M, CM * NEXT CASE Terminates the reading of input cards and begins execution of * the case. Case data are destroyed following execution of a case, * unless the SAVE card is present. * TRIM Trim calculations will be performed for each subsonic Mach number * within the case. A vehicle may be trimmed by deflecting a control * device on the wing or horizontal tail or by deflecting an * all-moveable horizontal stabilizer. * DAMP Provides dynamic-derivative results in addition to the standard * static-derivative output * CASEID Provides a case identification that is printed as part of the * output header. * DUMP ALL Prints the contents of the named arrays in the foot-pound-second * system. See Appendix C for list of arrays and their contents. * DUMP CASE will print all the arrays that are used during case * execution prior to the conventional output. (Not particularily * useful, as all data is in nasty arrays) * DUMP INPT will print dump of all input data blocks used for the case. * (also not useful) * DUMP IOM will print all the output arrays for the case. (not useful) * DUMP ALL will print all program arrays, even if not used for the case. * (surprisingly, not useful) * DERIV DEG Defines the output units of measure for the static and dynamic * stability derivatives, either RAD or DEG. The following parameters * are affected: CLa, Cma, Cyb, Cnb, Clb, CLq, Cmq, Clp, Cyp, Cnp, * Cnr, Clr, CLad, CMad. JSBSim XML output is also switched between * degrees and radians for alpha, beta, etc. * PART Provides auxiliary and partial outputs at each Mach number in the * case. These outputs are automatically provided for all cases at * transonic Mach numbers. * BUILD This command provides configuration build-up data. Conventional * static and dynamic stability data are output for a LOT of items. * PLOT Causes data generated by the program to be written to logical * unit 13, which can be retained for input to the Plot Module. * (Looks like it dumps the data arrays out in column format. Not * too useful). DIM M PART DERIV RAD *DUMP ALL DAMP ********************** * Flight Conditions * ********************** * WT Vehicle Weight * LOOP Program Looping Control * 1 = vary altitude and mach together, default) * 2 = vary Mach, at fixed altitude * 3 = vary altitude, at fixed Mach * NMACH Number of Mach numbers or velocities to be run, max of 20 * Note: This parameter, along with NALT, may affect the * proper setting of the LOOP control parameter. * MACH Array(20) Values of freestream Mach number * VINF Array(20) Values of freestream speed (unit: l/t) * NALPHA Number of angles of attack to be run, max of 20 * ALSCHD Array(20) Values of angles of attack, in ascending order * RNNUB Array(20) Reynolds number per unit length * Freestream Reynolds numbers. Each array element must * correspond to the respective Mach number/freestream * speed input, use LOOP=1.0 * NALT Number of atmospheric conditions to be run, max of 20 * input as either altitude or pressure and temperature * Note: This parameter, along with NMACH, may affect the * proper setting of the LOOP control parameter. * ALT Array(20) Values of geometric altitude * Number of altitude and values. Note, Atmospheric conditions * are input either as altitude or pressure and temperature. (MAX 20) * PINF Array(20) Values of freestream Static Pressure * TINF Array(20) Values of freestream Temperature * HYPERS =.true. Hypersonic analysis at all Mach numbers > 1.4 * STMACH Upper limit of Mach numbers for subsonic analysis * (0.6 3) $WGPLNF TYPE=1.0, CHRDR=6.80, CHRDTP=3.10, SSPN=17.37, SSPNE=15.87, SAVSI=4.27, CHSTAT=0.0, TWISTA=-3.0, DHDADI=3.6$ ********************************************** * Wing Sectional Characteristics Parameters * pg 39-40 ********************************************** * The section aerodynamic characteristics for these surfaces are * input using either the sectional characteristics namelists WGSCHR, * HTSCHR, VTSCHR and VFSCHR and/or the NACA control cards. Airfoil * characteristics are assummed constant for each panel of the planform. * * To avoid having to input all the airfoil sectional characteristics, * you can specify the NACA airfoil designation. Starts in Column 1. * * NACA x y zzzzzz * * where: * column 1-4 NACA * 5 any deliminator * 6 W, H, V, or F Planform for which the airfoil * designation applies: Wing, Horizontal * tail, Vertical tail, or Ventral fin. * 7 any deliminator * 8 1,4,5,6,S Type of airfoil section: 1-series, * 4-digit, 5-digit, 6-series, or Supersonic * 9 any deliminator * 10-80 Designation, columns are free format, blanks are ignored * * TOVC Maximum airfoil section thickness fraction of chord * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * DELTAY Difference between airfoil ordinates at 6% and 15% chord, * percent chord (% correct ???) * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * XOVC Chord location of maximum airfoil thickness, fraction of chord * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CLI Airfoil section design lift coefficient * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * ALPHAI Angle of attack at section design lift coefficient, deg * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CLALPA Airfoil section lift curve slope dCl/dAlpha, per deg (array 20) * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CLMAX Airfoil section maximum lift cofficient (array 20) * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CMO Section zero lift pitching moment coefficient * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * LERI Airfoil leading edge radius, fraction of chord * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * LERO RLE for outboard panel, fraction of chord * [Required input]. * Not required for straight tapered planforms. * CAMBER Cambered airfoil flag flag * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * TOVCO t/c for outboard panel * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * Not required for straight tapered planforms. * XOVCO (x/c)max for outboard panel * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * Not required for straight tapered planforms. * CMOT Cmo for outboard panel * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * Not required for straight tapered planforms. * CLMAXL Airfoil maximum lift coefficient at mach = 0.0 * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CLAMO Airfoil section lift curve slope at Mach=0.0, per deg * [Not required for subsonic speed regime. Required input * for transonic speed regime, user supplied or computed if * NACA card supplied] * TCEFF Planform effective thickness ratio, fraction of chord * [Not required for subsonic speed regime. Required input * for transonic speed regime, user supplied or computed if * NACA card supplied] * KSHARP Wave-drag factor for sharp-nosed airfoil section, not * input for round-nosed airfoils * [Not required for subsonic speed regime. Required input * for transonic speed regime, user supplied or computed if * NACA card supplied] * SLOPE Airfoil surface slope at 0,20,40,60,80 and 100% chord, deg. * Positive when the tangent intersects the chord plane forward * of the reference chord point * [Not required for subsonic speed regime. Required input * for transonic speed regime, user supplied or computed if * NACA card supplied] * ARCL Aspect ratio classification (see table 9, pg 41) * [Optional input] * XAC Section Aerodynamic Center, fraction of chord * [Optional input, computed by airfoil section module if airfoil * defined with NACA card or section coordinates] * DWASH Subsonic downwash method flag * = 1.0 use DATCOM method 1 * = 2.0 use DATCOM method 2 * = 3.0 use DATCOM method 3 * Supersonic, use DATCOM method 2 * [Optional input] * See figure 9 on page 41. * YCM Airfoil maximum camber, fraction of chord * [Required input, user supplied or computed by airfoil * section module if airfoil defined with NACA card or * section coordinates] * CLD Conical camber design lift coefficient for M=1.0 design * see NACA RM A55G19 (default to 0.0) * [Required input] * TYPEIN Type of airfoil section coordinates input for airfoil * section module * = 1.0 upper and lower surface coordinates (YUPPER and YLOWER) * = 2.0 Mean line and thickness distribution (MEAN and THICK) * [Optional input] * NPTS Number of section points input, max = 50.0 * [Optional input] * XCORD Abscissas of inputs points, TYPEIN=1.0 or 2.0, XCORD(1)=0.0 * XCORD(NPTS)= 1.0 required. * [Optional