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We Provide Best Services. App Store Preview. Screenshots iPhone iPad. Dec 16, Version Minor bug fixes and improvements. Ratings and Reviews. App Privacy. Information Seller American Automobile Association. Size MB. Category Lifestyle. Compatibility iPhone Requires iOS Languages English. Price Free. App Support. The purpose of this module is to present a Class II method for estimating aircraft component weights.
These methods employ empirical equations, which relate component weights to aircraft design characteristics. In the Class II weight estimate module, the weight estimation methods are identified as follows:.
Because aircraft component weight modeling in the software is a function of take-off weight, the Class II weight estimation module includes a weight iteration calculation which iterates between the take-off weights and component weights to converge on a solution.
Component Center of Gravity option allows the user to calculate the aircraft center of gravity by entering weight components and their locations in a preformatted table. Tables are available for structure weight, fixed equipment weight, powerplant weight and total weight. Center of gravity of wing, canard, horizontal tail, vertical tail, V-tail, fuselage, tailbooms, nacelles, stores, floats and ventral fins are calculated.
These are based on geometry of the components and on the methods described in Chapter 8 of Airplane Design Part V. Detailed center of gravity calculations based on empty weight and various loading scenarios are included in the software.
These calculations include the determination of the most forward and most aft center of gravity based on the minimum and maximum weights of passengers, fuel, baggage, cargo etc. The software includes the option to plot the C. The plot feature can also be used to plot the empty weight C. The Lift submodule can be used for estimating the lifting characteristics of aircraft lifting surfaces and high lift devices.
Maximum lift coefficients for airfoils and wing, horizontal tail, vertical tail, V-tail and canard are calculated. The increase in the maximum lift coefficients caused by multiple high lift device segments is calculated. The user can define multiple unique segments of high lift devices on the wing. The user can select from the following list of high lift devices.
The size of the trailing edge flap segment needed to meet the desired maximum lift coefficient for take-off and landing conditions can be calculated.
The plotting feature can be used to plot the high lift device maximum lift coefficient as a function of flap deflection and flap chord ratio. The methodology used to determine the lifting surface spanwise lift distribution is based on lifting line theory. Lift as a function of angle of attack, elevator canardvator, ruddervator and horizontal tail canard, V-tail incidence can be calculated. The lift coefficient at zero-angle-of-attack can be used to determine the lift coefficients, downwash angle and upwash angle at zero aircraft angle of attack and the zero lift angle of attack.
This module includes flap effects. The Class I Drag submodule can be used for a first estimate of the aircraft drag. The Class I Drag module has the following seven options:. The program will check to see if the aircraft has retractable or fixed landing gear. The methodology used to calculate the drag polar can be found in Chapter 3 of Airplane Design Part I. The Class I Drag module relates the total aircraft lift coefficient to the total aircraft drag coefficient by a parabolic drag polar equation.
The purpose of the Class II Drag submodule is to supply a Class II method for predicting drag polars of aircraft during the preliminary design phase. A detailed drag polar can be calculated for the subsonic, transonic and supersonic flow regimes.
The following drag calculations can be performed:. The program will account for fuselage base area change as a function of inoperative engines. Changes in nacelle drag are also accounted for due to inoperative engines. Dihedral is accounted for in the wetted areas of lifting surfaces. Drag can also be plotted as a function of angle of attack for an untrimmed aircraft. The drag polar can be approximated with a trendline. The Moment submodule calculates the spanwise moment distribution on the wing, horizontal tail, vertical tail and canard.
The ground effects on aircraft moment are also determined. Moment as function of angle of attack, elevator canardvator and horizontal tail canard incidence can be calculated. This pitching moment at zero-angle-of-attack submodule can be used to determine the zero lift aircraft pitching moment coefficients and pitching moment coefficients at zero aircraft angle of attack.
To account for each aircraft component, the pitching moment coefficient components are calculated separately. The user can estimate the zero lift pitching moment coefficient, the effect of trailing and leading edge flaps on lift and pitching moment coefficients. The Aerodynamic Center submodule can be used to calculate aerodynamic center locations of individual aircraft components and the aerodynamic center shift due to components.
The Dynamic Pressure Ratio module allows the user to predict the horizontal tail, V-tail and vertical tail dynamic pressure ratio as a variation of angle of attack.
The wake effects are also accounted for in these calculations. The purpose of the Power Effects module is to calculate the effects of the operating propeller on aerodynamic properties of the aircraft.
The effects of power are calculated for the following parameters:. The Pitch Break module uses the wing aspect ratio and wing quarter chord sweep angle to determine whether the pitch break is stable or unstable based on Figure 8.
The purpose of the Performance Sizing module is to allow for a rapid estimation of those aircraft design parameters which have a major impact on aircraft performance. Aircraft are usually required to meet performance objectives in different categories depending on the mission profile.
Meeting these performance objectives normally results in the determination of a range of values for:. The variables listed above are plotted in the form of a performance-matching plot. These plots depend on the type of aircraft, the applicable specification and the applicable regulation s.
With the help of such a plot, the combination of the highest possible wing loading and the smallest possible thrust or highest power loading, which meets all performance requirements, can be determined. The methodology used for performance sizing can be found in Airplane Design Part I.
The purpose of the Analysis submodule is to provide the user with Class II analysis methods for predicting the performance characteristics of an aircraft. The purpose of the geometry module is to help the user define the geometry of the fuselage, wing, horizontal tail, vertical tail, V-tail and canard and calculate related parameters. The methodology used to calculate the aircraft component geometries is described in Airplane Design Part II.
Two-dimensional geometry can be defined for straight and cranked lifting surfaces. Get a TripTik. How can I get a TripTik? Launch TripTik. At a AAA branch. Find a branch. On the Auto Club App. Download the app. Digital TourBook guides. View TourBooks.
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