Lancert E. Foster and John D. Saunders, Jr.

Glenn Research Center, Cleveland, Ohio

Bobby W. Sanders and Lois J. Weir

TechLand Research, Inc., North Olmsted, Ohio

Highlights From a Mach 4 Experimental

Demonstration of Inlet Mode Transition

for Turbine-Based Combined Cycle

Hypersonic Propulsion

NASA/TM—2012-217724

December 2012
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Lancert E. Foster and John D. Saunders, Jr.

Glenn Research Center, Cleveland, Ohio

Bobby W. Sanders and Lois J. Weir

TechLand Research, Inc., North Olmsted, Ohio

Highlights From a Mach 4 Experimental

Demonstration of Inlet Mode Transition

for Turbine-Based Combined Cycle

Hypersonic Propulsion

NASA/TM—2012-217724

December 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

48th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

Atlanta, Georgia, July 30 to August 1, 2012
Acknowledgments

The CCE LIMX development was led by NASA with many partners playing an integral role in the design and development. 

Project coordination and funding occurred under the aegis of the NASA Fundamental Aeronautics Program under the technical 

leadership of Mr. Scott Thomas, Dr. Kenneth Suder, and Dr. Jim Pittman. Techland Research, Inc. was the primary contractor 

for the inlet aerodynamic fl owpath design. Alliant Techsystems, Inc. (ATK) had the primary responsibility for the mechanical 

design and integration of the inlet hardware and also provided a candidate scramjet design which was carried through detailed 

design. Arctic Slope Regional Corporation (ASRC) also helped with mechanical design and were responsible for much of the 

design and fabrication of the facility hardware including the strong back. Metalex, Inc. fabricated the majority of the inlet test rig 

hardware. Williams International is providing the (high Mach) turbine engine via an interagency agreement between NASA and 

the Air Force Research Lab (AFRL) to develop high speed turbine engines. The integrated turbine engine nozzle was designed 

and fabricated by Spiritech, Inc. Boeing, Inc. contributed with CFD simulations which helped with test preparations. And fi nally 

and most importantly, the authors wish to sincerely thank the NASA GRC in-house facility engineering and technical support 

team that was led by Christine Pastor-Barsi, G. Scott Williamson, and Daniel Kovach. Their efforts allow quality data to be culled 

from the most complex inlet model testing in the 10 by 10-Foot Supersonic Wind Tunnel.

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

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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—2012-217724 1 

Highlights From a Mach 4 Experimental Demonstration of 

Inlet Mode Transition for Turbine-Based Combined 

Cycle Hypersonic Propulsion 

Lancert E. Foster and John D. Saunders, Jr. 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Bobby W. Sanders and Lois J. Weir 

TechLand Research, Inc. 

North Olmsted, Ohio 44070 

Abstract 

NASA is focused on technologies for combined cycle, air-breathing propulsion systems to enable 

reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems 

offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff 

and return mission segments along with improved safety. Among the most critical TBCC enabling 

technologies are: 1) mode transition from the low speed propulsion system to the high speed propulsion 

system, 2) high Mach turbine engine development and 3) innovative turbine based combined cycle 

integration. To address these challenges, NASA initiated an experimental mode transition task including 

analytical methods to assess the state-of-the-art of propulsion system performance and design codes. One 

effort has been the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE-LIMX) 

which is a fully integrated TBCC propulsion system with flowpath sizing consistent with previous NASA 

and DoD proposed Hypersonic experimental flight test plans. This experiment was tested in the NASA 

GRC 10 by 10-Foot Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key 

hypersonic combined-cycle engine issues including: (1) dual integrated inlet operability and performance 

issues—unstart constraints, distortion constraints, bleed requirements, and controls, (2) mode-transition 

sequence elements caused by switching between the turbine and the ramjet/scramjet flowpaths (imposed 

variable geometry requirements), and (3) turbine engine transients (and associated time scales) during 

transition. Testing of the initial inlet and dynamic characterization phases were completed and smooth 

mode transition was demonstrated. A database focused on a Mach 4 transition speed with limited offdesign elements was developed and will serve to guide future TBCC system studies and to validate higher 

level analyses.

