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An experimental study is conducted on bluff-body stabilized premixed flames in a curved, square cross section duct. High flow velocities coupled with a small radius of curvature of the duct induce high centrifugal acceleration normal to the flame sheet. A cylindrical flame holder spans the width of the square cross section and is positioned at the channel midheight. Flame shear layers are stabilized on the radially inward (upper) and outward (lower) edges of the flame holder. Side-view high-speed Schlieren images and high-speed pressure measurements are captured. Static stability, overall pressure loss, and statistics and velocimetry results from the Schlieren images are reported, and results are compared to a straight configuration with no centrifugal acceleration. Two bluff-body diameters are studied to show the effect of flame holder diameter on static stability. For the curved configuration, blowout velocities are higher for the smaller bluff-body diameter, likely due to flow acceleration effects. Blowout velocities are lower for the curved configuration compared to the straight configuration which may be due to the destabilizing Rayleigh–Taylor (RT) effect on the upper flame layer. Overall pressure loss is slightly higher for the curved configuration than the straight configuration. High-speed Schlieren results show centrifugal acceleration causes significant structural and velocimetric asymmetry in the bluff-body wake. In the curved configuration, the upper flame layer displays destabilizing RT instabilities, and the lower flame layer displays stabilizing RT effects. The upper flame shows vigorous RT instabilities which broaden the flame brush and sustain a flame leading edge independent of inlet Reynolds number or velocity. Conversely, the lower flame exhibits suppression of Kelvin–Helmholtz and flame-generated instabilities in the wake, which confines the flame brush and significantly reduces transverse flame velocities. The lower flame edge profile moves toward the channel centerline with increasing inlet Reynolds number. The upper flame in the curved configuration shows higher flame edge velocities than the straight configuration while the lower flame shows velocities closer to zero. The empirical constant to the power law relation for upper flame edge velocities agrees with RT-dominated flame growth theory for this experimental scale and agrees with other RT-dominated flame studies.

The detrended fluctuation analysis, a technique for detecting long-range correlations in fractal time series, is used to characterize thermoacoustic oscillations of impending combustion instabiliti...

The structure and dynamics of a hydrogen-air rotating detonation engine (RDE) are described based on 100-kHz laser absorption spectroscopy measurements of water temperature at four simultaneous locations within the detonation channel. The analysis focuses on the evolution of the flowfield over a 200 ms period for three separate air mass flow rate cases. Two-dimensional unwrapped visualizations of the temperatures show a flowfield structure containing regions with the detonation front, combustion products, oblique shock, and refilling reactants, qualitatively agreeing with previous simulations and experiments. A major conclusion is that water from the combustion products is measured throughout all space and time in the RDE, including near the injector, implying the presence of performance loss processes such as burning upstream of the detonation wave or the back recirculation of combustion products with fresh fuel–air. By analyzing the elevated temperatures of the reactants during the refill process, one estimation for the mass fraction of combustion products in the reactants is as high as 20–30% on average. This product mass fraction is found to be inversely proportional to the bulk air mass flow rate and decreases as time progresses. This indicates these non-ideal processes are more significant closer to RDE ignition for poorer performing operating conditions. For the largest air mass flow case, water temperatures near the nominally cold plenum conditions likely corroborate the presence of a recirculation region on the RDE inner body. Analysis of inter- and intra-cycle temperature dynamics further support non-ideal processes occurring behind the detonation wave and during the refill process. As a whole, the data indicates that the RDE performance is better as time progresses away from ignition or for higher air mass flow rates. These data are also important for comparison with numerical models.

New observations are reported of the processes leading up to lean blowoff (LBO) in a two-dimensional confined bluff body flame by visualizing LBO from a spanwise perspective. Time-resolved OH* chemiluminescence imaging simultaneous with particle image velocimetry are used to characterize the spanwise structure and dynamics. For reacting conditions, newly observed large-scale corner vortex structures (CVS) are present downstream of the bluff body inside the recirculation zone near the span walls. The CVS are orthogonal to the conventional recirculation zone observed in a bluff body flow, and the CVS are absent under non-reacting conditions. As the equivalence ratio is decreased, the vorticity and circulation of the CVS increase, contributing to significant nonuniformity across the spanwise direction both inside the recirculation zone and on the surrounding combustion. Approaching LBO, increased quantities of vanishing combustion near the span walls implies increased quantities of reactants entering and diluting the recirculation zone. Since cold reactant entrainment into the recirculation zone has been suggested as a possibly important LBO contributing factor, the current study suggests that this wake disruption mechanism begins much earlier than previously thought. The CVS are also the final regions of the flow to extinguish before LBO is complete, emphasizing that LBO is a highly three-dimensional process. A bluff body with an aspect ratio three times larger was also investigated, and the LBO process is observed to remain qualitatively similar with significant spanwise nonuniformity and containing the CVS. This paper intends to improve the three-dimensional understanding of two-dimensional bluff body stabilized premixed flames approaching LBO.