input] * YUPPER Ordinates of upper surface, TYPEIN=1.0, fraction of chord, and * requires YUPPER(1)=0.0 and YUPPER(NPTS)=0.0 * [Optional input] * YLOWER Ordinates of lower surface, TYPEIN=1.0, fraction of chord, * and requires YLOWER(1)=0.0 and YLOWER(NPTS)=0.0 * [Optional input] * MEAN Ordinates of mean line, TYPEIN=2.0, fraction of chord, and * requires MEAN(1)=0.0 and MEAN(NPTS)=0.0 * [Optional input] * THICK Thickness distribution, TYPEIN=2.0, fraction of chord, and * requires THICK(1)=0.0 and THICK(NPTS)=0.0 * [Optional input] * *NACA W 5 23014 * Göttingen 436 airfoil. The S.23 had a modified Gö 436 airfoil. $WGSCHR TYPEIN=1.0, DWASH=1.0, NPTS=29.0, XCORD = 0.000000000, 0.000022952, 0.000070284, 0.000162514, 0.000334249, 0.000624353, 0.001056325, 0.002176391, 0.005038420, 0.009613580, 0.012500000, 0.015660360, 0.019609120, 0.025000000, 0.032253180, 0.050000000, 0.075000000, 0.100000000, 0.150000000, 0.200000000, 0.300000000, 0.400000000, 0.500000000, 0.600000000, 0.700000000, 0.800000000, 0.900000000, 0.950000000, 1.000000000, YUPPER = 0.000000000, 0.001267582, 0.001931812, 0.002638020, 0.003427223, 0.004345490, 0.005417153, 0.007736135, 0.012362700, 0.018755210, 0.022310000, 0.025631780, 0.028950410, 0.032630000, 0.036917100, 0.046250000, 0.057870000, 0.066500000, 0.079250000, 0.087500000, 0.092500000, 0.089500000, 0.083000000, 0.072000000, 0.058500000, 0.041000000, 0.022000000, 0.011250000, 0.000000000, YLOWER = 0.000000000,-0.001254635,-0.001890570,-0.002540153,-0.003221198, -0.003953656,-0.004746233,-0.006344466,-0.009132541,-0.012681680, -0.014690000,-0.016597310,-0.018336890,-0.019870000,-0.021155550, -0.022750000,-0.022830000,-0.022500000,-0.021250000,-0.020000000, -0.017500000,-0.015000000,-0.012500000,-0.010000000,-0.007500000, -0.005000000,-0.002500000,-0.001250000, 0.000000000, $ ****************************** * Ground effects parameters ****************************** * NGH Number of ground heights to be run, maximum of 10. * GRDHT Values of ground heights. Ground heights equal altitude * of reference plane relative to ground. Ground effect output * may be obtained at a maximum of ten different ground heights. * According to the DATCOM, the ground effects become neglible * when the ground height exceeds the wing span. Through * testing, there is a minimal effect up to twice the wing * span, so to keep our tables smooth, let's make the last * point 1.5b, and the output adds a point at 2b of 0.0. The * smallest value should NOT be 0.0, which would be the wing * sitting on the ground. It should be the height of the wing * with the aircraft sitting on the ground. * * $GRNDEF NGH=10.0, GRDHT=5.0,10.0,15.0,20.0,25.0,30.0,40.0,50.0,60.0,77.55$ SAVE ****************************************** * Symetrical Flap Deflection parameters ****************************************** * DATCOM pg 47 states : * * "In general, the eight flap types defined using SYMFLP * (variable FTYPE) are assumed to be located on the most * aft lifting surface, either horizontal tail or wing if * a horizontal tail is not defined." * * FTYPE Flap type * 1.0 Plain flaps * 2.0 Single slotted flaps * 3.0 Fowler flaps * 4.0 Double slotted flaps * 5.0 Split flaps * 6.0 Leading edge flap * 7.0 Leading edge slats * 8.0 Krueger * NDELTA Number of flap or slat deflection angles, max of 9 * * DELTA Flap deflection angles measured streamwise * (NDELTA values in array) * PHETE Tangent of airfoil trailine edge angle based on ordinates at * 90 and 99 percent chord * PHETEP Tangent of airfoil trailing edge angle based on ordinates at * 95 and 99 percent chord * CHRDFI Flap chord at inboard end of flap, measured parallel to * longitudinal axis * CHRDFO Flap chord at outboard end of flap, measured parallel to * longitudinal axis * SPANFI Span location of inboard end of flap, measured perpendicular * to vertical plane of symmetry * SPANFO Span location of outboard end of flap, measured perpendicular * to vertical plane of symmetry * CPRMEI Total wing chord at inboard end of flap (translating devices * only) measured parallel to longitudinal axis * (NDELTA values in array) * Single-slotted, Fowler, Double-slotted, leading-edge * slats, Krueger flap, jet flap * CPRMEO Total