Nomenclature 

AAS Air-breathing Access to Space 

AoA Angle of Attack 

AIP Aerodynamic Interface Plane 

CCE Combined Cycle Engine 

CCE-LIMX Combined Cycle Engine Large Scale Inlet Mode Transition Experiment 

DMSJ Dual Mode Scramjet 

FAP Fundamental Aeronautics Program 

Isp Specific Impulse 

RBCC Rocket-Based Combined Cycle 
NASA/TM—2012-217724 2 

TBCC Turbine-Based Combined Cycle 

TSTO Two-Stage-to-Orbit 

VIGV Variable Inlet Guide Vanes 

1010 SWT 10 by 10-Foot Supersonic Wind Tunnel 

Introduction 

The NASA Fundamental Aeronautics Hypersonics project is focused on technologies for combined 

cycle, air-breathing propulsions systems to enable Air-breathing Access to Space (AAS). A recent study 

(Ref. 1), which included an extensive survey of past studies, determined that a turbine-based combined 

cycle (TBCC) (Ref. 2) was the concept with the greatest promise. A TBCC propulsion system uses a 

turbine-based engine to accelerate the vehicle from takeoff to the mode transition flight condition 

(approximately Mach 3 or 4), at which point, the propulsion system performs a “mode transition” from 

the turbine to a dual mode scramjet (DMSJ) engine, which would then accelerate the vehicle on to 

hypersonic speed levels. This type of propulsion system has promising potential for Two-Stage-to-Orbit 

(TSTO) AAS. One benefit of a TBCC system would be a significant improvement in specific impulse 

(Isp) over rocket-based propulsion systems, during the subsonic takeoff and return phases of a mission, as 

shown in Figure 1. Turbine-based systems also have the potential for more operational flexibility, 

allowing for adverse weather launch, cross-range takeoff and landing, and powered landing. There is also 

a dual-use potential, as turbine engines are serviceable for both low-speed accelerator missions and longrange cruise missions. Another benefit would be the reduced system maintenance. 

Mode transition is defined as the smooth change of engine operation from a gas turbine to a dualmode scramjet; this being said, many variations of implementation are possible. These include: mission, 

vehicle, specific engine cycles and even inlet design. Therefore, this study employed a generic 

implementation of these variations, such that the resulting database would be usefully representative of 

TBCC technology. The trade space for vehicle design is complex with unknown tradeoffs between the 

desired high effective specific impulse versus the increased dry weight fraction due to an air-breathing 

propulsion system. To maximize benefits, the air-breathing transition Mach numbers were pushed to 

practical limits: Mach 4 for transition and Mach 7 for staging. Although there are other options, this 

experimental model was designed with the intent to ground a solid inlet and propulsion system database 

for future conceptual and preliminary mission studies. 

Figure 1.—Comparison of specific impulse for turbine-based and 

rocket-based propulsion systems. 
NASA/TM—2012-217724 3 

Figure 2.—Operating modes of the TBCC propulsion system. 

Based on the Next Generation Launch Technology (NGLT) studies (Ref. 3), the NASA FAP 

Hypersonics Multidisciplinary Design and Analysis Optimization (MDAO) discipline selected an overunder TBCC system (turbine engine over ram/scramjet) as the reference configuration for a TBCC TSTO 

vehicle. This over-under TBCC configuration is illustrated in Figure 2. This over-under design, in which 

the two engines share a common inlet and nozzle, saves weight, compared to a design featuring two 

separate, non-integrated inlets feeding the separate engine systems. This study investigates a twodimensional inlet concept designed (Ref. 4) with a mode transition Mach number of 4 and a staging Mach 

number of 7. From takeoff to Mach 4, both inlet ducts would be open with thrust generated primarily by 

the turbine engine. Once the vehicle accelerated to a condition that would allow ramjet combustion, the 

high-speed engine would be ignited. Although the high-speed engine would only provide minimal thrust 

at these lower Mach conditions, combustion might be necessary to fill the nozzle and reduce base drag. 

Once the vehicle accelerated to Mach 4, the designed inlet mode transition point, the high-speed flowpath 

would be open to accept appropriate airflow to for a fully operational scramjet engine. The turbine engine 

would spool down and the “low-speed” flowpath would go from fully open to almost fully closed. From 

this point, the DMSJ would provide the thrust to accelerate the vehicle to all Mach numbers beyond the 

mode transition condition. Smooth mode transition is accomplished when flow is diverted from one 

flowpath to the other, with neither experiencing unstart or buzz (Ref. 5). 

A number of enabling technologies require further development for TBCC vehicle systems for 

hypersonics applications to become a functional reality (Ref. 6). Chief among those is mode transition 

from the low-speed propulsion system to the high-speed propulsion system. A hypersonic TBCC 

propulsion system requires a high Mach turbine engine, capable of accelerating a vehicle to near Mach 4, 

where the scramjet can take over. Conversely, a scramjet must be developed with low ignition speed that 

can operate at speeds less than Mach 4. Other challenges include transonic aero-propulsion performance 

as well as the need for innovative three-dimensional flow concepts. In addition, innovative TBCC 

propulsion/airframe integration concepts will be required to minimize propulsion system dry weight, 

which stands counter to the projected Isp (wet weight) benefit. 

