Dennis S. Fox, Robert A. Miller, Dongming Zhu, and Michael Perez

Glenn Research Center, Cleveland, Ohio

Michael D. Cuy

ASRC Aerospace Corporation, Cleveland, Ohio

R. Craig Robinson

Jacobs Technology Inc., Cleveland, Ohio

Mach 0.3 Burner Rig Facility at the NASA Glenn

Materials Research Laboratory

NASA/TM—2011-216986

March 2011
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Dennis S. Fox, Robert A. Miller, Dongming Zhu, and Michael Perez

Glenn Research Center, Cleveland, Ohio

Michael D. Cuy

ASRC Aerospace Corporation, Cleveland, Ohio

R. Craig Robinson

Jacobs Technology Inc., Cleveland, Ohio

Mach 0.3 Burner Rig Facility at the NASA Glenn

Materials Research Laboratory

NASA/TM—2011-216986

March 2011

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors would like to acknowledge the efforts of others that provided key contributions to the facility throughout its 

history: Carl A. Stearns, James (Jimmy) R. Johnston, Thomas Schneider, Dennis L. Weismantel, Steve DeBarr,

Richard D. Rinehart, R. Craig Robinson, Robert T. Pastel, Ruben J. Ramos, Eric W. Sockel,

Daniel L. Deadmore, Carl E. Lowell, Richard L. Ashbrook, and Leonard J. Bunyak.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-216986 1 

Mach 0.3 Burner Rig Facility at the NASA Glenn 

Materials Research Laboratory 

Dennis S. Fox, Robert A. Miller, Dongming Zhu, and Michael Perez 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Michael D. Cuy 

ASRC Aerospace Corporation 

Cleveland, Ohio 44135 

R. Craig Robinson 

Jacobs Technology Inc. 

Cleveland, Ohio 44135

Abstract 

This Technical Memorandum presents the current capabilities of the state-of-the-art Mach 0.3 Burner 

Rig Facility. It is used for materials research including oxidation, corrosion, erosion, and impact. 

Consisting of seven computer controlled jet-fueled combustors in individual test cells, these relatively 

small rigs burn just 2 to 3 gal of jet fuel per hour. The rigs are used as an efficient means of subjecting 

potential aircraft engine/airframe advanced materials to the high temperatures, high velocities and thermal 

cycling closely approximating actual operating environments. Materials of various geometries and 

compositions can be evaluated at temperatures from 700 to 2400 °F (370 to 1316 °C). Tests are conducted 

not only on bare superalloys and ceramics, but also to study the behavior and durability of protective 

coatings applied to those materials. 

1.0 Introduction 

The state-of-the-art Mach 0.3 Burner Rig Facility is located within the Materials Research Laboratory 

(MRL, building 34) at the National Aeronautics and Space Administration’s Glenn Research Center 

(GRC). It is used for materials research including oxidation, corrosion, erosion, and impact. Consisting of 

seven computer controlled jet-fueled combustors in individual test cells, these relatively small rigs burn 

just 2 to 3 gal of jet fuel per hour. The name of the facility comes from the velocity of the exhaust gas 

stream (~335 ft/sec, or ~102 m/s). The rigs are used as an efficient means of subjecting potential aircraft 

engine/airframe advanced materials to the high temperatures, high velocities and thermal cycling closely 

approximating actual operating environments. Materials of various geometries and compositions can be 

evaluated at temperatures from 700 to 2400 °F (370 to 1316 °C). In addition, materials can be subjected 

to thermal cycling, which duplicates the flight cycles experienced by aircraft making daily takeoffs and 

landings. This facility is primarily used in test programs for NASA Fundamental Aeronautics Programs. 

Additional research has been conducted with aircraft gas turbine engine companies including General 

Electric Aviation, Pratt & Whitney, Rolls-Royce, LibertyWorks (formerly Allison Advanced Development 

Co.), Honeywell Aerospace (formerly AlliedSignal Engines) and Williams International; the United States 

Air Force, Navy and Army; various material suppliers; other NASA field centers; and academia. 

The primary focus of this Technical Memorandum is to present the current capabilities of the facility. 

The backbone of each facility—the burner itself—is constant throughout the facility. However, the 

materials test configuration is unique in each test cell. Current capabilities include cyclic oxidation 
NASA/TM—2011-216986 2 

(hot-cold, or idle-takeoff-cruise), corrosion, erosion, impact, and tensile loading. Tests are conducted not 

only on bare superalloys and ceramics, but also to study the behavior and durability of protective coatings 

applied to those materials. 

The secondary focus of this memorandum is to document past studies conducted in the facility, 

primarily through an extensive listing of papers in the bibliography. The Appendix summarizes the work 

conducted using more unique NASA GRC burner rigs. The high pressure burner rig is still in operation, 

while the 4-atmosphere corrosion rig and Mach 1 rig have been dismantled. Finally, some archival photos 

are presented. The Mach 0.3 burner rig facility is located within the Durability & Protective Coatings 

Branch: http://www.grc.nasa.gov/ WWW/StructuresMaterials/DPC/, part of the Structures & Materials 

Division: http://www.grc.nasa.gov/WWW/StructuresMaterials/, in the Research and Technology 

Directorate: http://rt.grc.nasa.gov/. 

2.0 Description and Operation of Mach 0.3 Burner Rig 

A schematic representation of a standard Mach 0.3 burner rig is presented in Figure 1. Note that the 

standard exhaust nozzle diameter is 1 in. (2.5 cm) in diameter, though other nozzles have been used 

previously. A detailed description of burner operation is found in References 1 and 2, and it is 

summarized here. Maximum sample temperature that can be achieved on a single fixed sample is 

nominally 2400 F (1316 C). Each rig utilizes 120 psig (800 kPa) filtered shop air supplied via the GRC 

central air system. Air flow is measured with Sponsler (Lake Bluff, IL) type 2503 turbine flow meters 

(electronic output to computer) and Fisher & Porter (Warminster, PA) Precision Bore rotometers (visual 

float-type air flow indication on rig). To run at Mach 0.3 approximately 2.5 lb/min airflow is used. For 

higher Mach numbers up to 7.0 lb/min airflow is available. Air is preheated to 500 F (260 C) to 

Figure 1.—Mach 0.3 burner rig cross-sectional schematic. 
NASA/TM—2011-216986 3 

minimize coking within the combustor can, and it is delivered through a 0.75 in. (1.9 cm) diameter line 

and forced through a co-rotating swirl plate. Jet fuel is provided from a 5000 gal (19,000 liter) 

underground storage tank. A low pressure fuel pump (35 psig) delivers filtered fuel to the building. A 

high pressure pump (350 psig) then delivers fuel to each test cell, where it is injected into the combustor 

using off-the-shelf Goodrich-Delavan oil burner fuel nozzles located in the center of the swirl plate. The 

most commonly used fuel nozzle delivers 1.5 gph, with others on-hand ranging from 0.5 to 3.5 gph. 