Reported is a three-step procedure (also referred to as the observer) for fast and unbiased determination of the dominant frequencies and amplitude of thermoacoustic oscillations. The first step is...

10 kHz two-color OH planar laser-induced fluorescence (PLIF) thermometry was demonstrated in both laminar Hencken flames and turbulent premixed jet flames using two injection-seeded optical parametric oscillators (OPOs) pumped by a high-speed three-legged burst-mode laser. The two burst-mode OPOs generate ∼ 5 m J / p u l s e at 282 nm and 286 nm to excite the Q 1 ( 5 ) and Q 1 ( 14 ) transitions of the A 2 Σ + ← X 2 Π (1,0) system of OH, respectively. PLIF images were collected simultaneously from each of the two transitions and ratios of intensities from the two images were used to determine local temperatures. Analyses of flame curvature, temperature, and the correlation in time of these two quantities are also discussed. The results from this work are promising for the use of this technique in more complex flow environments and at, potentially, even higher repetition rate.

This study characterizes the structure and dynamics of a confined, bluff-body-stabilized turbulent premixed flame by simultaneously employing formaldehyde (CH2O) and hydroxyl (OH) planar laser-induced fluorescence (PLIF) and particle image velocimetry (PIV), at a rate of 10 kHz. The large field-of-view (>170 cm2) CH2O-PLIF is enabled by use of a burst-mode laser delivering 10-kHz pulse trains of 355-nm at 350 mJ/pulse, resulting in a CH2O signal-to-noise of 47:1 during PIV seed flow. Two cases illustrative of the CH2O dynamics are presented. A statistically stationary turbulent combustion case highlights the development of the CH2O layers in space and time. Notably, presumed CH2O-vortex dynamic interactions are observed where the CH2O accumulates broadly near the Kelvin–Helmholtz vortex core and remains thin near the vortex braid, contributing to a distribution of CH2O preheat zone thickness from 1 to 10 times of the calculated laminar value. The second case highlights the CH2O dynamics during a self-excited combustion instability. Two short-duration increases in CH2O are produced during the elevated velocity portion of the acoustic cycle. The first CH2O increase is caused by the reactant mass flux impulse as the velocity starts to increase. The second CH2O increase is the result of the upper and lower shear layers merging downstream and entraining fresh reactants that burn in intense, distributed regions inside the wake. Estimating the time delay between the CH2O and the heat release, it is suggested that the secondary CH2O increase may contribute to damping of the acoustic instability, because of its out-of-phase relationship with pressure, while the first CH2O increase appears to drive the instability.

This work presents fully coupled computational fluid dynamics (CFD) simulations and thermodynamic cycle analyses of a small-scale turbojet engine at several conditions along the equilibrium running line. The CFD simulations use a single mesh for the entire engine, from the intake to the exhaust, allowing information to travel in all directions. The CFD simulations are performed along the equilibrium running line by using the iterative Secant method to compute the fuel flow rate required to match the compressor and turbine power. The freestream pressure and temperature and shaft angular speed are the only inputs needed for the CFD simulations. To evaluate the consistency of the CFD results with thermodynamic cycle results, outputs from the CFD simulations are prescribed as inputs to the cycle model. This approach enables on-design and off-design cycle calculations to be performed without requiring turbomachinery performance maps. In contrast, traditional off-design cycle analyses require either scaling, calculating, or measuring compressor and turbine maps with boundary condition assumptions. In addition, the CFD simulations and the cycle analyses are compared with measurements of the turbojet engine. The CFD simulations, thermodynamic cycle analyses, and measurements agree in terms of total temperature and pressure at the diffuser–combustor interface, air and fuel mass flow rate, equivalence ratio, and thrust. The developed methods to perform CFD simulations from the intake to the exhaust of the turbojet engine are expected to be useful for guiding the design and development of future small-scale gas turbine engines.