This study investigates a two-dimensional inlet concept designed with a mode transition Mach 

number of 4 and a staging Mach number of 7. The purpose of this study is to understand, demonstrate, 

and control the mode transition between the low-speed turbine engine and the high-speed DMSJ for a 

relevant TBCC over-under propulsion configuration. This requires a comprehensive understanding of the 

characteristics of the dual inlet design and its interactions with the engine and nozzle so that the behavior 

of the TBCC propulsion system can be modeled and predicted. An integrated performance and operability 

database must be developed that assesses factors such as: unstart constraints, bleed requirements, and 

controls characterization. This database will provide realistic distortion characteristics throughout the 

mode transition Mach number range and assess the impact of this distortion on turbine performance and 

operability. This will allow validation of computational tools used to predict TBCC propulsion system 

behavior. Lastly, the hardware from this effort will serve as a testbed for future mode transition controls 

research and integrated inlet/engine propulsion system studies. 
NASA/TM—2012-217724 4 

The planned test series includes four phases which are as follows: 

Phase I: Parametric inlet characterization—completed 

Phase II: Inlet controls system identification—completed 

Phase III: Demonstration of Closed Loop Controls—future work 

Phase IV: Fully integrated controlled mode transition with turbine engine—future work 

This paper focuses on the inlet characterization portion of testing (Phase I). The crux of supersonic inlet 

performance characterization is the effect of design and flow on inlet losses; inlet losses can be measured 

with total pressure rakes mounted at the engine inlet face. A second important inlet characteristic is 

“operability”. High Mach number inlets use a mixture of compression, both forward and aft of the cowl lip. 

Aft compression, also known as internal contraction, is necessary at speeds beyond Mach 2 to reduce losses. 

However, a major negative side effect of internal compression is the “unstart” phenomenon. With the inlet 

trimmed to match engine flow, unstart can be triggered by either upstream or downstream perturbations. An 

inlet’s ability to resist these triggering perturbations and maintain started flow is termed operability or 

stability. The “smooth” mode transition is developed by maintaining performance, avoiding unstart with 

adequate margins, and providing flow with acceptable levels of distortion. 

Downstream perturbations are caused by irregularity in the engine flow which may stem from engine 

control variations, mechanical degradation or flow instability. Flow instability is perhaps more prevalent 

in the scramjet duct where shock train and wall separations are present intentionally. The “stability index” 

is a measure of turbine flowpath stability and is often quantified as the percent change in corrected flow. 

Scramjet stability has not been extensively examined as of yet. Perturbations upstream of the inlet are 

typically caused by freestream changes due to atmospheric gusts (thermal upwellings or flow angularity). 

These disturbances can trigger unstart when they effect a drop in inflow Mach number, an increase in 

local angle of attack, or an increase in sideslip. In wind tunnel testing, stability for upstream disturbance is 

usually quantified as the change in Mach or AoA that causes unstart. 

Engine compatibility is also an important measure of inlet performance. Clearly, an inlet must match the 

inflow demands of the engine but distortion is a critical factor in assessing compatibility. The turbine engine 

community has a well-developed standard for measuring non-uniformities at subsonic and lower supersonic 

speeds (Ref. 7). Still, engine distortion tolerance is coupled to the individual turbine engine design and 

limits are usually determined empirically with the engine mounted behind distortion screens (Ref. 8). 

Turbine distortion metrics are particularly challenging as modern engines rely on dynamic data and may 

require upwards of 150,000 time slices even to characterize a nominally “steady-state” operating point. 

Military Specification (Mil-spec) inlet performance, as measured by total pressure recovery, can be 

achieved for an inlet carefully designed for single point operation. However for a TBCC system, the inlet 

must operate over a wide range of flight conditions and distribute the flow with acceptable levels of 

distortion and unsteadiness to both the turbine engine and DMSJ. Common practice is to use variable 

geometry and designed bleed configurations to enable a wide range of inlet operability and meet 

performance requirements, such as maximizing pressure recovery and minimizing flow spillage. However, 

this added complexity required to achieve high performance and operability results in additional weight, 

complex vehicle integration, and increased drag. Figure 3 illustrates that the inlet pressure recovery can vary 

by a factor of 4X (assuming Mil-spec can be achieved over the Mach flight range) depending on the inlet 

complexity. Component performance levels (such as total pressure recovery) are used in system trade 

studies for each vehicle configuration and mission such as air-breathing access to space (AAS). NASA is 

focusing on the development of validated design and analysis tools as well as parametric databases which 

are required to perform these system trade studies. Specific questions related to dual inlet technology 

addressed by CCE-LIMX inlet characterization (Phase I) testing include: 

 Is Mil-Spec performance achievable over the wide operating range that is required for TBCC space 

access? 