Type “A” nozzles have a solid cone of spray particles, whereas Type B’s spray pattern is only located on 

the perimeter of the cone. The nozzles come in cone spray patterns from 45 to 80, with an 80 “hollow” 

Type “B” cone providing the best overall combustion performance (e.g., soot-free flame, minimal coking 

of the combustor can). 

Pressure within the combustor is measured with a transducer, and is usually held 1 psi above ambient. 

This ∆P determines the Mach 0.3 flame velocity. An aircraft-type igniter initiates combustion. The inner 

liner is made of Inconel 601 superalloy, with its outer diameter cooled with bypass air. Downstream this 

air is added to the combustion process through liner perforations. A Type “K” thermocouple is used to 

confirm the presence of a flame. For corrosion studies, an optional salt-solution injector can be used. 

Sample temperature can be measured in a variety of ways. Routinely, an optical pyrometer is used as 

illustrated in Figures 1 and 2. Another monitoring technique is the use of a video camera recording 

system. Since the rigs are employed around-the-clock, video can be used to capture certain events 

occurring outside of first shift, such as spallation of a thermal barrier coating. There are currently seven 

burner rigs in the facility, each housed separately in a test cell that has a 10 by 10 ft (3 by 3 m) footprint 

and is 12 ft (~3.7 m) high. 

Figure 2.—Mach 0.3 burner used in thermal blanket durability study. 

(NASA photo C-1994-5005, November 1994, from Ref. 33). 
NASA/TM—2011-216986 4 

The Mach 0.3 burners at GRC are based on a Pratt & Whitney (East Hartford, CT) design. The 

earliest ones came on line in the early 1970s. At first, the rigs were operated in a constantly attended 

mode. With the advent of personal computers (PCs) in the early 1980s, digital control of the burner rigs 

became viable. This made them highly flexible and provided accurate means of 24/7 unattended material 

durability testing. A Technical Memorandum (TM) containing detailed information of both Mach 0.3 

burner hardware and the first version of digital temperature and velocity control (including computer 

code) was published in 1985 (Ref. 2). Figure 3 is a simple schematic diagram of the burner rig control 

system hardware from that paper. The Mach 0.3 Burner Rig Operator’s Manual (Ref. 3) was published in 

1987, and Figure 4 is a promotional viewgraph with both authors pictured. 

Figure 3.—Schematic diagram of the burner rig control system hardware (from Ref. 3). 
NASA/TM—2011-216986 5 

Figure 4.—Promotional viewgraph (circa 1987) for the PC controlled burner rigs. Note the IBM PC with 

two 5.25 in. floppy disk drives. Reference 3 authors shown. 

The computer control system used today was installed in 1987. As of this writing, antiquated but 

rugged 486-based machines are still used to run a DOS-based program with version 1.1 installed in July 

of 1987. The control system is a hybrid—both computer (digital) and electromechanical (analog). Via a 

suitable interface, the PC acquires specimen temperature (usually via an optical pyrometer), fuel pressure, 

fuel flow, burner can pressure, and other rig parameters. Specimen temperature is then controlled by 

digital adjustment of fuel flow using a PID control algorithm. Also included in the computer control 

functions are safety features such as emergency shutdowns, test condition measurements, record keeping 

of test data, display of testing status, and mathematical calculations. 

Five different test modes can be selected when starting each test: 

1. Pyrometer/Fuel—This mode is used the majority of the time. It maintains a specific sample 

surface temperature (±6 C/±10 F), as measured with the pyrometer, by adjusting fuel flow. 

2. Fuel/Air ratio—This mode maintains a specific fuel-to-air ratio throughout the test, with regard to 

the desired temperature of a thermocoupled dummy sample. The software adjusts temperature of 

the dummy (minimum 60 sec needed) and then pivots the burner onto the actual sample. 

3. Thermocouple/Fuel—This mode maintains a specific sample temperature (±6 C/±10 F; sample 

is thermocoupled) by adjusting fuel flow. 

4. Thermocouple/Air—In this mode, the temperature of a hollow thermocoupled specimen is held 

steady by adjusting cooling air flow through the sample. A preselected fuel-to-air ratio is 

maintained throughout the test. 

5. Pyrometer/Air—In this mode, the temperature of a hollow specimen as measured with an optical 

pyrometer is held steady by adjusting cooling air flow through the sample. A preselected fuel-toair ratio is maintained throughout the test. 
NASA/TM—2011-216986 6 

Figure 5.—Top left: Assortment of carousels used for cylindrical, rectangular, and 

occasional odd-shaped samples (such those in Fig. 20). Top right: Cylindrical sintered 

α-silicon carbide samples in test carousel (NASA GRC photo C-1991-10893). Bottom: 

Superalloy sheets in carousel test (C-1999-2482). 

Another test mode choice is that of thermal cycling, wherein the burner can be pivoted on-and-off the 

sample a specified number of times. Also, forced air cooling can be applied to the specimen when the 

burner is pivoted off of the sample(s). Single samples are fixed in place and impinged on by the flame, 

with the impingement angle specified by the customer. If a number of specimens are to be tested, a 

spinning sample carousel (100 to 200 rpm) is utilized (see Fig. 5). Carousels are fabricated in-house 

according to customer needs. It should be noted that a single specimen can also be rotated in the flame, if 

desired. 

In Figure 6 are photographs of the Mach 0.3 burner rig facility control room and one of the test cells. 
NASA/TM—2011-216986 7 

Figure 6.—Top: Mach 0.3 burner rig facility control room. Each of the seven 

rigs is housed in individual test cells behind the operator as he sits. 

