Air force 17. 1 Small Business Innovation Research (sbir) Phase I proposal Submission Instructions

AIR FORCE SBIR 17.1 Topic Descriptions

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AIR FORCE SBIR 17.1 Topic Descriptions


TITLE: Variable Pressure/Flow Control Firefighting Nozzle

TECHNOLOGY AREA(S): Materials/Processes

OBJECTIVE: Develop a hose-end nozzle system capable of allowing the operator to hand select discharges of Ultrahigh-Pressure (UHP; 1100 to 1500 PSI @ 15 to 20 GPM) and low-pressure (LP; 100 PSI @ 15 to 20 GPM) water streams at variable flow rates through a single hose line and application nozzle.

DESCRIPTION: UHP and LP nozzle technologies already exist and are commercially available, although no one nozzle is capable of providing both UHP and LP water streams. A variable-pressure/flow-control nozzle will allow the first responder the option to utilize, as applicable, either technology to respond to fire incidents. The nozzle shall be designed to be compatible with the existing 4-stage centrifugal pump and 200 ft of 1-in high-pressure hose line mounted on the Air Force P-34 Rapid Intervention Vehicle (RIV).

The variable flow nozzle should be easily handled with similar design to existing firefighting nozzles i.e. with pistol grips and bails as needed to operate the nozzle. The ability to control the nozzle discharge in either UHP or LP shall be by a simple means of a nozzle bail or trigger and or combination of each for the discharge of water thorough the nozzle at variable flow rates. The application nozzle shall also be capable of straight and or fog patterns as selected by the operator.

Candidate technologies should balance commercial considerations with DoD requirements.

PHASE I: R&D goals of phase I are to identify the possibility of using a single pump, hose and nozzle system to provide both UHP and low-pressure waster flows between 15 and 20 GPM at variable pressures from 100 to 1500 PSI Phase I deliverable is design specifications. Offerors are encouraged to work with AFCEC's Fire Research Group during the design stage.

PHASE II: Produce a full-scale prototype nozzle. Provide detailed specifications of max and min pressure, GPM and capabilities of the nozzle. Conduct a full-scale demonstration of the nozzle at AFCEC's Fire Research Facilities

PHASE III DUAL USE APPLICATIONS: Deliver a marketable product with detailed installation/operator guidance manual.


1. National Fire Protection Association (NFPA) Standards




KEYWORDS: Fire, Firefighting, Mist, Nozzle, Pressure, Stream, Water


TITLE: Wide-angle Retroreflectometer

TECHNOLOGY AREA(S): Materials/Processes

OBJECTIVE: A hand‐held system that fits light‐tight over reflective beads marking runway surfaces and measures light intensity reflected at angles from 0–45° from a beam incident at 0–45 degrees and displaced 0–30 degrees from the reflected beam.

DESCRIPTION: Retroreflective glass beads are widely employed as a technology for improving the visibility of highway road markings to drivers in unlighted areas at night. In the context of that application, two limiting geometries are regularly important: as the vehicle approaches a road sign or other horizontal marker, light arrives perpendicular to the reflective surface and returns along a nearly antiparallel optical path to the observer; at centerline striping and edge markings, light from the vehicle arrives nearly parallel to the road surface and reflects back to the driver at an angle that is only slightly larger. Technology for high‐precision measurement of perpendicular reflectance that operates in a range of ~0.5 degree from the normal is available and widely used to evaluate the condition (inferred from 90‐degree reflectance of light incident at ~90 degrees) of both categories of highway markings.

Until the adoption of centerline runway illumination, reflective glass beads were also the tool of choice to provide observability to allow pilots to land aircraft at night and taxiing. Their utility in for this purpose was effectively eliminated when runway lighting came into use, but use of the beads has been continued as a backup in the event of a lighting failure, and because they afford some measure of visibility on the ground, reflecting landing and taxiing lights. As for the second highway case, measurement with current technology is marginally representative at best because illumination and reflection for aircraft on runway and taxiway surfaces occur at geometries vastly different from that measured by conventional retroreflectometers, for which measured reflectances typically vary widely in field measurements.

The instrument sought will perform measurements of reflectance at geometries representative of aircraft operations, from which criteria indicating need for and that restriping treatments must pass for acceptance will be developed to replace the FAA recommendations based on perpendicular reflectance. The instrument sought will exclude external light and allow independent adjustment to the nearest degree of the incident angle, reflected angle, and radial displacement of the latter from the vertical plane of the former. It will operate on rechargeable batteries, preferably solar, for at least 4 hours, and onboard software will calculate means and standard deviations for sets of nine adjacent measurements defining a square array at each sampling point.

PHASE I: Develop a breadboard prototype. Demonstrate proof of concept in a darkened laboratory: reproducibly (+/- 20% for six sampling points in each angular dimension) map reflectance of a 1-inch square each of a commercial reflective tape and a runway coating to which highway beads have been applied to the supplier's specifications. Deliverables include a constructible design for a field-testable prototype.