wing chord at outboard end of flap (translating devices * only) measured parallel to longitudinal axis * (NDELTA values in array) * Single-slotted, Fowler, Double-slotted, leading-edge * slats, Krueger flap, jet flap * CAPINS (double-slotted flaps only) * CAPOUT (double-slotted flaps only) * DOSDEF (double-slotted flaps only) * DOBCIN (double-slotted flaps only) * DOBCOT (double-slotted flaps only) * SCLD Increment in section lift coefficient due to * deflecting flap to angle DELTA[i] (optional) * (NDELTA values in array) * SCMD Increment in section pitching moment coefficient due to * deflecting flap to angle DELTA[i] (optional) * (NDELTA values in array) * CB Average chord of the balance (plain flaps only) * TC Average thickness of the control at hinge line * (plain flaps only) * NTYPE Type of nose * 1.0 Round nose flap * 2.0 Elliptic nose flap * 3.0 Sharp nose flap * JETFLP Type of flap * 1.0 Pure jet flap * 2.0 IBF * 3.0 EBF * CMU Two-dimensional jet efflux coefficient * DELJET Jet deflection angle * (NDELTA values in array) * EFFJET EBF Effective jet deflection angle * (NDELTA values in array) * $SYMFLP FTYPE=2.0, NDELTA=9.0, * DELTA(1)=0.0,5.0,10.0,15.0,20.0,25.0,30.0,35.0,40.0, * PHETE=0.0522, PHETEP=0.0391, * CHRDFI=2.0, CHRDFO=1.6, * SPANFI=5.78, SPANFO=15.3, * CPRMEI(1)=8.1,8.1,8.2,8.2,8.3,8.3,8.3,8.4,8.4, * CPRMEO(1)=3.7,3.7,3.8,3.8,3.9,3.9,3.9,4.0,4.0, * NTYPE=1.0$ * At this point, we are going to terminate the case so that we can get * the flap effects. We can't save this data, as we are * also going to do control surfaces on the horizontal tail. CASEID FLAPS: Short S-23 Aircraft NEXT CASE ************************************************************* * Asymmetrical Control Deflection parameters : Ailerons ************************************************************* * STYPE Type * 1.0 Flap spoiler on wing * 2.0 Plug spoiler on wing * 3.0 Spoiler-slot-deflection on wing * 4.0 Plain flap aileron * 5.0 Differentially deflected all moveable horizontal tail * NDELTA Number of control deflection angles, required for all controls, * max of 9 * DELTAL Defelction angle for left hand plain flap aileron or left * hand panel all moveable horizontal tail, measured in * vertical plane of symmetry * DELTAR Defelction angle for right hand plain flap aileron or right * hand panel all moveable horizontal tail, measured in * vertical plane of symmetry * SPANFI Span location of inboard end of flap or spoiler control, * measured perpendicular to vertical plane of symmetry * SPANFO Span location of outboard end of flap or spoiler control, * measured perpendicular to vertical plane of symmetry * PHETE Tangent of airfoil trailing edge angle based on ordinates * at x/c - 0.90 and 0.99 * CHRDFI Aileron chord at inboard end of plain flap aileron, * measured parallel to longitudinal axis * CHRDFO Aileron chord at outboard end of plain flap aileron, * measured parallel to longitudinal axis * DELTAD Projected height of deflector, spoiler-slot-deflector * control, fraction of chord * DELTAS Projected height of spoiler, flap spoiler, plug spoiler and * spoiler-slot-deflector control; fraction of chord * XSOC Distance from wing leading edge to spoiler lip measured * parallel to streamwise wng chord, flap and plug spoilers, * fraction of chord * XSPRME Distance from wing leading edge to spoiler hinge line * measured parallel to streamwise chord, flap spoiler, * plug spoiler and spoiler-slot-deflector control, fraction * of chord * HSOC Projected height of spoiler measured from and normal to * airfoil mean line, flap spoiler, plug spoiler and spoiler- * slot-reflector, fraction of chord $ASYFLP STYPE=4.0, NDELTA=9.0, DELTAL(1)=-32.0,-20.0,-10.0,-5.0,0.0,5.0,10.0,20.0,32.0, DELTAR(1)=32.0,20.0,10.0,5.0,0.0,-5.0,-10.0,-20.0,-32.0, SPANFI=9.66, SPANFO=15.86, CHRDFI=1.18, CHRDFO=0.50, * Note: PHETE is from the Citation. I don't understand the definition. PHETE=0.05228$ * Terminates the reading of input cards and begins execution of * the case. Case data are destroyed following execution of a case, * unless the SAVE card is present. CASEID AILERONS: Short S-23 Aircraft SAVE NEXT CASE ************************************************* * Horizontal Tail Sectional