 What is the trade-off between bleed and inlet performance/operability? 

 Can design and analysis tools adequately predict operability limits both with and without bleed? 
NASA/TM—2012-217724 5 

 Can these tools predict low-speed/high-speed flowpath interaction effects due to varying inlet 

backpressures and low-speed/high-speed cowl positions? 

 What is the process for safe and optimum mode transition? 

In order to maintain propulsion system performance through mode transition, inlet/engine interactions 

and low-speed/high-speed engine interactions must be controlled to avoid inlet unstarts and engine stalls. 

All the while, proper airflow quality and distribution must be maintained while the turbine engine is 

spooled down as the DMSJ takes over. Specific questions with regard to inlet/engine interaction include: 

 What impact does inlet distortion have on performance and operability? 

 Can design and analysis tools predict distortion into the engine both with and without bleed? 

 Can these tools predict the impact of distortion on engine performance and operability? 

 What is the trade-off between bleed and inlet exit profile? Are vortex generators required? 

Validated design and analysis tools are necessary to enable modeling of TBCC propulsion system 

characteristics over a wide flight range and during mode transition. Specific questions must be addressed 

with regard to understanding and demonstrating mode transition such as: 

 Can design and analysis tools model components adequately to perform a controlled mode transition? 

 Are these predictions of sufficiently high fidelity to guide the testing and avoid inlet unstart and/or 

engine stall? 

 Can these design tools be effectively used to extrapolate the experimental data in order to optimize 

the propulsion system configuration? 

Figure 3.—Potential total pressure recovery for TBCC inlets. 
NASA/TM—2012-217724 6 

Experimental Apparatus 

Test Article Description 

Figure 4 shows the CCE-LIMX mounted in the NASA GRC 1010 SWT. Figure 5 shows a crosssectional view that illustrates the variable geometry features. The CCE-LIMX is an over-under dual inlet 

designed for mode transition testing, including controls and integrated inlet and engine testing (Refs. 9 

and 10). It features a high-speed flowpath designed to operate at up to Mach 7. The high-speed flowpath 

has a variable cowl and is mated with an isolator that transitions from rectangular to circular in crosssection. The low-speed flowpath is positioned directly above the high-speed flowpath and has a variable 

cowl that serves as the outer surface of the high-speed ramp; the low-speed cowl is sometimes referred to 

as the “splitter”, since the lower side serves as the ramp for the high-speed flowpath. The low-speed 

flowpath also has a variable ramp that allows the throat of the flowpath to be adjusted. Both the cowls and 

the ramp are controlled by hydraulic actuators. Each flowpath terminates with a constant area coldpipe 

and is modulated and measured by a variable exit conical plug, which serves as an engine simulator. The 

low-speed flowpath has 13 bleed regions, with plenums ducted to 15 remotely controlled coldpipe/massflow-plug assemblies. Figure 6 shows the arrangement of these bleed regions throughout the low-speed 

flowpath. The low-speed flowpath can also be fitted with vortex generators on its ramp, cowl, and 

sidewalls. Some configurations during Phase I testing did feature vortex generators. 

Figure 4.—CCE-LIMX installed in the NASA 1010 SWT. 

Figure 5.—CCE LIMX installation schematic with highlights of variable geometry features. 
NASA/TM—2012-217724 7 

Figure 6.—Bleed configuration for the low-speed flowpath. 

Figure 7.—High-speed AIP rake array. Figure 8.—Low-speed AIP rake array. 