PHASE II: Produce a full-scale prototype and mar, as in phase I, five randomly selected sites each on new and old coatings on asphalt and concrete highway surfaces and new and old coatings on asphalt and concrete runway surfaces. For each, agreement between averages of the prototype's and conventional instrument's measurements in perpendicular orientation must agree within 20%. Deliverables include a 50% design for a manufacturable prototype

PHASE III DUAL USE APPLICATIONS: Deliver a marketable product that minimally trained personnel deployed in an austere environment can operate. Unit will survive being dropped 4 feet onto a concrete surface and rolled down a 20-foot, 15% incline, and will function in rain or salt spray and at temperatures from -30 to 140 degrees F.


1. Airfield Marking Handbook, Innovative Pavement Research Foundation Project 05-01 Report, available from

2. "Retroreflection from spherical glass beads in highway pavement markings. 2. Diffuse reflection (a first approximation from calculation)." Vedam, K., and Stout, M.D., Applied Optics, (1978) 17(12):1859-1869.

3, Pavement Markings and Safety, Smadi, O.; Hawkins, N.; Nlenanya, I.; and Aldemir-Bektas, B., Iowa Highway Research Board Project TR 580, Nov 2010; available from

KEYWORDS: Beads, markings, observability, reflectance, reflection, runway, taxiway, visibility


TITLE: Alternative Method of Surface Activation of High Strength Steels for Electroplating

TECHNOLOGY AREA(S): Materials/Processes

OBJECTIVE: Develop a surface activation technology for electroplating landing gear parts that shall reduce the amount of material removed from the parts, improve plating adhesion, reduce waste, and lower risks to both personnel & landing gear parts.

DESCRIPTION: In the preparation of landing gear parts to be electroplated, the parts are usually prepared using a blasting technique. The blasting technique utilizes either aluminum oxide or garnet and this process is an aggressive technique which activates the surface. The activation is obtained by abrading off the surface layer by one ten-thousandth to one thousandth of an inch per second.

As material is abraded from the base material, some particulates of the base metal become part of the blast media. This cross contamination of the blast material creates possible problems such as corrosion of landing parts by dissimilar metals. In order to minimize the problems associated with cross contamination, independent blasting systems are used for each metal type. Additionally, the blast media is a non-renewable material and after cycling through the filtration system, the particles that are too small are discarded. These attributes increase the operational cost for preparing landing gear parts for electroplating.

The Air Force is interested in other methods for preparing the surface for electroplating. Because of the high dynamic stresses that landing gear undergoes, compatibility of the process must be such that the surface activation process does not over-heat the base metal (not to exceed 425 deg. F.), cause embrittlement of the base metal, or other danger to the landing gear. In addition, the removal of the base material while activating the surface should be minimal. It is preferred that only the first few atoms of the surface will be removed instead of up to a ten-thousandth of an inch.

The new technology needs to be able to function on multiple landing gear parts that consists of different high strength steel (heat treat = 180ksi) alloys in a variety of sizes and shapes. Conduct feasibility verification and validation testing of the alternative technology to include, but not limited to, adhesion of electroplated materials to base metal, temperature testing, fatigue testing, corrosion testing and hydrogen embrittlement testing. Along with the safety of the landing gear parts; personnel safety must be taken into consideration.

PHASE I: Develop a solution that meets above requirements and conduct preliminary business case analysis (BCA) to determine implementation costs, including a return-on-investment (ROI) calculation that compares anticipated savings to expected costs. Mishap avoidance shall not be included in cost calculations. Proof-of-concept prototype(s) shall be developed to demonstrate conformance to the requirements.

PHASE II: Optimize the developed solution. Perform testing on large diameter surface areas as well as small surface areas, both outer diameters and inner diameters. Prepare specification so the technology can be demonstrated verified and validated. Continue to refine the BCA to determine costs for implementation; BCA should include an ROI calculation that compares the anticipated savings to the expected costs.

PHASE III DUAL USE APPLICATIONS: Continue to verify the technology’s activation of high strength steels for electroplating. Perform cost benefit analysis. Implement the new technology. Expand workload across all electroplating platforms which may include other Air Force bases.


1. “Surface treatment of metals using an atmospheric pressure plasma jet and their surface characteristics,” M.C. Kima, b, S.H. Yanga, J.-H. Boob, J.G. Hanb; Nano Surface Technology Team, Korea Institute of Industrial Technology, Cheonan 330-825, South Korea.

2. “Plasma jet treatment of five polymers at atmospheric pressure: surface modifications and the relevance for adhesion”; International Journal of Adhesion and Adhesives, Volume 24, Issue 2, April 2004, Pages 171-177 Michael Noeske, Jost Degenhardt, Silke Strudthoff, Uwe Lommatzsch.

3. MIL-STD-1504.

4. A-A-1722.

KEYWORDS: electroplating, surface activation, landing gear


TITLE: Electroplating 3D Printed Materials

TECHNOLOGY AREA(S): Materials/Processes

OBJECTIVE: Research and develop an appropriate process to include specifications, equipment requirements, optimal applications and cost/benefit for metal plating 3D Fused Deposition Modeling (FDM) and Stereo Lithography (SLA) printed parts.