High-speed synchronized stereo particle-imaging velocimetry and OH planar laser-induced fluorescence (PIV/OH-PLIF) measurements are performed on multiple $R{-}\unicode[STIX]{x1D703}$ planes downstream of a high-Reynolds-number swirling jet. Dynamic-mode decomposition (DMD) – a frequency-resolved data-reduction technique – is used to identify and characterize recurrent flow structures. Illustrative results are presented in a swirling flow field for two cases – the nominal flow dynamics and where self-excited combustion driven oscillations provide strong axisymmetric narrowband forcing of the flow. The robust constituent of the nominal reacting swirl flow corresponds to a helical shear-layer disturbance at a Strouhal number ($St$) of ${\sim}0.30$, $St=fD/U_{0}$, where $f$, $D$ and $U_{0}$ denote the precessing vortex core (PVC) frequency (${\sim}800~\text{Hz}$), the swirler exit diameter (19 mm) and the bulk velocity at the swirler exit ($50~\text{m}~\text{s}^{-1}$) respectively. Planar projections of the PVC reveal a pair of oscillating skew-symmetric regions of velocity, vorticity and OH-PLIF intensity that rotate in the same direction as the mean tangential flow. During combustion instabilities, the large-amplitude acoustics-induced axisymmetric forcing of the flow results in a fundamentally different flow response dominated by a nearly axisymmetric disturbance and almost complete suppression of the large-scale helical shear-layer disturbances dominating the nominal flow. In addition, reverse axial flows around the centreline are significantly reduced. Time traces of the robust constituent show reverse axial flows around the centreline and negative axial vorticity along the inner swirling shear layer when the planar velocity is in the same direction as the mean tangential flow. For both stable and unstable combustion, recurrent flow structures decay rapidly downstream of the air swirler, as revealed by the decreasing amplitude of the velocity, axial vorticity and OH-PLIF intensity.

The detonations propagating through the annular channel of an optically accessible rotating detonation engine (RDE) operating on hydrogen–air are visualized using OH* chemiluminescence imaging. The images are useful for observing the instantaneous size and shape of the detonation structure, oblique shock wave, and possible presence of deflagration between the fuel-fill zone and expansion region containing detonation products. The detonation increases in height as the air flow rate is increased for low flow rates, experiences subtle changes for intermediate flow rates, and transitions from one to two waves for higher flow rates. The two detonation waves typically propagate in the same azimuthal direction. Counter-rotating waves resulting in detonation–detonation interactions are observed for some configurations with a reduced number of fuel injection jets. Time-dependent static pressure measurements show that acoustic interactions between the detonation channel and air plenum are important for low air flow rates and large air injection areas. The OH* chemiluminescence images, pressure, and wave speed measurements provide benchmark data that are useful for evaluating RDE models and simulations, improving fundamental understanding of the detonation structure in RDEs, and identifying critical design parameters that influence RDE operation and performance.

A numerical optimization study is performed on a small-scale high-swirl cavity-stabilized combustor. A parametric geometry is created in CAD software that is coupled with meshing software. The latter automatically transfers meshes and boundary conditions to the solver, which is coupled with a post-processing tool. Steady, incompressible three-dimensional simulations are performed using a multi-phase Realizable k-ϵ Reynolds-averaged Navier-Stokes (RANS) approach with the non-adiabatic flamelet progress variable (FPV). There are nine input parameters based on geometrical control variables. There are five output parameters, viz., pattern factor (PF), RMS of the profile factor deviation, averaged exit temperature, averaged exit swirl angle, and total pressure loss. An iterative design of experiments (DOE) with a recursive Latin Hypercube Sampling (LHS) is performed to filter the most important input parameters. The five major input parameters are found with Spearman’s order-rank correlation and R2 coefficient of determination. The five input parameters are used for the adaptive multiple objective (AMO) optimization. The AMO algorithm provided a candidate design point with the lowest weighted objective function. This design point was verified through CFD simulation. The combined filtering and optimization procedures improve the baseline design point in terms of pattern and profile factor. The former halved from that of the baseline design point whereas the latter turned from an outer peak to a center peak profile, closely mimicking an ideal profile. The exit swirl angle favorably increased 25%. The averaged exit temperature and the total pressure losses remained nearly unchanged from the baseline design point.

