In the modern world of railways, electrification is given great importance. Currently, more than 50% of all freight traffic carried out by electric traction. However, today, about half of the railways are not electrified, so it is necessary to use thermal engines to drive the locomotive. One of the possible variant is use gas turbine unit. The power of the gas turbine unit is given to the electric generator, and the electric motors drive the locomotive. In the present paper, as a power plant of a gas turbine locomotive, considered gas turbine unit with a twin -shaft gas generator of two schemes: 1- with steam supply to the inlet of the high-pressure turbine (into the combustion chamber) and 2- with steam supply to inlet of the free turbine. By CAE system of ASTRA, Collaboration operation lines calculated for different variants of steam injection. When the steam injected into the inlet of a free turbine and a high-pressure turbine. in the case of steam supply to the input of the free turbine and the high-pressure turbine there is a significant shift in Collaboration operation lines, which can lead to a decrease in the gas-dynamical stability of the compressors, and efficiency. To maintain the position of Collaboration operation lines, was applied the correction of the throughput capacity of free turbine nozzle vanes (by 15%). In the case of steam supply to the inlet of a free turbine, to ensure gas-dynamic stability of the compressors, a change in the throughput capacity of its nozzle vanes is required.

Continuous improvement of fuel efficiency of aircraft engines is the main global trend in modern engine construction. To date, aviation gas turbine engines have reached a high degree of thermodynamic and design-and technology perfection. One of the promising ways to further improve their fuel efficiency is the use of complex thermodynamic cycles with turbine exhaust heat regeneration and with intermediate cooling in the process of air compression. Until recently, the use of cycles with a recuperative heat exchanger and an intercooler in aircraft gas turbine engines was restrained by a significant increase in the mass of the power plant due to the installation of heat exchangers. Currently, it has become technologically possible to create compact, light, high-efficiency heat exchangers for use on aircraft without compromising their performance. An important target in the design of engines with heat recovery is to select the parameters of the working process that provide maximum efficiency of the aircraft system. The article focuses on the statement of the task of optimization and choice of rational parameters of the working process of a bypass three-shaft turbojet engine with an intercooler and a recuperative heat exchanger. On the basis of the developed method multi-criteria optimization was carried out by means of numerical simulations. The results of optimization of thermodynamic cycle parameters of a bypass three-shaft turbojet engine with an intercooler and a recuperative heat exchanger in the aircraft system according to such criteria as the total weight of the engine and fuel required for the flight, and the aircraft specific fuel consumption per ton - kilometer of the payload are presented. A passenger aircraft of the Airbus A310-300 type was selected. The developed mathematical model for calculating the mass of a compact heat exchanger, designed to solve optimization problems at the stage of conceptual design of the engine is presented. The developed methods and models are implemented in the ASTRA program. The possibility of improving the efficiency of turbofan engines due to the use of complex thermodynamic cycles is shown.

With the tightening of international environmental requirements for civil aviation and the implementation of initiatives aimed at reducing specific greenhouse gas emissions, the transition to hybrid power plants for regional aircraft is becoming increasingly relevant. This paper proposes an approach to the integrated energy assessment of a parallel hybrid turboprop power plant at the conceptual design stage while taking the aircraft mission profile into account. The considered power plant includes a gas turbine engine, a reversible electric machine located on the same shaft as the reduction gearbox and propeller, an electrical energy storage system, and power electronics. The mission profile is represented as a sequence of segments—takeoff, climb, cruise, descent, and approach/landing. For each segment, energy balances are formulated and allowable operating ranges for the gas turbine and electric subsystems are defined. The key parameter is the hybridization factor, which determines the share of power transmitted to the propeller from the electric machine in different mission segments. The primary integrated performance metric is the energy consumption per ton-kilometer of transported payload. The analysis shows that for ranges up to 500 km, the hybrid configuration reduces specific energy consumption per ton-kilometer by up to 9%.

The benefit of exhaust heat recuperation in a micro turbojet depends on the fuel saved at the required thrust and on recuperator mass and pressure losses. A 1000 N engine with a semicircular-channel matrix is analysed in ASTRA at constant static thrust for temperature recovery ratios θ = 0–0.7. Compressor and turbine characteristics are interpolated, scaled and extrapolated during component matching. Original, half and zero pressure-loss cases retain the same matrix mass. With the original losses, the lowest calculated engine-plus-fuel mass for one hour occurs at θ = 0.3, but the compressor stability margin of 5.856% is below the adopted 10% threshold. Among points meeting this thresh-old, θ = 0.2 leaves 1.565 kg for additional installation mass and requires a turbine inlet temperature (TIT) of 1234.38 K, 65.71 K above the reference. For one hour with 90% nominal operation and no fuel-flow difference during the remaining time, the allowance falls to 0.962 kg. Half-loss and ze-ro-loss cases give one-hour allowances of 4.198 and 7.716 kg. These sensitivities define targets for recuperator design; the mass, thermal strength and life of the complete installed assembly remain to be established.