Instrumentation 

Both the high-speed and low-speed flowpaths are equipped with longitudinal rows of static pressure 

taps on both their cowl and ramp flow surfaces. Also, each flowpath is fitted with an aerodynamic 

interface plane (AIP) rake array as shown in Figures 7 and 8. Each of the area-weighted tubes in the lowspeed array includes a dynamic pressure transducer, in addition to the steady-state tube; this allows 

determination of industry standard distortion profiles of the flow being fed to the turbine engine. This data 

is critical for assessing the inlet’s compatibility with the CCE Turbine Engine, discussed later in this 

paper. The CCE Turbine Engine will be mounted to the low-speed duct during Phase IV testing. 
NASA/TM—2012-217724 8 

1010 SWT 

The NASA 1010 SWT can operate over a Mach number range of 2 to 3.5 (Ref. 11). Because the 

tunnel does not operate at Mach 4, this condition must be simulated. This is accomplished by a large 

trapezoidal plate mounted on the front of the model, clearly visible in Figure 8; this plate is called the 

“pre-compression ramp”. In an actual Mach 4 flow, there would be Mach 3.533 flow over the precompression ramp due to the oblique shock caused by the 6.5° ramp angle, as shown in Figure 9. The 

CCE-LIMX is mounted to the facility strut and the angle of attack (AoA) can be varied from –15 to +5. 

With the tunnel operating at Mach 3.5 and the CCE-LIMX at an approximate AoA of –6.5, as shown in 

Figure 10, a very small expansion angle on the pre-compression ramp gives us Mach 3.533 flow. In 

regard to the inlet, the only aerodynamic difference between actual Mach 4 freestream flow and simulated 

Mach 4 is the total pressure recovery across the oblique shock. The variation in total pressure loss is 

corrected for in the recorded data. 

Figure 9.—Inlet and attitude at Mach 4 flight condition. 

Figure 10.—CCE-LIMX at simulated Mach 4 tunnel condition. 
NASA/TM—2012-217724 9 

Mach 3 Capable Turbine Development 

Phase IV of CCE LIMX testing includes removing the coldpipe and mass-flow plug assembly from 

the low-speed flowpath and replacing it with an actual Mach 3 capable turbine engine. Therefore, 

Williams International has developed and built an engine for this purpose. The CCE Turbine Engine is a 

modified version of Williams’ WJ38 Turbine Engine with some adaptations for Mach 3 operation. 

Figure 11 shows a photo of the engine and the integrated nozzle assembly. The production WJ38 is not 

suited to accommodate the increased distortion levels and temperature; the CCE engine must tolerate a 

600 °F max inlet temperature. Thus, a distortion- and temperature-tolerant fan, incorporating forwardswept blades, was added along with variable inlet guide vanes (VIGV). To accommodate the increased 

temperature, material changes were made in some compressor and lubrication system components. Flame 

spray abradable was added to some compressor surfaces and Thermal Barrier Coating (TBC) was added 

to the combustor. Lastly, an afterburner and Single Expansion Ramp Nozzle (SERN) were added to the 

engine, allowing for improved supersonic operation. Because the CCE engine was designed for ground 

testing in the NASA GRC 1010 SWT with no particular performance specifications in mind, it was not 

weight-optimized as a comparable flight engine would be. The engine is currently available and has 

undergone sea level static (SLS) checkout tests. A final series of checkout/acceptance testing was 

conducted with an afterburner and integrated nozzle installed in May 2011. Following post-test 

inspections, the engine was delivered to NASA GRC on August 2, 2011. 

Figure 11.—Williams International High Mach WJ38 turbine engine and 

integrated nozzle assembly. 
NASA/TM—2012-217724 10 

Results and Discussion 

The inlet performance and operability tests (Phase I) encompassed 30 run nights and 145 hr on 

condition. Comprehensive data sets were obtained at Mach 4 and 3, while limited data sets were obtained 

at Mach 3.5, 2.5, and 2. Note that different configurations in terms of inlet features such as ramp/cowl 

settings, low-speed cowl leading edge, boundary layer bleed patterns, and vortex generator design were 

studied. The system identification/dynamics tests (Phase II) encompassed 17 run nights and 90 hr on 

condition. A total of 651 dynamic experiments were conducted; 495 at Mach 4 and 156 at Mach 3. 

Due to security restrictions on the data, quantitative details cannot be presented in this open paper. 

Full detailed test results will be presented later; however, Mach 4 results are currently available (Ref. 5). 

Figure 12 shows a qualitative representation of the mode transition data acquired during testing. This 

mode transition map for the low-speed flowpath is composed of curves of total pressure recovery versus 

duct exit flow ratio. The curves are termed “inlet performance canes”. Each successive inlet cane (as the 

cowl lip was moved toward the closed position) moved to a lower mass-flow ratio with a reduced 

recovery at the data point that was recorded just prior to inlet unstart (minimum stable). The ultimate 

objective is to achieve full system mode transition without incurring inlet unstart, inlet buzz or engine 

stall. Initial test results indicate that successful and smooth inlet mode transition was demonstrated with a 

split-flow TBCC inlet system. 

