With the advancements in miniaturization and temperature capabilities of piezoresistive pressure sensors, pneumatic probes—which are the long established standard for flow-path pressure measurements in gas turbine environments—are being replaced with unsteady pressure probes. Any measured quantity is by definition inherently different from the “true” value, requiring the estimation of the associated errors for determining the validity of the results and establishing respective confidence intervals. In the context of pressure measurements, the calibration uncertainty values, which differ from measurement uncertainties, are typically provided. Even then, the lack of a standard methodology is evident as uncertainties are often reported without appropriate confidence intervals. Moreover, no time-resolved measurement uncertainty analysis has come to the attention of the authors. The objective of this paper is to present a standard method for the estimation of the uncertainties related to measurements performed using single sensor unsteady pressure probes, with the help of measurements obtained in a one and a half stage low pressure (LP) high speed axial compressor test rig as an example. The methodology presented is also valid for similar applications involving the use of steady or unsteady sensors and instruments. The static calibration uncertainty, steady measurement uncertainties, and unsteady measurement uncertainties based on phase-locked average (PLA) and ensemble average are presented in this contribution. Depending on the number of points used for the averaging, different values for uncertainty have been observed, underlining the importance of having greater number of samples. For unsteady flows, higher uncertainties have been observed at regions of higher unsteadiness such as tip leakage vortices, hub corner vortices, and blade wakes. Unfortunately, the state of the art in single sensor miniature unsteady pressure probes is comparable to multihole pneumatic probes in size, preventing the use of multihole unsteady probes in turbomachinery environments. However, the angular calibration properties of a single sensor probe obtained via an aerodynamic calibration may further be exploited as if a three-hole directional probe is employed, yielding corrected total pressure, unsteady yaw angle, static pressure, and Mach number distributions based on the PLAs with the expense of losing the time-correlation between the virtual ports. The aerodynamic calibration and derivation process are presented together with the assessment of the uncertainties associated to these derived quantities by the authors in Dell'Era et al. (2016, “Assessment of Unsteady Pressure Measurement Uncertainty—Part II: Virtual Three Hole Probe,” ASME J. Eng. Gas Turbines Power, 138(4), p. 041602). In the virtual three-hole mode, similar to that of a single sensor probe, higher uncertainty values are observed at regions of higher unsteadiness.

The present paper proposes a concept for a water-cooled high temperature unsteady total pressure probe intended for measurements in the hot sections of industrial gas turbines or aero-engines. This concept is based on the use of a conventional miniature piezoresistive pressure sensor, which is located at the probe tip to achieve a bandwidth of at least 40 kHz. Due to extremely harsh conditions and the intention to immerse the probe continuously into the hot gas stream, the probe and sensor must be heavily cooled. The short term objective of this design is to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (1100–1400 K) and in the long term at combustor exit (2000 K or higher).

This paper presents the first experimental engine and test rig results obtained from a fast response cooled total pressure probe. The first objective of the probe design was to favor continuous immersion of the probe into the engine to obtain time series of pressure with a high bandwidth and therefore statistically representative average fluctuations at the blade passing frequency. The probe is water cooled by a high pressure cooling system and uses a conventional piezo-resistive pressure sensor which yields therefore both time-averaged and time-resolved pressures. The initial design target was to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (1100–1400K) with a bandwidth of at least 40kHz and in the long term at combustor exit (2000K or higher). The probe was first traversed at the turbine exit of a Rolls-Royce Viper turbojet engine, at exhaust temperatures around 750 °C and absolute pressure of 2.1bars. The probe was able to resolve the high blade passing frequency (≈23kHz) and several harmonics up to 100kHz. Besides the average total pressure distributions from the radial traverses, phase-locked averages and random unsteadiness are presented. The probe was also used in a virtual three-hole mode yielding unsteady yaw angle, static pressure and Mach number. The same probe was used for measurements in a Rolls-Royce intermediate pressure burner rig. Traverses were performed inside the flame tube of a kerosene burner at temperatures above 1600 °C. The probe successfully measured the total pressure distribution in the flame tube and typical frequencies of combustion instabilities were identified during rumble conditions. The cooling performance of the probe is compared to estimations at the design stage and found to be in good agreement. The frequency response of the probe is compared to cold shock tube results and a significant increase in the natural frequency of the line-cavity system formed by the conduction cooled screen in front of the miniature pressure sensor were observed.

This paper presents the first experimental engine and test rig results obtained from a fast response cooled total pressure probe. The first objective of the probe design was to favor continuous immersion of the probe into the engine to obtain a time series of pressure with a high bandwidth and, therefore, statistically representative average fluctuations at the blade passing frequency. The probe is water cooled by a high pressure cooling system and uses a conventional piezoresistive pressure sensor, which yields, therefore, both time-averaged and time-resolved pressures. The initial design target was to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (800–1100°C) with a bandwidth of at least 40 kHz and in the long term at combustor exit (2000 K or higher). The probe was first traversed at the turbine exit of a Rolls-Royce Viper turbojet engine at exhaust temperatures around 750°C and absolute pressure of 2.1 bars. The probe was able to resolve the high blade passing frequency (≈23 kHz) and several harmonics of up to 100 kHz. Besides the average total pressure distributions rom the radial traverses, phase-locked averages and random unsteadiness are presented. The probe was also used in a virtual three-hole mode yielding unsteady yaw angle, static pressure, and Mach number. The same probe was used for measurements in a Rolls-Royce intermediate pressure burner rig. Traverses were performed inside the flame tube of a kerosene burner at temperatures above 1600°C. The probe successfully measured the total pressure distribution in the flame tube and typical frequencies of combustion instabilities were identified during rumble conditions. The cooling performance of the probe is compared with estimations at the design stage and found to be in good agreement. The frequency response of the probe is compared with cold shock-tube results and a significant increase in the natural frequency of the line-cavity system formed by the conduction cooled screen in front of the miniature pressure sensor were observed.

The present paper proposes a concept for a water cooled high temperature unsteady total pressure probe, intended for measurements in the hot sections of industrial gas turbines or aero-engines. This concept is based on the use of a conventional miniature piezo-resistive pressure sensor, located at the probe tip to achieve a bandwidth of at least 40kHz. Due to extremely harsh conditions and the intention to immerse the probe continuously into the hot gas stream, the probe and sensor must be heavily cooled. The short term objective of this design is to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (1100–1400K) and in the long term at combustor exit (2000K or higher).