# Feasibility of a Helium Closed-Cycle Gas Turbine for UAV Propulsion
## 4 Discussion  
### System Performance  
The helium closed-cycle gas turbine looks suitable for UAV missions. It is compact, efficient and offers high reliability since it requires only one rotating equipment (turbo set). The choice of helium as a working fluid enhances the overall system design, and the operating envelope does not adversely affect the working fluid. The top-level requirement of high payload, high altitude, long-endurance, and range are representative of the mission requirements which could be reasonably expected but sufficiently stringent to test the real application of the closed-cycle gas turbine system for UAV propulsion. The system mass flow rate, temperature, and pressure, as well as component characteristic, were simulated for targeted thrust and system performance based on pressure ratio, turbine entry temperature (TET), and altitude variation. The choice of mass flow has a direct relationship with the thrust, the power required, weight and system size. In the parametric assessment, a fixed mass flow of 50 kg/s was chosen as a representative for a thrust and payload requirement. The mass flow value can be optimised in relationship with TET to obtain a lightweight UAV. The proposal to use a single-shaft arrangement also reduces the complexity of weight. The payload estimate refers to equipment for which the UAV provides platform and transportation. In this context, the payload is a measure of the size, weight, and power available to perform take-off, fly around and landing. Figure 3 shows the strong relationship between specific fuel consumption and the TET and pressure ratio. At take-off, a simultaneous increase in the TET from 700 to 950 \({}^{\circ}\)C, and pressure ratio both decreases the SFC requirement. However, there is a limit for an optimum UAV mission requirement considering the implication of size for pressure ratio and the cooling need for both turbine and the precooler heat exchanger. A temperature between 700 and 800 \({}^{\circ}\)C seems to be conservative to high payload mission as the TET have a direct effect on the power required and thrust. The parametric study shows the system performance peak pressure ratio of 4 and a conservative TET of 800 \({}^{\circ}\)C was chosen as the final baseline for the analysis. A pressure ratio of 4 was selected for the gas turbine to keep the turbo set weight and size within reasonable values. An increase in power density for the helium closed-cycle gas turbine provides a small system size and less cycle pressure losses compared with other propulsion systems. However, a maximum take-off thrust of 110 kN at a TET of 800 \({}^{\circ}\)C was chosen due to cooling requirement and heat sink design constraints. In the preliminary design, one-third of the pumping power is assigned to the reactor and one-sixth to the precooler. With much high thermal conductivity and permissible flow velocity; helium heat exchangers are compact. The problem with Ranking rejection at almost constant temperature is in contrast with the closed-cycle gas turbine as heat is rejected over a wide range of temperature. The success of the closed-cycle application for UAV is tied to the heat sink. The radiator area should be compact and efficient at rejecting the waste for recycling of the helium working fluid into the compressor. In the preliminary analysis, pressure losses were not considered. Table 3 shows the system performance at 30,000 m and a Mach number of 0.7. The control at different thrust requirement can be easily achieved with pressure level control, heat source and by-pass control. For most cases, pressure level control is performed at almost constant SFC and a constant temperature in all component. Parametric sizing and weight estimates were made for the turbomachinery and precooler and reactor.