The results schematically represented in Figure 12 for the low-speed flowpath, along with the highspeed flowpath results, are being compared to pre-test computational solutions and to the small scale inlet 

mode transition experiment (IMX) results previously obtained in the NASA GRC 11 SWT (Ref. 12). 

The unstart and distortion limits are depicted as shaded regions in Figure 12. 

Figure 12.—Schematic plot of mode transition sequence representative of low-speed inlet test data. 
NASA/TM—2012-217724 11 

Internal contraction ratio limits were found for both flowpaths with limited variation of bleed, Mach 

number, and ramp geometry. With this information, two mode transition tracks were determined for the 

high-speed flowpath: an aggressive one, with better thrust performance, and a conservative one, with 

more stability. Delta Mach and AoA limits were also found. Parametric testing provided a design 

database that includes: bleed versus recovery trades, bleed versus engine inlet distortion trades, and 

unstart and distortion limits. 

Dynamic data were obtained with the low-speed AIP rake at conditions appropriate for operation of 

the CCE Turbine Engine. This was done in a manner consistent with turbine industry standards for 

assessing engine inflow quality, which allows the calculation of distortion indices. Distortion profiles will 

be incorporated into analytical tools to predict engine performance and operability constraints, in 

preparation for fully integrated testing. 

For a scramjet engine, a distortion profile can be used to calculate the mass flow distribution. These 

profiles might enable scramjet designs that include differential fuel distribution to approximate constant 

fuel/air ratio combustion. To date, aspects of even basic scramjet control continue to be researched 

(Ref. 13). Nonetheless, dynamic data from the CCE-LIMX high-speed flowpath will prove useful as the 

isolator duct is coupled to a relevant TBCC dual inlet. 

Further development of the TBCC propulsion concept is needed to address the transient nature of 

mode transition. Certain aspects of the mode transition remain unknown, as they are dependent on the 

selection of a specific mission and vehicle. Ideally, acceleration would be maintained or even enhanced 

during mode transition and in trimmed flight, AoA would probably vary as well. If net thrust production 

were to lag during the transition, vehicle deceleration could result. To simplify these unknowns in the 

current testing, Mach and AoA were generally kept constant during mode transition. Development of this 

database will help to remove some of these unknowns. The primary focus was on a transition Mach 

number of 4 to match the inlet design point. However, a secondary constant Mach transition was also 

investigated at Mach 3. This alternate Mach condition matched the upper performance limit of the CCE 

turbine engine built for this project. 

Limited data were also gathered at M3.5, 2.5, and 2.0. These intermediate points provide skeletal 

TBCC performance information across the operating range: from starting Mach numbers (around Mach 2) 

through the main turbine thrust production (Mach 2.5) and on to the mid-point of mode transition speed 

(Mach 3.5). 

Eventually, variable Mach mode transition must be investigated as well. A representative accelerating 

mode transition was conducted, although a more comprehensive study of Mach and AoA variations will 

be called for in the future. Data from this study reveals specific details of Mach and AoA variation, which 

enables their incorporation into the mode transition modelling of further mission studies. 

Dynamic data were obtained to capture the details of transients present during mode transition. These 

data were used validate design and analysis approaches (Refs. 14 and 15) and are available to prepare 

control laws for close-looped mode transition. 

With the resulting database, propulsion system design approaches can be validated. The inlet design 

approach was outlined in the work of Sanders and Weir (Ref. 4). In the completed test to date, the inlet 

design intent can be compared to the test results. Where the inlet deviates from expected behavior, the 

large database will be scrutinized and modifications or refinements to the design methods will be reported 

and incorporated into the next generation of designs. 

Alternate model configurations were tested to widen the applicability of the database and provide 

some means to improve performance. These variations included: alternate cowl lips, vortex generators, 

and bleed variations for the turbine flowpath. Depending on baseline configuration results, testing of 

alternate configurations could be extensive or very limited. Six run nights, or about 5 percent of the 

testing, were devoted to this investigation. The effect of bleed complexity was also investigated; by 

closing various bleed exits, the effect of reduced bleed complexity on low-speed flowpath performance 

could be quantified. 

 
NASA/TM—2012-217724 12 

Using selected configurations, the inlet was characterized for performance and operability. A large 

amount of test time was devoted to testing at Mach 4 conditions. A lesser but still significant amount of 

the test matrix was used to look at selected geometry at the turbine engine compatible speed of Mach 3. 