Bottom: Looking through the soundproof door into one of the test cells. 
NASA/TM—2011-216986 8 

3.0 Current Research and Specific Test Cell Configurations 

3.1 3 by 3 in. Plate Oxidation 

Superalloys are metals used at high temperatures in the hot sections of gas turbine engines. This class 

of material retains significant strength to temperatures near 1830 °F (1000 °C). An understanding of the 

oxidation and hot corrosion behavior of superalloys is of great importance. In this demanding application, 

components such as turbine blades and vanes are often protected with Thermal Barrier Coatings (TBCs). 

These thin 5 to 10 mil (0.125 to 0.25 mm) zirconia-based ceramic coatings, when coupled with internal 

air cooling, allow the components to be operated at temperatures 100 to 200 °F higher than bare 

components. 

In this test cell, square oxidation samples (coated or uncoated) are held at a 45 angle to the flow 

direction and held in a superalloy “picture frame.” The burner is fixed in place (doesn’t pivot). Via 

electromechanical actuation, the sample can be moved close to the 1 in. diameter burner exit nozzle to 

model the high temperatures encountered during full power engine conditions such as takeoff, far from 

the nozzle to mimic engine idle conditions, and to an intermediate position to model cruise conditions. 

The sample can be rapidly translated (within seconds) from 1 to 18 in. away from the burner nozzle. 

Some customers specify up to three positions, with the corresponding temperatures at those positions 

relating to idle, cruise and takeoff/landing conditions. Preferred sample size is 3 by 3 in. (7.6 by 7.6 cm) 

as shown in Figure 7, though 2 by 2 in. (5 by 5 cm) and 4 by 4 in. (10.2 by 10.2 cm) samples have also 

been tested. 

Figure 7.—Durability testing of a 3 by 3 in. superalloy plate protected with a thermal 

barrier coating. 
NASA/TM—2011-216986 9 

Figure 8.—Exposure of a 2 by 2 in. carbon fiber reinforced silicon carbide matrix 

composite “sandwich” to a top face temperature of 2300 F. Heat transfer across 

the panel is measured using an optical pyrometer on the top face and two Type “K” 

thermocouples on the back face. 

Ceramics such as silicon carbide and silicon nitride are generally stable to higher temperatures than 

most metals and alloys. These materials form a stable film of silica (SiO2) in an oxidizing environment. 

The development of high temperature structural ceramics for advanced gas turbine engines has been an 

area of active research for many years. In such applications, ceramics are exposed not only to high 

temperatures but also aggressive gases (including water vapor) and possibly deposits (sea salt). Water 

vapor attack of the silica scale results in paralinear oxidation and recession of the substrate at 

unacceptable rates. Environmental Barrier Coatings (EBCs) have been developed to address this problem. 

In the case of sea salt hot corrosion, the concern of corrosion is loss of load bearing capability due to 

pitting of the surface. Oxidation studies of silicon-based ceramics are also conducted with this plate 

oxidation rig, an example shown in Figure 8. 

3.2 Corrosion 

Hot corrosion is an accelerated attack experienced by superalloys and silicon-based ceramics that is 

initiated by the deposition of condensed sodium sulfate (Na2SO4) on engine parts. The source of the salt is 

direct ingestion with intake air or by its formation during combustion from sodium chloride-contaminated 

air and sulfur in the fuel. Most superalloys/ceramics are susceptible to hot-corrosion attack to some extent. 

The phenomenon depends on many factors, including temperature, cyclic conditions, material composition, 

impurities in and pressure of air ingested into the engine, and impurities in the fuel. In this test, samples are 

normally held in a sample carousel. The one illustrated in Figures 9 and 10 holds 18 superalloy pins 

nominally 0.25 in. diameter by 3 in. tall (0.64 cm diameter by 7.6 cm tall). A synthetic sea salt solution 

(mixed in-house) is pumped through an air atomizer and injected into the combustion can as shown in the 

Figure 1 schematic. Solutions in the 0.1 to 10 parts-per-million range have been used; higher concentrations 

tend to clog the atomizer. 
NASA/TM—2011-216986 10 

Figure 9.—Carousel holding eighteen specimens, each 1/4 

diameter by 3 in. tall. Samples were undergoing synthetic 

sea salt corrosion to determine the effectiveness of protective 

metallic coatings. 

Figure 10.—Results of the above 900 C (1650 F), 500 hr 

corrosion test with 8 ppm synthetic sea salt. The most 

corrosion resistant samples are located at the forefront. 

3.3 Impact Gun and TBC Particle Impact 

A gas turbine engine ingests enormous amounts of air. At the same time, it can also suck in foreign 

objects such as dust, dirt, volcanic ash, sand, rocks, ice, hail, and man-made objects (e.g., nuts, bolts). 

This can cause foreign object damage (FOD), a problem so severe that a technical society is devoted to 

the issue (National Aerospace FOD Prevention, Inc., http://www.fodnews.com/). An impact gun based on 

a small particle gun first developed in the Ballistic Impact Laboratory (http://rt.grc.nasa.gov/main/rlc/ 

ballistic-impact-laboratory/) is used to fire 1/16 in. diameter steel balls at samples heated by the burner rig 

(Figs. 11 and 12). Velocities range from 50 m/sec to a maximum of 400 m/sec. Studies of the impact 

resistance of both uncoated and environmental barrier coated SiC/SiC composites using the small particle 

gun are found in References 40 and 41. 
NASA/TM—2011-216986 11 

Figure 11.—Top: The small cylinder is filled with helium from the “K” bottle, 

and the “Red Hat” valve is actuated to accelerate the round projectile toward 

the sample down a hypodermic metal tube (1/16 in. ID) inserted in a larger 

diameter (1/8 in. ID) stainless steel barrel. Bottom: Business end of impact 

gun. Exit of barrel is protected from the flame with ceramic insulation. In this 

case, a superalloy target sample is held in a “C” clamp arrangement. 
NASA/TM—2011-216986 12 

Figure 12.—Uncoated CMC tensile “dogbone” specimens (6 by 1/2 by 1/8 in.) after impact with a 

1/16 in. diameter steel ball bearing at various temperatures and projectile velocities. 

Aircraft gas turbine engines are subject to damage from erosion (generally caused by particles tens of 

microns in diameter) or impact damage (generally from particles hundreds of microns in diameter). This 

rig is also being modified to allow another type of impact testing, similar to the rig discussed in 

Section 3.4. The impact gun will be temporarily moved, and 560 μm particles will be added to the 

combustor can to model particle impact (as opposed to erosion) of TBC coated specimens. In the rig 

described in the next Section 3.4 smaller 50 to 100 μm alumina grit is added to the combustor can for 

TBC erosion studies that model sand ingestion by helicopter turbine engines. 

