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Current engine condition monitoring (ECM) systems for jet engines include the analysis of on-wing gas path data using steady-state performance models. Such data, which are also referred to as performance snapshots, usually are taken during cruise flight and during takeoff. Using steady-state analysis, it is assumed that these snapshots have been taken under stabilized operating conditions. However, this assumption is reasonable only for cruise snapshots. During takeoff, jet engines operate in highly transient conditions with significant heat transfer occurring between the fluid and the engine structure. Hence, steady-state analysis of takeoff snapshots is subject to high uncertainty. Because of this, takeoff snapshots are not used for performance analysis in current ECM systems. We quantify the analysis uncertainty by transient simulation of a generic takeoff maneuver using a performance model of a medium size two-shaft turbofan engine with high bypass ratio. Taking into account the influence of the preceding operating regimes on the transient heat transfer effects, this takeoff maneuver is extended backward in time to cover the aircraft turnaround as well as the end of the last flight mission. We present a hybrid approach for thermal calculation of both the fired engine and the shutdown engine. The simulation results show that takeoff derate, ambient temperature, taxi-out (XO) duration and the duration of the preceding aircraft turnaround have a major influence on the transient effects occurring during takeoff. The analysis uncertainty caused by the transient effects is significant. Based on the simulation results, we propose a method for correction of takeoff snapshots to steady-state operating conditions. Furthermore, we show that the simultaneous analysis of cruise and corrected takeoff snapshots leads to significant improvements in observability. [ABSTRACT FROM AUTHOR]

This paper addresses the unsteady formation of secondary flow structures inside a turbine rotor passage. The first stage of a two-stage, low-pressure turbine is investigated at a Reynolds Number of 75,000. The design represents the third and the fourth stages of an engine-representative, low-pressure turbine. The flow field inside the rotor passage is discussed in the relative frame of reference using the streamwise vorticity. A multistage unsteady Reynolds-averaged Navier-Stokes (URANS) prediction provides the time-resolved data set required. It is supported by steady and unsteady area traverse data acquired with five-hole probes and dual-film probes at rotor inlet and exit. The unsteady analysis reveals a nonclassical secondary flow field inside the rotor passage of this turbine. The secondary flow field is dominated by flow structures related to the upstream nozzle guide vane. The interaction processes at hub and casing appear to be mirror images and have characteristic forms in time and space. Distinct loss zones are identified, which are associated with vane-rotor interaction processes. The distribution of the measured isentropic stage efficiency at rotor exit is shown, which is reduced significantly by the secondary flow structures discussed. Their impacts on the steady as well as on the unsteady angle characteristics at rotor exit are presented to address the influences on the inlet conditions of the downstream nozzle guide vane. It is concluded that URANS should improve the optimization of rotor geometry and rotor loss can be controlled, to a degree, by nozzle guide vane (NGV) design. [ABSTRACT FROM AUTHOR]

A mathematical model for the prediction of the maximum speed of a high pressure turbine following a shaft failure event was developed. The model predicts the high pressure compressor and ducting system pre- and poststall behaviors such as rotating stall and surge after the shaft breakage. The corresponding time-dependent high pressure turbine inlet conditions are used to calculate the turbine maximum speed, taking into account friction and blade and vane tip clearance variations as a result of the rearward movement of the turbine and destruction of the turbine blading. The compressor and ducting system is modeled by a one-dimensional, stage-by-stage approach. The approach uses a finite-difference numerical technique to solve the nonlinear system of equations for continuity, momentum, and energy including source terms for the compressible flow through inlet ducting, compressor, and combustor. The compressor blade forces and shaft work are provided by a set of quasisteady state stage characteristics being valid for prestall and poststall operations. The maximum turbine speed is calculated from a thermodynamic turbine stand-alone model, derived from a performance synthesis program. Friction and blade and vane tip clearance variations are determined iteratively from graphical data depending on the axial rearward movement of the turbine. The compressor and ducting system model was validated in prestall and poststall operation modes with measured high pressure compressor data of a modern two-shaft engine. The turbine model was validated with measured intermediate pressure shaft failure data of a three-shaft engine. The shaft failure model was applied on a modern two-shaft engine. The model was used to carry out a sensitivity study to demonstrate the impact of control system reactions on the resulting maximum high pressure turbine speed following a shaft failure event. [ABSTRACT FROM AUTHOR]