Thin film heat flux gauges (HFGs) have been used for decades to measure surface temperatures and heat flux in test turbines with the majority being used in facilities that are short-duration. These gauges are typically composed of two resistive temperature devices deposited on opposing sides of a dielectric. However, because these sensors have been traditionally applied for measurements in transient-type facilities, the challenges facing adaptation of this technology for a steady facility warrant investigation. Those challenges are highlighted, and solutions are presented throughout the paper. This paper describes the nanofabrication process for heat flux gauges and a new calibration method to address potential deterioration of gauges over long runtimes in continuous-duration facilities. Because a primary uncertainty of these sensors arises from the ambiguity of the thermal properties, emphasis is placed on the property determination. Also, this paper presents a discussion on the use of impulse response theory to process the data showing the feasibility of the method for steady-duration facilities after an initial settling time. The latter portion of the paper focuses on comparing well-established heat flux gauges developed for short-duration turbine test facilities to recently developed gauges fabricated using modern nanofabrication techniques for a continuous turbine test facility. The gauges were compared using the test case of an impinging jet over a range of Reynolds numbers. The comparison between the PSU gauge and the reference device indicated agreement within 14%, and similar results were achieved through comparison with established sensors from partner institutions.

Large rotor tip clearances and the associated tip leakage flows are known to have a significant effect on overall compressor performance. However, detailed experimental data reflecting these effects for a multistage compressor are limited in the open literature. As design trends lead to increased overall compressor pressure ratio for thermal efficiency benefits and increased bypass ratios for propulsive benefits, the rear stages of the high-pressure compressor will become physically small. Because rotor tip clearances cannot scale exactly with blade size due to the margin needed for thermal growth considerations, relatively large tip clearances will be a reality for these rear stages. Experimental data have been collected from a three-stage axial compressor to assess performance with three-tip clearance heights representative of current and future small core machines. Trends of overall pressure rise, stall margin, and efficiency are evaluated using clearance derivatives, and the summarized data presented here begin to narrow the margin of tip clearance sensitivities outlined by previous studies in an effort to inform future compressor designs. Furthermore, interstage measurements show stage matching changes and highlight specific differences in the performance of rotor 1 and stator 2 compared to other blade rows in the machine.

No abstract

No abstract

As engine development continues to advance toward increased efficiency and reduced fuel consumption, efficient use of compressor bypass cooling flow becomes increasingly important. In particular, optimal use of compressor bypass flow yields an overall reduction of harmful emissions. Cooling flows used for cavity sealing between stages are critical to the engine and must be maintained to prevent damaging ingestion from the hot gas path. To assess cavity seals, the present study utilizes a one-stage turbine with true-scale engine hardware operated at engine-representative rotational Reynolds number and Mach number. Past experiments have made use of part-span (PS) rather than full-span (FS) blades to reduce flow rate requirements for the test rig; however, such decisions raise questions about potential influences of the blade span on sealing effectiveness measurements in the rim cavity. For this study, a tracer gas facilitates sealing effectiveness measurements in the rim cavity to compare data collected with FS engine airfoils and simplified, PS airfoils. The results from this study show sealing effectiveness does not scale as a function of relative purge flow with respect to main gas path flow rate when airfoil span is changed. However, scaling the sealing effectiveness for differing spans can be achieved if the fully purged flow rate is known. Results also suggest reductions of purge flow may have a relatively small loss of seal performance if the design is already near a fully purged condition. Rotor tip clearance is shown to have no effect on measured sealing effectiveness.

No abstract

Rotor tip clearance height and the associated tip leakage flow have a significant effect on the performance and stability of compressors. Existing studies considering tip clearance effects on stability have been primarily limited to low-speed compressors, and many of these evaluated single-stage machines, which may not adequately represent stall trends for enginescale compressors. Furthermore, test campaigns for engine-scale compressors cannot provide instrumentation accessibility required for detailed stage performance and stall investigations. Using results collected from a three-stage intermediate-speed axial compressor with appreciable density rise, this study addresses these needs. In this paper, three rotor tip clearances are tested, ranging from 1.5% to 4% span (1% to 3% chord). Previous studies have primarily shown a transition from short-length-scale spikes to longlength-scale modes as the clearance is increased, whereas the present study shows the opposite: a transition from modes to spikes with increased tip clearance. As a result, these data emphasize that a definitive trend does not exist between the stall inception mechanism and increasing tip clearance. Instead, the clearance effects alter stage matching with speed and change the stall inception mechanism. These results also elicit future research by preliminarily suggesting that stall inception mechanisms may be predictable from steady performance measurements collected in the stalling stage.

No abstract