3.4 Erosion 

Considerable advances in engine performance are required over the next several decades. Advanced 

TBCs will be an important part of all future gas turbine engines. As stated in the previous section, aircraft 

gas turbine engines are subject to ingested dirt/dust/sand damage from erosion (generally caused by 

particles tens of microns in diameter) or impact damage (generally from particles hundreds of microns in 

diameter). Particulate ingestion may be more severe at low altitudes, from hovering of rotorcraft during 

takeoff and landing, and from operation under harsh conditions. The shedding of particles from the 

combustor is another important source of particulate matter. Unless future TBCs are erosion resistant, 

coating loss from erosion and impact damage will be a factor limiting future advancements. Reference 4 

describes current turbine blade TBC development at GRC, as well as the study of the coating’s impact 

and erosion resistance validated in laboratory-simulated engine erosion and/or thermal gradient 

environments using the unique Mach 0.3 burner rig described here. 
NASA/TM—2011-216986 13 

A standard Mach 0.3 burner rig was modified to allow Mach 0.3 to 1.0 velocities and erosion 

capabilities to study advanced turbine blade TBCs in a simulated turbine engine-relevant erosion 

environment (Figs. 13 to 15). Rig development emphasized increasing the erodent particle velocities at 

high temperatures by increased gas mass flow rate, erodent flow uniformity with improved feeding 

systems, and relevant thermal gradients found in turbine environments. The high velocity burner erosion 

rig includes a 1.9 mm (3/4 in.) diameter highly efficient burner nozzle, a specimen holder fixture and 

precision erodent feeder. 

The burner nozzle inner surface contour was designed based on an ANSI/ASME nozzle standard to 

achieve better flame stability and uniformity, with the overall configuration modified to accomodate 

increased burner mass flow and higher heat flux environments. The nozzle was made of a single crystal 

nickel-base superalloy turbine blade material to ensure high temperature durability. 

To achieve high erosion particle velocities, a computational fluid dynamics modeling (CFD) approach 

in conjunction with experimental investigations was employed to optimize the burner and erodent 

injection design. The CFD model (Fluent/ANSYS; Canonsburg, PA) was used to calculate the gas and 

particle velocities for the burner rig temperatures and pressures of interest. The modeling and 

experimental testing were conducted in relevant burner velocities ranging from Mach 0.3 to 0.9, 

representative of turbine engine conditions. Alumina (Al2O3) particles were used as the erodent and 

nominal particle sizes ranged from 27 to 560 μm in order to understand the broad coating erosion and 

impact behavior. 

Figure 13.—Modified burner rig for Mach 0.3 to 1.0 erosion testing. The dust 

collection system is not shown. 
NASA/TM—2011-216986 14 

Figure 14.—2.5 cm diameter sample (in clamshell holder) exposed to erosion rig burner exhaust 

nozzle operating at Mach 0.5. 

 

Figure 15.—Left: 2.5 cm diameter superalloy button with improved TBC after erosion testing; Right: Digital image 

of eroded specimen and corresponding micron scale from contour tracer (Zygo, Middlefield, CT) illustrating 

penetration depth. 
NASA/TM—2011-216986 15 

A modified burner rig configuration with an extension duct (300 mm in length and 19 mm inner 

diameter) was explored to further accelerate the erodent particle and increase the velocity using 50 and 

560 μm size Al2O3 particles. Erosion rates are determined by interrupted coating erosion thickness 

recession and weight-loss measurements, and also by the amount of erodent required for penetration of 

the coating to expose the substrate. In Reference 4 the coating erosion evaluations was primarily 

conducted at 1800 and 2000 °F (coating surface ~1900 and ~2100 °F, respectively) using 27 μm size 

Al2O3 particles, with additional tests conducted at 2200 °F.

3.5 Tensile Loading 

In combustion environments, silicon-based ceramics are prone to enhanced recession due to the 

reaction of its protective silica (SiO2) scale with water vapor to form gaseous Si(OH)4. Environmental 

barrier coatings (EBCs) were developed to protect Si-based Ceramic Matrix Composites (CMCs) from 

such attack. This Mach 0.3 burner rig test cell contains a very simple mechanical test frame (Fig. 16) used 

to apply tensile loads primarily to ceramic matrix composites (bare or coated). Imposed crack formation 

allows penetration of the combustion gases to the substrate, allowing oxidation of the reinforcing fiber 

and interphase to occur. Metal samples have been tested in this rig as well. A very unique material studied 

in this rig was an Inconel 718 Lattice Block Structure. 

3.6 Thermal Barrier Coating Durability 

Life prediction studies of TBCs and their metallic bond coats - primarily as applied to solid 

superalloy rods - have been conducted in the past. These insulating coatings thermally protect 

components and allow higher operating temperatures. At the same time, component life is extended by 

reductions in oxidation and thermal fatigue. In conjunction with active film cooling, TBCs permit 

working temperatures higher than the melting point of the metal blade/vane. TBCs are applied to the 

superalloy substrate using the plasma spray technique. A powder such as zirconia is introduced into a 

plasma jet emanating from a torch, the temperature of which is on the order of 5000 F. The zirconia 

powder melts and is propelled towards a substrate, wherein molten droplets of the ceramic solidify to 

develop the desired coating. There are a large number of variables in this technique including powder 

type/size, powder manufacturer, plasma gas composition, gas flow rates, energy input, standoff distance 

from the torch to the substrate, and powder deposition rate. The life of the coating is a function of all of 

these variables, and the burner rig is an excellent tool to conduct lifing studies. 

Exposure of single rotating specimens can be conducted (Fig. 17), but a higher amount of sample 

throughput is achieved using an eight-place sample carousel (Fig. 18). Samples can be exposed for 

hundreds of hours of unattended thermal cycling. A video camera with time stamping can be triggered to 

take ~30 sec of video every time the burner pivoted on the sample set. A TBC delamination appears as a 

hot spot, and the rig operator can use the videotape to determine time of failure. The failed specimen can 

be replaced with a new specimen, and the carousel test continued. 

The application of thick thermal barrier coatings (TTBCs) to low heat rejection diesel engine 

combustion chambers was a subject of study in the early 1990s by companies such as Caterpillar 

(Peoria, IL) and Cummins Engine Company, Inc., (Columbus, IN). Areas of TTBC technology examined 

at the time included powder characteristics and chemistry; bond coat composition; coating design, 

microstructure and thickness as they affect properties, durability, and reliability; and TTBC “aging” 

effects (microstructural and property changes) under diesel engine operating conditions (Ref. 43). 