Where performance expectations were met, mode transition was investigated and documented at each of 

these two “design” Mach numbers. 

The validation of advanced CFD-based analytical tools is of critical importance. Though RANS 

solvers have become production tools for many flows, prediction of supersonic inlet flow remains 

challenging. In particular, the interplay of the normal and oblique shock with the bleed zones in the 

turbine flowpath, the prediction of an isolator shock-train in the scramjet flowpath and the 

characterization of the distortion field at the inlet duct exits are goals that this inlet database will help 

address. Beyond RANS solvers, future algorithms are under development that may be confirmed with the 

inlet database. 

A significant amount of pre-test CFD analysis was completed in preparation for the objective of CFD 

tool validation (Refs. 16 to 18). Most of this effort was focused on the transition design point of Mach 4. 

On-going CFD will continue this effort and investigate the depth of the database at off-design Mach 

numbers and alternate configurations. 

Conclusion 

One of the missions of the NASA Aeronautics Research Directorate (ARMD) FAP Hypersonics 

project is to develop propulsion technology for AAS. The focus of this effort is in developing an inlet 

database for the TBCC TSTO propulsion system. There are many technical challenges involved with the 

development of a TBCC propulsion system and these were discussed in this paper. The CCE-LIMX 

article tested to date included mass-flow plug and cold pipe assemblies in both the low-speed and highspeed flowpaths. For Phase IV, a high Mach capable turbine engine and single expansion ramp nozzle 

(SERN) would be incorporated into the low-speed flowpath. This engine and nozzle have also been 

developed as a part of the TBCC discipline. 

Overall this test activity has provided a rather comprehensive database on this TBCC design. 

Specific results were: 

1. Phases I and II of TBCC testing (Parametric inlet characterization tests and Inlet system identification 

tests) were successfully completed. 

2. A database was compiled for a dual inlet for TBCC mode transition which includes performance and 

operability information. Total pressure recovery, unstart limits, and distortion levels are documented 

for both flowpaths. Limited configuration effects from bleed, cowl lip, vortex generator, and 

contraction ratio were investigated. 

3. Smooth mode transition was demonstrated in the inlet at a variety of Mach numbers. 

4. Data was taken at conditions corresponding to the requirements of the CCE Turbine Engine. 

5. The data collected enabled Phase III controls efforts: 1) Identify the frequency range where a closedloop control system is needed to enhance the inlet stability margin; and 2) Determine control design 

linear models (CDMs) for closed loop control of the CCE-LIMX to simulate a hypersonic vehicle 

propulsion system mode transition. 

Future tests are now possible which will serve to demonstrate controlled mode transition, a 

requirement for an advanced air-breathing propulsion system that enables hypersonic flight. Beyond the 

first two phases of this current study, the CCE LIMX testbed will be available to address future national 

needs in combined cycle propulsion.
NASA/TM—2012-217724 13 

References

1. Bartolotta, P., Wilhite, A.W., Schaffer, M., Voland, R.T., and Huebner, L., “Horizontal Launch: A 

Versatile Concept for Assured Space Access,” NASA/SP-2011-2151994; DFRC-E-DAA-TN4418, 

December 2011. 

2. Science Applications International Corporation, “Integrated Concept Model 2 (ICM-2): A twostage-to-orbit turbine based combined cycle concept,” ASAP Team Final Report, 21 February 2003. 

3. Bilardo, V.J., Curran, F.M., Hunt, J.L., Lovell, N.T., Wilhite, A.W., and McKinney, L.E., “The 

Benefits of Hypersonic Airbreathing Launch Systems for Access to Space,” AIAA 2003-5265, 39th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Huntsville, AL, 20-23 July 2003. 

4. Sanders, B.W., and Weir, L.J.: “Aerodynamic Design of a Dual-Flow Mach 7 Hypersonic Inlet 

System for a Turbine-Based Combined-Cycle Hypersonic Propulsion System,” NASA-CR-2008-

215214, June 2008. 

5. Sanders, B.W., Weir, L.J., Saunders, J.D., and Foster, L.E., “Mach 4 Experimental Demonstration of 

Inlet Mode Transition Technology for a Turbine-Based Combined Cycle Hypersonic Propulsion 

System,” NASA/TM—2012-217647; September 2012. 

6. NIA Report, Boeing Contractor Report, Contract No. 2900-HY04, February 18, 2005. 

7. SAI S-16 Committee, “Aerospace Recommended Practice: Gas Turbine Engine Inlet Flow 

Distortion Guidelines,” Society of Automotive Engineers America Recommended Practice, 

ARP 1420B, March 2002. 

