The integration of a propulsion system into the airframe can contribute to the overall aircraft performance and is beneficial for military applications due to a reduction of the aircraft radar cross section. On the other hand, engine integration calls for serpentine engine intake systems in which generally combined pressure-swirl distortions are provoked, which affect the performance of the propulsion system. The military engine intake research duct was designed to provoke a large-scale flow separation as well as a combined pressure-swirl distortion. The serpentine research duct is tested in both a remote- and close-coupled setup with the Larzac 04 turbofan engine to assess the aerodynamic interactions between the flow within the intake duct and the compression system. The upstream effect of the compression system on the steady-state duct flow is within the range of expectations. Less is known from the open literature with regard to the unsteady character of the flow in such a setup. Three distinct unsteady flow phenomena caused by the serpentine shape of the duct are identified up- and downstream of the low pressure compressor in the remote-coupled setup. An additional distinct low frequency unsteadiness is provoked with a configuration which features vortex generators for flow control. All phenomena are largely attenuated by the upstream effect of the compressor in the close-coupled setup. Nonetheless, the surge margin is massively reduced due to the inflow distortion and stall cells within the first stage of the low pressure compressor are visualized at high thrust settings of both the remote- and close-coupled setup, which are expected to impact the durability of the compression system.

This paper presents an analysis of the stability margin improvement (SMI), which is also known as stall margin improvement, achieved by continuous tip air injection. New piezoelectric actuators were designed and manufactured with a new engine inlet for the Larzac 04 C5 jet engine. It has noninvasive injection positions that do not have any measurable effect on the inlet air flow when it is switched off. The main focus of the system design was to achieve high power of the injected air and, as a result, a high SMI. The results presented enable a maximum SMI of 99%. A variety of engine operating conditions and injection positions were experimentally tested and discussed regarding SMI. Additionally, the complex relationship between SMI gains and thrust specific fuel consumption (TSFC) is explored in a power balance analysis, revealing a trade-off between SMI improvement and increased energy consumption.

The role of unmanned aerial systems (UAS) is becoming increasingly important in aerospace development across a range of disciplines. The integration of complete engines into aircraft fuselages, along with highly curved intake ducts, offers significant advantages due to their compact size and minimal impact on structural weight. A compact intake design can lead to substantial flow distortions and negatively affect the performance and flow stability of the engine. Today’s industry practice for flow distortion quantification in propulsion system integration relies on isolated pressure measurements at the engine front. Recent studies have shown that this low-accuracy approach is not adequate to complex future aircraft designs, as these assessment methods cannot capture pertinent features of flow distortion. Traditional distortion descriptors like the DC60 can provide a simplified total pressure distortion based on a 60-degree weighted angular section. More complex distortion patterns beyond the coverage of these sections are underrated. The phenomena arising from the instability of secondary flows cannot be captured adequately by traditional descriptors. Frequencies associated with certain flow distortion phenomena significantly affect the engine fan’s response. This study uses additional newly defined flow distortion descriptors to close these gaps, such as the area integral of the secondary kinetic energy or the resulting load of the wall pressure distribution at the aerodynamic interface plane. Using the coefficient for the secondary vorticity also makes it possible to identify the dominant frequencies of the vortex systems that occur in the coupled intake system and affect the engine. The coupling shifts the first mode of the central flow separation from 15.7 to 25.1 Hz and introduces a new 75.2 Hz mode. A dominant 150.4 Hz mode, close to the measured 153 Hz is also found. The results demonstrate that advanced, time-resolved distortion metrics enable an effective analysis of intake–compressor interaction and are therefore essential for evaluating highly integrated propulsion systems.