The capability of current turbomachinery-based engines limit the obtainable altitude and flight Mach number of modern aircraft. Maturing hypersonic technologies such as ram and scramjets allow greatly increased flight velocity but are not able to power themselves from the ground and, thus, rely on lift aircraft. Combined-cycle engines incorporate turbomachinery and ramjet technologies to allow both high and low flight Mach numbers. These high Mach capable systems are typically flown using specialty fuels such as JP-7 that are more applicable to hypersonic applications than conventional gas-turbine fuels such as JP-8/Jet A or JP-5. Potential issues exist, however, when operating a combined cycle that employs a legacy main combustor since these platforms were not designed for operation with these fuels. Of particular concern is the re-ignition performance of the turbomachinery core. The relight of the gas-turbine combustor occurs at high altitudes and relatively high vehicle speeds, yielding low pressure and temperature in the combustor as well as high combustor-dome velocities. All of these conditions are unfavorable for ignition. Additionally, heavy fuels require more energy for atomization and vaporization, which increases the probability that ignition will become a problem in a turbine-based combined-cycle (TBCC) system. Successful demonstration of legacy main-combustor technologies in hypersonic combined-cycle applications will eliminate the need for costly design of new main-burner technology. The literature does not provide information on the effects of running specialty fuels such as JP-7 and JP-10 in burners with conventional aerodynamic features. To fill that gap, a three-cup sector of a conventional fighter-class swirl-stabilized combustor configured to provide optical access through the sidewall was used in the study. Two test conditions representative of varying flight Mach number and altitude are evaluated. For each flight condition, the combustor pressure drop is varied to characterize ignition as a function of burner inlet velocity. Data are correlated by loading parameter and equivalence ratio at ignition, and at each test point the conditions are varied until ignition cannot be achieved.

One of the challenges of gas turbine combustor research is to accurately measure and model air mass flow rates through complex air injection schemes. Accurate measurements and computations of air mass flow rates are necessary for determining air and fuel distributions, which influence a range of combustor operation and performance characteristics. Experimental and computational studies were performed on a representative gas turbine combustor swirler. The swirler geometry consists of four component flows: two co-rotating annular axial swirls, one radial swirl, and cooling on the periphery of the swirler. The purpose of this study is to compare measured and computed air mass flow rates in a realistic swirler with a complex geometry and to quantify the magnitude of the interaction effects between air passages. The measurements of the air mass flow rates were performed using a calibrated air flow stand. The computations were performed using commercial Computational Fluid Dynamics (CFD) tools. Comparisons between measured and computed air mass flow rates show good agreement for the individual and total flow configurations. Significant interaction effects among the swirling flows are observed when all of the air passages are open. The radial swirl mass flow rate decreases by 2.7% and the outer axial swirl mass flow rate increases by 3.8% when the individual component flow configuration is compared to the total flow configuration. The computed mass flow rates demonstrate that the interactions among the swirl flows create a significant change in mass flow distribution within the swirler.

Cycle-resolved measurements of H2O temperatures and number densities taken within the detonation channel of a hydrogen—air rotating detonation engine (RDE) at a 100 kHz repetition rate using laser absorption spectroscopy are presented. The laser source used is an MEMS-tunable Vertical-Cavity Surface Emitting laser which scans from 1330 to 1360 nm. Optical access into and out of the RDE is achieved using a dual-core fiber optic. Light is pitched into the RDE through a sapphire window via a single-mode core, retroreflected off the mirror-polished inner radius of the RDE annulus, and collected with the multi-mode fiber core. The resulting absorption spectra are used to determine gas temperatures as a function of time. These measurements allow characterization of the transient-temperature response of the RDE.

This article describes an investigation of flame shapes and flow configurations in a premixed, swirl-stabilized dump combustor. High swirl, annular nozzle flows of this nature enable a variety of different flame configurations and heat release distributions with their associated flow fields. These differences are significant, since each of these configurations, in turn, has different thermoacoustic sensitivities and influences on combustor emissions, nozzle lifetime, and liner heating. These different configurations arise because multiple flame stabilization locations are present, associated with the inner and outer shear layers of the annulus, and the stagnation point of the vortex breakdown region. We present results from high-speed luminosity imaging, particle image velocimetry (PIV), and OH-planar laser induced fluorescence (PLIF) to illustrate time-averaged and instantaneous flame shapes and flow fields associated with the different configuration “families.” Selected cases are compared with large edd...