Figure 19 is a photo of burner rig testing of such coatings, circa 1997. 
NASA/TM—2011-216986 16 

Figure 16.—Tensile test in conjunction with burner rig exposure. Left, coated CMC tensile “dogbone” specimen. 

Right, uncoated Si-based CMC. 

Figure 17.—Burner rig exposure of a single spinning thermal barrier coated superalloy. 

Nominal sample size of solid rod: 1/2 in. diameter, 5 in. tall. GRC photo numbers 

C-1999-2485 and C-1999-2487. 
NASA/TM—2011-216986 17 

Figure 18.—Eight place carousel for burner rig 

exposure of a multiple thermal barrier coated 

superalloy specimens. Nominal size of coated 

solid rod: 1/2 in. diameter, 5 in. tall. 

Figure 19.—Burner rig exposure of diesel thermal barrier 

coatings (NASA photo C-1997-2825). 
NASA/TM—2011-216986 18 

4.0 Historic NASA Aeronautics Program Research 

As noted, the main customers of the facility are the fundamental NASA aeronautics programs. In the 

mid 1990’s as part of the Enabling Propulsion Materials Program, it was used to support Task B – 

Exhaust Nozzle. The objective of this task was to provide the necessary material technology required for 

an economically viable and environmentally acceptable High Speed Civil Transport HSCT exhaust 

system (Ref. 33). Mach 0.3 burner rig tests were performed to evaluate the insulating capability of the 

thermal protection system, or TPS (see Fig. 2). After exposure, the samples were examined for any 

physical degradation, such as loss of flexibility or mechanical integrity. 

As part of the General Aviation Program in the late 1990s, the oxidation and corrosion of turbine 

engine blades and vanes manufactured by Williams International for the lightweight FJX-2 turbofan 

engine was studied (see Fig. 20). The testing was cyclic in nature, with each cycle consisting of 60 min in 

the flame and 10 min with the burner pivoted off the sample. Also, parts-per-million amounts of a 

synthetic sea salt solution were injected into the burner can for the first few minutes of the 1-hr hot cycle. 

This engine was to power General Aviation Program’s V-Jet demonstrator test bed aircraft. A description 

of the program is found here: http://www.docstoc.com/docs/1084022/The-General-Aviation-Propulsion-

(GAP)-Program. 

Figure 20.—Oxidation tests of engine vanes for General Aviation Propulsion 

Program (NASA photo C-28-270). 

 
NASA/TM—2011-216986 19 

Appendix—Other Unique NASA GRC Burner Rigs 

The High Pressure Burner Rig (HPBR) is currently housed in the Special Projects Laboratory 

(building 24) and is a combustion test rig used for high-temperature environmental durability studies of 

materials for advanced aircraft (Fig. 21). The facility burns jet fuel and air in controlled ratios to produce 

combustion gas chemistries and temperatures that are similar to those in gas turbine engines. In addition, 

the test section is also capable of simulating the pressures and gas velocities representative of today’s 

aircraft. A brief summary is given here, with a detailed test facility description and history of 

development is found in References 44 and 45 are summarized here. 

The HPBR typically operates with a fuel-lean gas mixture. However, with the proper scrubber and 

waste disposal systems, the facility can also be operated using fuel-rich gas mixtures. Test samples are 

easily accessible for ongoing inspection and documentation of weight change, thickness, cracking, and 

other metrics. Temperature measurements are available in the form of both Types “B” and “R” 

thermocouples as well as 2-color and 8 m optical pyrometers. The facility is also equipped with quartz 

windows for observation and videotaping. The facility has a wide range of operating conditions including: 

up to 2.0 lbm/sec airflow (combustion and/or secondary cooling); equivalence ratios 0.5 to 1.0 (leanburn); gas temperatures of 1500 to 3000 °F; test pressure range of 4 to 12 atmospheres; and gas flow 

velocities of 30 to 100 ft/sec. 

Figure 21.—High Pressure Burner Rig (HPBR). NASA photo C-1996-314, January 1996. 

. 
NASA/TM—2011-216986 20 

The HPBR is typically used as a materials test rig best suited for testing various sized coupons, bars, 

and disks of candidate materials, including metals, ceramics, and plasma sprayed coatings. As an 

example, HPBR data has enabled researchers to make major contributions in our knowledge of advanced 

materials such as silicon-based monolithic and composite ceramics (Refs. 46 to 48). In these experiments 

the effects of water vapor and scale volatility were studied on coupon-type test samples to provide 

oxidation mechanisms. However, complex-shaped components can also be accommodated including 

cylinders, airfoils, vanes, and film-cooled end walls. Examples of sub-component test samples are found 

in References 49 and 50. The facility has also been used to evaluate thin-film temperature measurements, 

microwave sensors, and embedded thermocouples. 

http://www.lerc.nasa.gov/WWW/EDB/Facilities/high_pressure_burner_facility.htm 

A now defunct natural gas/compressed air burner (Figs. 22 and 23) was used to produce velocities up 

to Mach 1 and specimen temperatures up to 2000 °F/1093 °C (Refs. 51 to 59). The burner was designed 

to burn natural gas with compressed air up to approximately 55 psia (0.038 MN/m ) with exit gas 

temperatures as high as 3000 °F (1649 °C). A double liner was used to permit the combustion air to 

efficiently cool the outer jacket as well as the flame tube. Water cooling was used only on the converging 

exit nozzle. Accordingly, heat losses were minimized and the resulting fuel/air ratio was as lean as 

possible. 

The now defunct Hot Corrosion Test Facility (Figs. 24 and 25) - also known as the 4-Atmosphere 

Burner Rig - at the NASA Lewis Special Projects Laboratory (SPL, or building 24) was a high-velocity, 

pressurized Mach 0.3 burner rig used to evaluate the hot corrosion of advanced ceramic materials such as 

silicon carbide and silicon nitride (Refs. 60 to 64). A salt water solution (2 to 5 parts per million) was 

injected into the burner flame. Maximum air flow was 250 kg/hr (550 lbm/hr), test pressure range was 

100 to 600 kPa (1 to 6 atm), and maximum gas temperature exceeded 1500 °C (2732 °F). The test 

samples in this rig were normally held at 400 kPa (60 psi, or 4 atm) since deposition of salt and 

subsequent corrosion is a function of total pressure dew point. The rig was dismantled in 2001. 