8. Rock, K.E., Nugent, M.R., Orme, J.S., and Calleja, J.F., “Propulsion Subsystem Control Law 

Development for the Hyper-X Mach 7 Flight Test,” Limited by Export Administration Regulations 

(EAR), JANNAF 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee 

Joint Meeting/Volume 2; 205-226; CPIA-Publ-Vol-2, October 1999. 

9. Saunders, J.D., Slater, J.W., Dippold, V.F., Weir, L.J., and Sanders, B.W., “Inlet Research Status for 

a Large-Scale Turbine-Based Combined Cycle Engine Testbed,” JANNAF, La Jolla, CA, December 

7-11, 2009. 

10. Saunders, J.D., Stueber, T.J., Thomas, S.R., Suder, K.L., Weir, L.J., and Sanders, B.J., “Testing of 

the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition 

Experiment (CCE LIMX),” NASA/TM—2012-217217, 2012. 

11. Soeder, R.H., Roeder, J.W., Linne, A.A., and Panek, J.W., “User Manual for NASA Glenn 10-by 

10-Foot Supersonic Wind Tunnel,” NASA/TM-2004-212697, May 2004. 

12. Saunders, J.D., Slater, J.W., Dippold, V.F., Lee, J., Sanders, B.W., and Weir, L.J., “Inlet Mode 

Transition Screening Test for a Turbine-Based Combined-Cycle Propulsion System,” JANNAF, 

Boston, MA, May 13, 2008. 

13. Prahst, P.S., and Suder, K.L., “Results of the High Mach Fan Rig Experiment,” NASA 2012 

Fundamental Aeronautics Program Annual Meeting, Cleveland, OH, March 2012. 

14. Csank, J., and Stueber, T.J., “A Turbine Based Combined Cycle Engine Inlet Model and Mode 

Transition Based on HiTECC Tool,” JPC, Atlanta, GA, July 30-August 2, 2012. 

15. Le, D.K., Vrnak, D.R., Slater, J.W., and Hessel, O.E., “A Simulation Scheme for Turbine-based 

Combined Cycles Inlet Mode-Transition Control,” JPC, Atlanta, GA, July 30-August 2, 2012. 

16. Slater, J.W., and Saunders, J.D., “CFD Simulation of Hypersonic TBCC Inlet Mode Transition,” 

AIAA-2009-7349, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and 

Technologies Conference, October 2009. 

17. Dippold, V.F., “Computational Analyses of the LIMX TBCC Inlet High-Speed Flowpath,” 

58th JANNAF Propulsion Meeting, April 2011. 

18. McDaniel, J., “An Overview of the National Center for Hypersonic Combined-Cycle Propulsion,” 

NASA 2012 Fundamental Aeronautics Program Annual Meeting, Cleveland, OH, March 2012. 
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Highlights From a Mach 4 Experimental Demonstration of Inlet Mode Transition for TurbineBased Combined Cycle Hypersonic Propulsion 

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14. ABSTRACT

NASA is focused on technologies for combined cycle, air-breathing propulsion systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle 

(TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments along with 

improved safety. Among the most critical TBCC enabling technologies are: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) 

high Mach turbine engine development and 3) innovative turbine based combined cycle integration. To address these challenges, NASA initiated an experimental mode 

transition task including analytical methods to assess the state-of-the-art of propulsion system performance and design codes. One effort has been the Combined-Cycle Engine 

Large Scale Inlet Mode Transition Experiment (CCE-LIMX) which is a fully integrated TBCC propulsion system with flowpath sizing consistent with previous NASA and 

DoD proposed Hypersonic experimental flight test plans. This experiment was tested in the NASA GRC 10 by 10-Foot Supersonic Wind Tunnel (SWT) Facility. The goal of 

this activity is to address key hypersonic combined-cycle engine issues including: (1) dual integrated inlet operability and performance issues-unstart constraints, distortion 

constraints, bleed requirements, and controls, (2) mode-transition sequence elements caused by switching between the turbine and the ramjet/scramjet flowpaths (imposed 

variable geometry requirements), and (3) turbine engine transients (and associated time scales) during transition. Testing of the initial inlet and dynamic characterization phases 

were completed and smooth mode transition was demonstrated. A database focused on a Mach 4 transition speed with limited off-design elements was developed and will 

serve to guide future TBCC system studies and to validate higher level analyses. 

15. SUBJECT TERMS

Hypersonic inlets; Supersonic inlets; Inlet design; Inlet test 

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