A mid-infrared fiber-coupled laser system constructed around three time-division-multiplexed quantum-cascade lasers capable of measuring the absorption spectra of CO, CO2, and N2O at 100 kHz over a wide range of operating pressures and temperatures is demonstrated. This system is first demonstrated in a laboratory burner and then used to measure temperature, pressure, and concentrations of CO, CO2, and N2O as a function of time in a detonated mixture of N2O and C3H8. Both fuel-rich and fuel-lean detonation cases are outlined. High-temperature fluctuations during the blowdown are observed. Concentrations of CO are shown to decrease with time for fuel-lean conditions and increase for fuel-rich conditions.

The fuel and air mixing processes in an optically accessible non-premixed rotating detonation engine (RDE) are visualized using acetone planar laser induced fluorescence (PLIF) imaging. The acetone PLIF images are used to observe the transient fuel injection processes and evaluate the extent of partial premixing upstream of the detonation wave. The acetone PLIF images complement past OH* chemiluminescence images which showed the instantaneous size and shape of the detonation structure, oblique shock wave, and possible presence of deflagration between the fuel-fill zone and expansion region containing detonation products. The acetone PLIF data presented in this work represents a recent and ongoing experimental investigation that provides insightful information on the transient processes in the RDE. The acetone PLIF images of the non-reacting flow show an impinging jet in crossflow consistent with the fuel injection scheme of the current RDE design. A recirculation zone with minimal fuel concentration is observed in the outer corner near the fuel injection surface of the annular detonation channel. The acetone PLIF images of the reacting flow indicate that there is a purging period (60 – 75 μs corresponding to 18 – 22 % of the cycle) in which fuel is not being injected into the channel after the detonation wave travels past a particular fuel jet. This observation suggests that the high-pressure detonation wave inhibits the inflow of fuel during the purging period. The application of established experimental techniques such as acetone PLIF and OH* chemiluminescence imaging is providing new insights into RDEs. The results provide benchmark measurements that are useful for evaluating RDE models and simulations, improving fundamental understanding of the detonation structure in RDEs, and identifying critical design parameters that influence RDE operation and performance.

Small internal combustion engines, particularly those ranging in power from 1 kW to 10 kW, propel many remotely piloted aircraft (RPA) platforms that play an increasingly significant role in the Department of Defense. Efficiency of these engines is low compared to conventional scale engines and thermal losses are a significant contributor to total energy loss. Existing thermal energy loss models are based on data from much larger engines. Whether these loss models scale to the engine size class of interest, however, has yet to be established. The Small Engine Research Bench (SERB) was used to measure crank angle resolved gas temperature inside the combustion chamber of a small internal combustion engine (ICE). A 55 cc, twostroke, spark-ignition ICE was selected for this study. The engine was modified for optical analysis using sapphire rods 1.6 mm in diameter on opposite sides of the combustion chamber. The engine modification was found to have no measurable impact on indicated mean effective pressure or heat rejection through the cylinder. FTIR absorption thermometry was used to collect mid-infrared absorption spectra. The FTIR was allowed to scan continuously while simultaneously recording the scanning mirror position and crank angle associated with each data point, then data was re-sorted by crank angle. Measured spectra were compared with lines generated using CDSD-4000 and HITEMP line list databases. The line of best fit corresponded to the mean gas temperature through the combustion chamber. In this way temperature was determined as a function of crank angle for three operating conditions: 4,300, 6,000, and 7,500 revolutions per minute, all at wide open throttle. High cycle-to-cycle variation in the regions of combustion and gas exchange degraded temperature measurements at the affected crank angles. Future research will attempt to improve signal to noise in these measurements.