Figures 26 to 30 are a collection of archival photographs.

 

Figure 22.—Mach 1 burner rig. NASA photo C-1969-3870, 

November 1969. 

Figure 23.—Mach 1 burner rig. NASA photo 

C-1977-1605. 
NASA/TM—2011-216986 21 

Figure 24.—Cross section drawing of 4-atmosphere corrosion burner rig. (CD-93-63596). 

Figure 25.—4-atmosphere corrosion rig circa 1992 (C-92-6666). 
NASA/TM—2011-216986 22 

Collection of Archival Photographs

Figure 26.—Mach 0.3 burner rig in operation circa 1979 

(C-1979-2390). 

Figure 27.—Mach 0.3 burner rig in operation circa 1982 (C-1982-6121). 
NASA/TM—2011-216986 23 

Figure 28.—Mach 0.3 burner rig deposition experiment 1982 (C-1982-

6122). Prior to 1985, there were just four Mach 0.3 burner rigs, and 

they were located in cell 4 of building 24 (Special Projects Laboratory). 

As shown above, the rigs were in pairs on two different tables. In the 

course of oxidation and corrosion testing, this arrangement was found 

to be unsatisfactory in that it allowed “cross-talk” between rigs and 

skewed experimental results. This is the reason behind each rig being 

currently housed in its own separate cell in building 34/MRL. 

Figure 29.—Oxidation of coated superalloy specimen circa 1986 

(C-1986-4728). 
NASA/TM—2011-216986 24 

Figure 30.—Burner rig/personal computer publicity photo circa 1990 

(C-1990-136). 

References 

Notes: Beginning with Reference 5, all papers through number 43 relate to historic Mach 0.3 studies 

and are listed in chronological order. References 44 to 50 relate to the high pressure burner rig. 

References 51 to 59 relate to the defunct Mach 1 burner rig. References 60 to 64 relate to the defunct 4-

atmosphere corrosion burner rig. Also, many of the historic papers in this technical memorandum can be 

found on the NASA Scientific and Technical website at http://www.sti.nasa.gov/nasaonly/resrch.html. 

1. Robinson, R.C.; Cuy, M.D., “Hot Corrosion Test Facility at the NASA Lewis Special Projects 

Laboratory,” NASA-CR-195323 May 1, 1994. 

2. Deadmore, D.L., “Digital Temperature and Velocity Control of Mach 0.3 Atmospheric Pressure 

Durability Testing Burner Rigs in Long Time, Unattended Cyclic Testing,” NASA TM-86959, 

March 1, 1985. 

3. Rinehart, R.D. and Cuy, M.D., “Mach 0.3 Burner Rig Operator’s Manual,” Sverdrup Technology, 

Inc. Contract NAS3-24105, October 7, 1987. 

4. Zhu, D., Miller, R.A.., and Kuczmarski, M.A., “Development and Life Prediction of Erosion 

Resistant Turbine Low Conductivity Thermal Barrier Coatings,” NASA/TM—2010-215669, 

February 2010. 

5. Lowell, C.E., Deadmore, D.L., “Effect of a Chromium-Containing Fuel Additive on Hot Corrosion,” 

NASA-TM-X-73465, July 1, 1976. 
NASA/TM—2011-216986 25 

6. Deadmore, D.L. and Lowell, C.E., “Burner Rig Alkali Salt Corrosion of Several High Temperature 

Alloys.” NASA TM-X-73659, May 13, 1977. 

7. Lowell, C.E., Deadmore, D.L., “Effect of a Chromium-Containing Fuel Additive on Hot Corrosion,” 

Corrosion Science, 18 [8] pp. 747–763, January 1, 1978. 

8. Deadmore, D.L., Lowell, C.E., Kohl, F.J., “The Effect of Fuel-To-Air Ratio on Burner-Rig Hot 

Corrosion,” NASA TM-78960, July 1, 1978. Also in Corrosion Science, 19, pp. 371–378. 

9. Deadmore, D.L., Lowell, C.E., “Inhibition of Hot Salt Corrosion by Metallic Additives,” NASA-TM78966, July 1, 1978. 

10. Miller, R.A., “Analysis of the Response of a Thermal Barrier Coating to Sodium and Vanadium 

Doped Combustion Gases,” NASA-TM-79205, January 1, 1979. 

11. Deadmore, D.L., Lowell, C.E., “Airfoil Cooling Hole Plugging By Combustion Gas Impurities of 

The Type Found In Coal Derived Fuels,” NASA-TM-79076, February 1, 1979. 

12. Hodge, P.E., Miller, R.A., Gedwill, M.A., “Evaluation of Hot Corrosion Behavior of Thermal Barrier 

Coatings,” NASA-TM-81520, January 1, 1980. 

13. Lowell, C.E., Sidik, S.M. and Deadmore, D.L., “Effect of Sodium, Potassium, Magnesium, Calcium 

& Chlorine on High Temperature Corrosion of IN-100, U-700, IN-792 & Mar M-509,” NASA TM79309, January 1, 1980. 

14. Deadmore, D.L. and Lowell, C.E., “Effects of Impurities in Coal-Derived Liquids on Accelerated Hot 

Corrosion of Superalloys,” NASA TM-81384, March 1, 1980. 

15. Hodge, P.E., Miller, R.A., Gedwill, M.A., “Evaluation of the Hot Corrosion Behavior of Thermal 

Barrier Coatings, “in Metallurgical coatings 1980, Proceedings of the Seventh International 

Conference, San Diego, CA, April 21–25, 1980. Volume 2. (A82-17251 05-23) Lausanne, Elsevier 

Sequoia, S.A., 1980, pp. 447–453. 

16. Santoro, G.J., Kohl, F.J., Stearns, C.A., Fryburg, G.C., Johnson, J.R., “Deposition and Material 

Response From Mach 0.3 Burner Rig Combustion of SRC 2 Fuels,” NASA-TM-81634, October 1, 

1980. 

17. Levine, S.R., Miller, R.A., Hodge, P.E., “Thermal Barrier Coatings for Heat Engine Components,” 

SAMPE Quarterly, 12, October, 1980, pp. 20–26. 

