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It is typical in aircraft engine design to explore new configurations in a constant effort to achieve greater efficiency with respect to various considerations. An integral component of this process requires a complete and robust simulation of rotor dynamics. Tuning the design with results of rotor dynamics simulations can be made possible with a tool that has adequate modeling techniques to capture the physics associated with engine behavior under various operating conditions accurately.

An important part of future air vehicle design will be the development of a transient integrated aircraft system model. DC electric power system modeling poses particular challenges because they are highly dynamic and employ short time constant line replaceable units [1, 2, 3]. This paper describes an approach to modeling an aircraft's electric power system that uses simplified non-physics based models of the line replaceable units that are part of future 270VDC aircraft power systems. The model is an alternative to physics based models and is particularly useful for the initial phases of aircraft development before hardware development has occurred. A 270VDC aircraft power system model is constructed as an example using the unit models. Selected results will be presented.

The optimal integration of vehicle subsystems is of critical importance in the design of future energy efficient fighter aircraft. The INVENT (INtegrated Vehicle ENergy Technology) program has been dedicated to this endeavor through modeling/simulation of thermal management, power generation & distribution, & actuation subsystems. Achieving dual cooling & power generation capability from a single subsystem would be consistent with current efforts in system integration optimization. In this paper, we present a reconfiguration of an archetypal closed-loop air cycle system for a modern fighter as an open-loop gas generator cycle operating interchangeably between refrigeration and auxiliary power modes. A numerical model was developed within NPSS to assess maximum power extraction capabilities of a system originally designed for cooling purposes under different operating conditions.

Current industry trends demonstrate aircraft electrification will be part of future platforms in order to achieve higher levels of efficiency in various vehicle level sub-systems. However, electrification requires a substantial change in aircraft design that is not suitable for re-winged or re-engined applications as some aircraft manufacturers are opting for today. Thermal limits arise as engine cores progressively get smaller and hotter to improve overall engine efficiency, while legacy systems still demand a substantial amount of pneumatic, hydraulic and electric power extraction. The environmental control system (ECS) provides pressurization, ventilation and air conditioning in commercial aircraft, making it the main heat sink for all aircraft loads with exception of the engine fuel thermal management system.

The study of Supercooled Large Droplets (SLD) has received greater attention in the Aviation industry since the ATR-72 accident in 1994, which was attributed to SLD. This type of icing cloud usually consists of droplets of up to a millimeter in diameter and mean volumetric diameter (MVD) greater than 40 microns1. The analyses of the ice accretion process with SLD have focused mainly on the wing and stabilizers, particularly on the leading edges where accretion can occur beyond the ice protected areas. There are several numerical and empirical models to predict the mass and shapes of ice accreted from SLD, but there are few published papers that focus on SLD accretion within aircraft turbofan engines2, 3, 4, 5, 6, 7, 8, 9. SLD droplets have higher inertia than conventional icing droplets, which leads to their trajectories being less influenced by the aerodynamic forces. However, large droplets are more likely to breakup than smaller droplets when subjected to highly shear flows.