Volatile fuel costs have initiated a global sprint for technologies that will increase fuel efficiency in gas turbine engines, which continue to be a primary propulsion system for commercial and military aircraft. The Ultra-Compact Combustor (UCC) is an advanced gas turbine combustor which integrates the high pressure turbine inlet guide vanes into the combustor. In comparison to conventional combustor systems, the UCC has the potential to shorten engine length, decrease engine weight, and reduce pressure losses entering the high pressure turbine rotor, all of which contribute subtantially to improving engine efficiency. The UCC has been studied in two forms in the literature – the Trapped-Vortex concept (TV) and the High-g concept (HG); the current study focuses on the latter. The HG-UCC utilizes a rectangular cavity wrapped around the outside circumference of the combustor where fuel and air are injected in a manner that generates a highly swirling flow within the circumferential cavity. The flow within the cavity experiences a centripetal acceleration (“g-load”) which varies with tangential velocity and radius from the combustor axis. This “high-g” effect has the potential to increase flame speeds and volumetric heating rates, as well as enhance fuel air mixing and spreading providing the potential to reduce fuel injector count without compromising performance. Experiments were performed on the HG-UCC combustor at higher operating pressures, temperatures, and equivalence ratios than previous high-g experiments conducted in a similar uncooled inter-turbine burner (ITB). Results show that the HG-UCC has high combustion efficiencies for the entire range of data considered. The effects of pressure and temperature, cavity driver angle, centripetal acceleration (g-load), and equivalence ratio are studied for their impacts on combustion efficiency and nitrogen oxide (NOx) emissions. An important result of this work is the construction of NOx correlations for the HG-UCC. Test data performance results and correlations are compared to previous trapped vortex and high-g UCC experiments.

During the past fiscal year, researchers with the Air Force Research Laboratorys Combustion Branch made substantial progress in numerous areas including: 1) ultra-compact combustors, 2) inter-turbine burner concepts, 3) bluff-body stabilized turbulent flames, 4) well-stirred reactors for chemical kinetics, and 5) detonation-stabilized turbulent flames.Lean blowout data was collected on propane and jet fuel bluff-body stabilized flames and was combined with data taken from past literature to create a database of over 1,100 data points. The best correlation for the overall dataset included pressure, temperature, the ratio of flameholder diameter to lip velocity, oxygen level, and the hydrogen-to-carbon ratio of the fuel as factors. The R-squared value of this correlation was 0.873. The exponents on the factors indicated that high pressure and temperature lowered the equivalence ratio at lean blowout, as did high levels of oxygen. The correlation for the ignition delay dataset with pressure, temperature, and the ratio of flameholder diameter to lip velocity as factors had an R-squared value of 0.942. This relatively small, uniform dataset correlates very well. When ignition delay time replaces pressure and temperature, the data continues to correlate well, with an R-squared value of 0.918. For this limited dataset, the ignition delay time is an adequate representation of the chemical timescale.

A novel technique for selecting wavelengths that optimize the temperature precision of absorption-based thermometry is developed. A single universal objective function considering continuous spectral lower-state energy E v ″ , spectral absorbance intensities and an experimental noise level is derived and applied in a brute-force approach to analyze two-wavelength, Doppler-broadened H2O vapor-absorbance thermometry in the 1.33–1.37 μm range. The results reveal the top wavelength pairs and their temperature precisions over a wide temperature range from 280 to 2800 K. Although the spectral database used includes over thousands of transitions in this spectral range, the top pairs are composed of only 12 unique wavelengths. This wavelength selection technique is able to help researchers choose key wavelengths that will perform well in practical applications.

As is well known, individual molecular spectroscopic transition can be characterized by a single, fixed parameter that describes the transition’s temperature sensitivity. This parameter is commonly known as the lower-state energy. In practical spectra, spectral lineshapes cause the effective lower-state energy to vary continuously as a function of frequency, and blending of adjacent features causes the effective lower-state energy to vary with temperature as well. Accordingly, a frequency-dependent lower-state energy can be defined. This lower-state energy is easily computed from measured or simulated spectra and is valuable for detailed assessment of the temperature sensitivity of practical spectra.

Gas-temperature measurements in the combustion zone of a high-pressure gas-turbine-combustor sector rig were made with a Fourier-domain mode-locked laser using wavelength-agile absorption-spectroscopy techniques. These measurements are among the first employing broadband high-resolution absorption spectroscopy in gas-turbine-engine environments. Compared with previous measurements in reciprocating engines and shock tubes, signal contamination from thermal emission was stronger in this combustor rig; methods for managing emission during experimental planning and postprocessing are discussed. H 2 O spectra spanning 1330―1380 nm (which includes the ν 1 + ν 3 and 2ν 1 overtone bands) are presented along with a method for calculating gas temperatures from the spectra. The resulting temperatures are reported for a variety of combustor conditions. These tests show promise for simple gas-turbine sensors and potential for more detailed experiments involving tomographic reconstruction or multispecies concentration measurements.