18. Lowell, C.E., Deadmore, D.J., Santoro, G.J., Kohl, F.J., “The Effects of Trace Impurities in CoalDerived Liquid Fuels on Deposition and Accelerated High Temperature Corrosion of Cast 

Superalloys,” NASA TM-81678, January 1, 1981. 

19. Hodge, P.E., Miller, R.A., Gedwill, M.A., Zaplatynsky, I., “Review of NASA Progress in Thermal 

Barrier Coatings for Stationary Gas Turbines,” NASA-TM-81716 January 1, 1981. 

20. Lowell, C.E., Sidik, S.M., Deadmore, D.L., “High Temperature Alkali Corrosion in High Velocity 

Gases,” NASA-TM-82591, May 1, 1981. 

21. Miller, R.A., Lowell, C.E, “Failure Mechanisms of Thermal Barrier Coatings Exposed To Elevated 

Temperatures,” NASA-TM-82905, January 1, 1982. 

22. Lowell, C.E., Deadmore, D.L., Whittenberger, J.D., “Long-Term High-Velocity Oxidation and Hot 

Corrosion Testing of Several NiCrAl and FeCrAl Base Oxide Dispersion Strengthened Alloys,” 

Oxidation of Metals, vol. 17, pp. 205–221, April, 1982. 

23. Levine, S.R., Miller, R.A., “Thermal-Barrier Coatings for Utility Gas Turbines,” NASA TM-85349, 

September 1, 1982. 

24. Deadmore, D.L., Lowell, C.E. Burner Rig Study of Variables Involved in Hole Plugging of Air 

Cooled Turbine Engine Vanes,” NASA-TM-83308, February 1, 1983. 

25. Santoro, G.J., Gokoglu, S.A., Kohl, F.J., Stearns, C.A., Rosner, D.E., “Deposition of Na2SO4 From 

Salt-Seeded Combustion Gases of a High Velocity Burner Rig,” NASA-TM-83751, January 1, 1984. 

26. Deadmore, D.L., “Application of Induction Coil Measurements to the Study of Superalloy Hot 

Corrosion and Oxidation,” NASA-TM-83560, January 1, 1984. 

27. Santoro, G.J., Kohl, F.J., Stearns, C.A., Gokoglu, S.A., Rosner, D.E., “Experimental and Theoretical 

Deposition Rates from Salt-Seeded Combustion Gases of a Mach 0.3 Burner Rig,” NASA-TP-2225, 

March 1, 1984. 
NASA/TM—2011-216986 26 

28. Deadmore, D.L., “In-Situ Measurements of Alloy Oxidation/Corrosion/Erosion Using a Video 

Camera & Proximity Sensor with Microcomputer Control,” NASA TM-83673, May 1984. 

29. Deadmore, D.L., “Effects of Alloy Composition on Cyclic Flame Hot-Corrosion Attack of Cast 

Nickel-Base Superalloys at 900 °C,” NASA TP-2338, July 1, 1984. 

30. Stearns, C.A., Deadmore, D.L., Barrett, C.A. Effect of Alloy Composition on the Sodium-Sulfate 

Induced Hot Corrosion Attack of Cast Nickel-Base Superalloys at 900 C,” in the proceedings of 

Alternate Alloying for Environmental Resistance, held in New Orleans, LA, Mar. 2–6, 1986. 

Warrendale, PA, Metallurgical Society, Inc., 1987, pp. 131–143. 

31. Brindley, W.J., Miller, R.A., “Thermal Barrier Coating Life And Isothermal Oxidation of LowPressure Plasma-Sprayed Bond Coat Alloys,” in Metallurgical Coatings and Thin Films 1990,

Proceedings of the 17th International Conference on Metallurgical Coatings and 8th International 

Conference on Thin Films, San Diego, CA, April 2–6, 1990. vol. 1 (A92-32376 12-23). London and 

New York, Elsevier Applied Science, 1990, pp. 446–457. 

32. Lei, J-F, Cuy, M.D., Wnuk, S.P., “Attachment of Free Filament Thermocouples for Temperature 

Measurements on CMC,” NASA TM-107488, June 1997. 

33. Enabling Propulsion Materials Program Annual Technical Progress Report, Volume 3: Task B – 

Exhaust Nozzle, pp. 142–145. January 31, 1998. 

34. Nesbitt, James A., “Hot Corrosion of Single-Crystal NiAl-X Alloys,” NASA TM 113128, May 1998. 

35. Nesbitt, James A., Darolia, Ram, Cuy, Michael D., “Burner Rig Hot Corrosion of a Single Crystal Ni48Al-Ti-Hf-Ga Alloy,” Mat. Res. Soc. Symp. Proc. vol. 522, pp. KK9.5.1 to 9.5.6, The Materials 

Research Society, 1999. 

36. Lei, J-F, Kiser, J.D., Singh, M. Cuy, M.D. Blaha, C.A., Androjna, D., “Durability Evaluation of a 

Thin Film Sensor System with Enhanced Lead Wire Attachments on SiC/SiC Ceramic Matrix 

Composites,” NASA/TM—2000-209917, March 2000. 

37. Nesbitt, James A., Harris, K., Helmink, R., and Erickson, G., “Burner Rig Hot Corrosion of Five NiBase Alloys Including Mar-M247,” Published in Corrosion 2001, pp. 01166/1-14, NACE 

International, Houston, TX (2001). 

38. Opila, E.J., Robinson, Raymond C., Cuy, Michael D., “High Temperature Corrosion of Silicon 

Carbide and Silicon Nitride in Water Vapor,” 10th International Conferences on Modern Materials 

and Technologies (CIMTEC 2002). 

39. Zhu, Dongming, Kuczmarski, Maria A., Miller, Robert A., Cuy, Michael D., “Evaluation of Erosion 

Resistance of Advanced Turbine Thermal Barrier Coatings,” January 22, 2007. 31st International 

Cocoa Beach Conference & Exposition on Advanced Ceramics and Composites, Daytona Beach, FL. 

40. Bhatt, R.T., Choi, S.R., Cosgriff, L.M., Fox, D.S., Lee, K.N., “Impact resistance of environmental 

barrier coated SiC/SiC composites,” Materials Science and Engineering A 476 (2008) 8–19. 

41. Bhatt, R.T., Choi, S.R., Cosgriff, L.M., Fox, D.S., Lee, K.N., “Impact resistance of uncoated SiC/SiC 

composites,” Materials Science and Engineering A 476 (2008) 20–28. 