Characteristics pg 39-40 ************************************************* * Same build up as wing, if you'd like to use that instead. * RAF 30 symmetric airfoil. The S.23 hstab. had a modified RAF 30 airfoil. * "The thickness to chord ratio was increased from the standard 12.64% to 13.75%" $HTSCHR TYPEIN=1.0, NPTS=17.0, XCORD = 0.0000000, 0.0125000, 0.0250000, 0.0500000, 0.1000000, 0.1500000, 0.2000000, 0.3000000, 0.4000000, 0.5000000, 0.6000000, 0.7000000, 0.8000000, 0.9000000, 0.9500000, 1.0000000, YUPPER = 0.0000000, 0.0180000, 0.0248000, 0.0346000, 0.0468000, 0.0544000, 0.0594000, 0.0632000, 0.0620000, 0.0566000, 0.0478000, 0.0370000, 0.0250000, 0.0130000, 0.0070000, 0.0000000, YLOWER = 0.0000000, -.0180000, -.0248000, -.0346000, -.0468000, -.0544000, -.0594000, -.0632000, -.0620000, -.0566000, -.0478000, -.0370000, -.0250000, -.0130000, -.0070000, 0.0000000$ ********************************************* * Horizontal Tail planform variables pg 37-38 ********************************************* * CHRDTP Tip chord * SSPNOP Semi-span outboard panel. Not required for straight * tapered planform. * SSPNE Semi-span exposed panel * SSPN Semi-span theoretical panel from theoretical root chord * CHRDBP Chord at breakpoint * CHRDR Chord root * SAVSI Inboard panel sweep angle * CHSTAT Reference chord station for inboard and outboard panel * sweep angles, fraction of chord * TWISTA Twist angle, negative leading edge rotated down (from * exposed root to tip) * SSPNDD Semi-span of outboard panel with dihedral * DHDADI Dihedral angle of inboard panel * DHDADO Dihedral angle of outboard panel. If DHDADI=DHDADO only * input DHDADI * TYPE 1.0 - Straight tapered planform * 2.0 - Double delta planform (aspect ratio <= 3) * 3.0 - Cranked planform (aspect ratio > 3) * SHB Portion of fuselage side area that lies between Mach lines * originating from leading and trailing edges of horizontal * tail exposed root chord (array 20). * Only required for supersonic and hypersonic speed regimes. * SEXT Portion of extended fueslage side area that lies between * Mach lines originating from leading and trailing edges of * horizontal tail exposed root chord (array 20) * Only required for supersonic and hypersonic speed regimes. * RLPH Longitudinal distance between CG and centroid of Sh(B) * positive aft of CG * Only required for supersonic and hypersonic speed regimes. $HTPLNF TYPE=1.0, CHRDR=2.92, CHRDTP=1.44, SSPN=4.83, SSPNE=4.26, SAVSI=11.47, CHSTAT=0.0, TWISTA=0.0, DHDADI=0.0$ ****************************************** * Vertical Tail planform variables pg 37-38 ****************************************** * CHRDTP Tip chord * SSPNOP Semi-span outboard panel * SSPNE Semi-span exposed panel * SSPN Semi-span theoretical panel from theoretical root chord * CHRDBP Chord at breakpoint * CHRDR Chord root * SAVSI Inboard panel sweep angle * SAVSO Outboard panel sweep angle * CHSTAT Reference chord station for inboard and outboard panel * sweep angles, fraction of chord * TYPE 1.0 - Straight tapered planform * 2.0 - Double delta planform (aspect ratio <= 3) * 3.0 - Cranked planform (aspect ratio > 3) * SVWB Portion of exposed vertical panel area that lies between * Mach lines emanating from leading and trailing edges of * wing exposed root chord (array 20) * Only required for supersonic and hypersonic speed regimes. * SVB Area of exposed vertical panel not influenced by wing or * horizontal tail (array 20) * Only required for supersonic and hypersonic speed regimes. * SVHB Portion of exposed vertical panel area that lies between Mach * lines emanating from leading and and trailing edges of * horizontal tail exposed root chord (array 20) * Only required for supersonic and hypersonic speed regimes. $VTPLNF TYPE=1.0, CHRDR=5.50, CHRDTP=2.10, SSPN=7.80, SSPNE=4.07, SAVSI=15.75, CHSTAT=0.0$ *********************************** * Elevator Deflection parameters *********************************** * FTYPE Flap type * 1.0 Plain flaps * 2.0 Single slotted flaps * 3.0 Fowler flaps * 4.0 Double slotted flaps * 5.0 Split flaps * 6.0 Leading edge flap * 7.0 Leading edge slats * 8.0 Krueger * NDELTA Number of flap or slat deflection angles, max of 9 * DELTA Flap deflection angles measured