42. Dongming Zhu, Robert A. Miller, and Maria A. Kuczmarski, “Development and Life Prediction of 

Erosion Resistant Turbine Low Conductivity Thermal Barrier Coatings,” NASA/TM—2010-215669, 

February 2010. 

43. M.B. Beardsley, “Thick thermal barrier coatings for diesel engines,” Journal of Thermal Spray 

Technology vol. 6 [2] 181-186 June 1997.

44. Stearns, C.A. and Robinson, R.C, “NASA Lewis Research Center Lean-, Rich-Burn Materials Test 

Burner Rig,” NASA CR 194437, February 1994. 

45. Robinson, R. Craig, “NASA GRC’s High Pressure Burner Rig Facility and Materials Test 

Capabilities,” NASA/CR 19990209411, December 1999. 

46. Smialek, J.L., Robinson, R.C., Opila, E.J., Fox, D.S., Jacobson, Nathan S., “SiC and Si3N4 Recession 

Due to SiO2 Scale Volatility under Combustor Conditions,” NASA/TP-1999-208696, July 1999. 

47. R.C. Robinson and J.L. Smialek, “SiC Recession Caused by SiO2 Scale Volatility under Combustion 

Conditions. Part I: Experimental Results and Empirical Model,” J. Am. Ceram. Soc., 82 [7] 1817-25 

(1999). 
NASA/TM—2011-216986 27 

48. E.J. Opila, J.L. Smialek, R.C. Robinson, D.S. Fox, and N.S. Jacobson, “SiC Recession Caused by 

SiO2 Scale Volatility under Combustion Conditions. Part II: Thermodynamics and Gaseous Diffusion 

Model,” J. Am. Ceram. Soc., 82 [7] 1826-34 (1999). 

49. R.C. Robinson and K.S. Hatton, “SiC/SiC Leading Edge Turbine Airfoil Tested Under Simulated Gas 

Turbine Conditions,” NASA/CR-1999-209314, September 1999. 

50. Zhu, Dongming; Fox, Dennis S.; Pastel, Robert T., “High Pressure Burner Rig Testing of Advanced 

Environmental Barrier Coatings for Si3N4 Turbine Components,” 31st International Cocoa Beach 

Conference & Exposition on Advanced Ceramics & Composites, 22–26 January. 2007, Daytona 

Beach, FL. 

51. Ashbrook, R.L.; Johnston, J.R, “Oxidation and Thermal Fatigue Cracking of Nickel- and Cobalt-base 

Alloys in a High Velocity Gas Stream,” NASA-TN-D-5376, August 1969. 

52. Deadmore, D.L.; Grisaffe, S.J.; Sanders, W.A., “Furnace and High-Velocity Oxidation of AluminideCoated Cobalt Super-Alloy WI-52.” NASA TN-D-5834, May 1970. 

53. Deadmore, D.L., “Cyclic Oxidation of Cobalt-Chromium-Aluminum-Yttrium and Aluminide 

Coatings on IN-100 and VIA Alloys in High Velocity Gases,” NASA TN-D-6842, July 1972. 

54. Deadmore, D.L., “High-Velocity-Oxidation Performance of Metal-Chromium-Aluminum (MCrAl), 

Cermet, And Modified Aluminide Coatings on IN-100 and type VIA alloys at 1093 C,” NASA TND-7530, February 1974. 

55. Deadmore, D.L.; Lowell, C.E, “High Gas Velocity Oxidation and Hot Corrosion Testing of Oxide 

Dispersion-Strengthened Nickel-Base Alloys,” NASA-TM-X-71835, November 1, 1975. 

56. Lowell, C.E.; Deadmore, D.L., “High Velocity Oxidation and Hot Corrosion Resistance of Some 

ODS Alloys.” NASA TM-X-73656, April 1, 1977. 

57. Sanders, William A., Johnston, James R., “High Velocity Burner Rig Oxidation and Thermal Fatigue 

Behavior of Si3N4- and SiC-based Ceramics to 1370 C,” NASA TM-79040, November, 1978. 

58. Young, S.G. and Deadmore, D.L., “An Experimental, Low-Cost, Silicon Slurry/Aluminide HighTemperature Coating For Superalloys,” NASA TM-79178, July 1979. 

59. Young, S.G. and Deadmore, D.L, “An Experimental, Low-Cost, Silicon-Aluminide HighTemperature Coating For Superalloys,” NASA TM-81455, January, 1980. 

60. Robinson, R.C.; Cuy, M.D., “Hot Corrosion Test Facility at the NASA Lewis Special Projects 

Laboratory,” NASA-CR-195323 May 1, 1994. 

61. Jacobson, N.S.; Stearns, C.A.; Smialek, J.L., “Burner rig corrosion of SiC at 1000 °C,” Advanced 

Ceramic Materials, vol. 1, April 1986, pp. 154–161.

62. Fox, D.S. and Smialek, J.L, “Burner Rig Hot Corrosion of Silicon Carbide and Silicon Nitride,” 

J. Am. Ceram. Soc., 73 [2] 303-11 (1990). 

63. Fox, D.S., Cuy, M.D. and Strangman, T.E., “Sea Salt Hot Corrosion and Strength of a YttriaContaining Silicon Nitride,” J. Am. Ceram. Soc. 80 [11] 2798-804 (1997). 

64. Fox, D.S., Cuy, M.D., and Nguyen, Q.N., “Sea Salt Hot Corrosion of a Sintered -Silicon Carbide,” 

J. Am. Ceram. Soc., 81 [6] 1565-70 (1998). 
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Mach 0.3 Burner Rig Facility at the NASA Glenn Materials Research Laboratory 

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This Technical Memorandum presents the current capabilities of the state-of-the-art Mach 0.3 Burner Rig Facility. It is used for materials 

research including oxidation, corrosion, erosion and impact. Consisting of seven computer controlled jet-fueled combustors in individual test 

cells, these relatively small rigs burn just 2 to 3 gal of jet fuel per hour. The rigs are used as an efficient means of subjecting potential 

aircraft engine/airframe advanced materials to the high temperatures, high velocities and thermal cycling closely approximating actual 

operating environments. Materials of various geometries and compositions can be evaluated at temperatures from 700 to 2400 °F. Tests are 

conducted not only on bare superalloys and ceramics, but also to study the behavior and durability of protective coatings applied to those 

materials. 

15. SUBJECT TERMS

High temperature; Degradation; Oxidation; Corrosion 

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