streamwise * (NDELTA values in array) * PHETE Tangent of airfoil trailine edge angle based on ordinates at * 90 and 99 percent chord * PHETEP Tangent of airfoil trailing edge angle based on ordinates at * 95 and 99 percent chord * CHRDFI Flap chord at inboard end of flap, measured parallel to * longitudinal axis * CHRDFO Flap chord at outboard end of flap, measured parallel to * longitudinal axis * SPANFI Span location of inboard end of flap, measured perpendicular * to vertical plane of symmetry * SPANFO Span location of outboard end of flap, measured perpendicular * to vertical plane of symmetry * CPRMEI Total wing chord at inboard end of flap (translating devices * only) measured parallel to longitudinal axis * (NDELTA values in array) * Single-slotted, Fowler, Double-slotted, leading-edge * slats, Krueger flap, jet flap * CPRMEO Total wing chord at outboard end of flap (translating devices * only) measured parallel to longitudinal axis * (NDELTA values in array) * Single-slotted, Fowler, Double-slotted, leading-edge * slats, Krueger flap, jet flap * CAPINS (double-slotted flaps only) (NDELTA values in array) * CAPOUT (double-slotted flaps only) (NDELTA values in array) * DOSDEF (double-slotted flaps only) (NDELTA values in array) * DOBCIN (double-slotted flaps only) * DOBCOT (double-slotted flaps only) * SCLD Increment in section lift coefficient due to * deflecting flap to angle DELTA[i] (optional) * (NDELTA values in array) * SCMD Increment in section pitching moment coefficient due to * deflecting flap to angle DELTA[i] (optional) * (NDELTA values in array) * CB Average chord of the balance (plain flaps only) * TC Average thickness of the control at hinge line * (plain flaps only) * NTYPE Type of nose * 1.0 Round nose flap * 2.0 Elliptic nose flap * 3.0 Sharp nose flap * JETFLP Type of flap * 1.0 Pure jet flap * 2.0 IBF * 3.0 EBF * CMU Two-dimensional jet efflux coefficient * DELJET Jet deflection angle * (NDELTA values in array) * EFFJET EBF Effective jet deflection angle * (NDELTA values in array) $SYMFLP FTYPE=1.0, NTYPE=1.0, NDELTA=9.0, DELTA(1)=-20.0,-15.0,-10.0,-5.0,0.0,5.0,10.0,15.0,20.0, SPANFI=0.58, SPANFO=4.83, CHRDFI=1.22, CHRDFO=0.61, * Note: PHETE and PHETEP are from the Citation. I don't understand the definition. PHETE=0.0522, PHETEP=0.0523, CB=0.1, TC=0.15$ ************************************** * Vertical Tail Sectional Characteristics pg 39-40 ************************************** * Same build up as wing, if you'd like to use that instead. * RAF 30 symmetric airfoil. $VTSCHR TYPEIN=1.0, NPTS=17.0, XCORD = 0.0000000, 0.0125000, 0.0250000, 0.0500000, 0.1000000, 0.1500000, 0.2000000, 0.3000000, 0.4000000, 0.5000000, 0.6000000, 0.7000000, 0.8000000, 0.9000000, 0.9500000, 1.0000000, YUPPER = 0.0000000, 0.0180000, 0.0248000, 0.0346000, 0.0468000, 0.0544000, 0.0594000, 0.0632000, 0.0620000, 0.0566000, 0.0478000, 0.0370000, 0.0250000, 0.0130000, 0.0070000, 0.0000000, YLOWER = 0.0000000, -.0180000, -.0248000, -.0346000, -.0468000, -.0544000, -.0594000, -.0632000, -.0620000, -.0566000, -.0478000, -.0370000, -.0250000, -.0130000, -.0070000, 0.0000000$ ***************************** * Control Tabs parameters DATCOM User's manual, pg 68/69 ***************************** * TTYPE Type * 1.0 Tab Control * 2.0 Trim Tab * 3.0 Both * CFITC Control tab inboard chord * CFOTC Control tab outboard chord * BITC Control tab inboard span location * BOTC Control tab outboard span location * CFITT Trim tab inboard chord * CFOTT Trim tab outboard chord * BITT Trim tab inboard span location * BOTT Trim tab outboard span location * B1 Ch(delta) for control tab * B2 Ch(delta) for trim tab * B3 Ch(alpha) for control tab * B4 Ch(alpha) for trim tab * D1 see table 11, pg 70-71 * D2 see table 11, pg 70-71 * D3 see table 11, pg 70-71 * GCMAX Maximum Stick gearing * KS Tab spring effectiveness * RL Aerodynamic boost link ratio (RL > 0). * To input RL infinity, set RL < 0 * BGR Control tab gear ratio * DELR -DELTAtc(max)/DELTAc(max) for a maximum control deflection * The value of DELR is positive because DELTAtc(max) and * DELTAc(max) will have opposite signs. * When R(L) 0, DELR = 1.0 * I'll skip propulsion effects for now. CASEID TOTAL: Short S-23 Aircraft