DESCRIPTION: Metal plating FDM parts with chromium, nickel or copper can greatly enhance strength, durability, surface hardness, and heat resistance. Industry has claimed that the metal plating thickness of a FDM or SLA part can be between 0.001 - 0.020 inches and can improve strength by a factor of 20. Open source web searches and inquiries to FDM and SLA equipment vendors indicate that the 3D part plating process is poorly defined and there is very little information available to specify or replicate the process. Additionally, the processes for plating complex 3D printed FDM and SLA parts is not well understood from a design perspective and internal areas may not be plated consistently or predictably.

The Air Force is interested in employing a metal plating technology of FDM and SLA printed parts in Depot maintenance processes. Examples of FDM and SLA material that may be metal plated are ABS, Nylon, Polycarbonate, Polypropylene, or ULTEM 9085. The new technology shall demonstrate the ability to metal plate on complex FDM and SLA parts. The prospective parts may include: casting molds, large metal press molds, parts with conductive surface requirements, tooling, fixtures, and increased size parts (larger than 3’x2’x3’) that are then plated and bonded together.

This effort shall provide opportunities for the Air Force to employ metal plating of FDM and SLA printed parts to enhance attributes of strength, durability, surface hardness, and heat resistance. Additionally, the efforts shall quantify material properties, structural integrity, process predictability and reproducibility. Furthermore, this effort shall develop process specifications and design guidelines. The efforts shall identify limitations, level of effort, space & utility requirements, equipment, and time requirements. Provide a business case analysis quantifying the expected cost for plating along with several alternative materials including part complexity and thickness of metal. Verify plating can be applied uniformly at any desired thickness range between 0.001 – 0.020 inches. Validate and verify plated prototype parts and evaluate the enhancements at varying plating thicknesses. The printed parts shall have bond lines where printed parts have been joined together. Validate finite element analysis for structure optimization.

PHASE I: Develop a solution that meets above requirements and conduct preliminary business case analysis (BCA) to determine implementation costs, including a return-on-investment (ROI) calculation that compares anticipated savings to expected costs. Proof-of-concept prototype(s) shall be developed to demonstrate conformance to the requirements.

PHASE II: Proof-of-concept prototype(s) shall be refined to installation-ready article and shall undergo testing to verify and validate all requirements. This process may require multiple iterations before a final design is selected. Refine BCA/ROI based on the final design.

PHASE III DUAL USE APPLICATIONS: If cost effective, implement developed technology.


1. Stratasys Finishing Processes Bond, seal and beautify 3D printed parts


KEYWORDS: 3D printing, electroplating, strength


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TITLE: Development of Additive Manufacturing for Landing Gear Components

TECHNOLOGY AREA(S): Materials/Processes

OBJECTIVE: Research and develop Additive Manufacturing (AM) technologies for landing gear components to enable testing and production of landing gear components using AM techniques.

DESCRIPTION: Due to the extreme operational loading and harsh environments, landing gear designs have led to manufacturing and fabrication processes that are lengthy (i.e. forgings) and expensive. Landing gear primary structural components include pistons, cylinders, truck beams, drag braces, side braces. Secondary components include actuators, links, gear boxes, brackets, bushings and bearings. Many components require forgings to ensure proper grain flow in order to prevent premature fatigue failures. The need for forgings adds significant costs, delays, and complexity to the manufacturing process. Additionally, current machining methods limit component designs to produce the most efficient geometry for strength and weight reduction. Many components are currently made from lower cost materials (low alloy steel and aluminum). Other alloys, such as titanium, could greatly improve the part performance, strength to weight ratio, weight, and corrosion resistance. In addition to the cost of materials for these high performance alloys, expense of manufacturing often offsets the feasibility of their use. In conjunction with material and fabrication challenges, USAF forecasting, planning, and procurement methods require purchasing of low quantity buys that eliminate economies of scale. This purchasing practice can inherently drive higher part costs, which is the case with costly forging and machining of landing gear components. With current landing gear manufacturing and fabrication processes, the cost to sustain the aging USAF fleet will continue to grow.

A game changing fabrication process is needed to replace current forging and machining processes to produce landing gear components of equal or improved strength, weight, and corrosion resistance. Metal additive manufacturing (AM, or 3D printing), technology has grown exponentially in the last several years and it appears now feasible to manufacture landing gear components using this technology. To achieve this, AM materials must be identified and tested to provide the required data that will satisfy USAF airworthiness requirements that will in turn, allow these materials to be used on current and future landing gear designs. Additionally, the required design processes, metal powder combinations, manufacturing techniques, post processing techniques, and qualification testing requirement must be developed for landing gear components.

Benefits to the use of AM technology for landing gear components include: cost savings, significant improvement in on-time deliveries of landing gear components, significant reduction in outside processing time, improved design alternatives, possibility for weight reduction, optimizing stress concentration areas, corrosion mitigation and improved production lead-time.

The purpose of this task is identification and development of AM material and process development for landing gear components. The task will investigate the following:

1. If a standard AM process is achievable for landing gear component materials and if the selected standard process will produce consistent repeatable mechanical properties as defined by industry standards

2. Design techniques/changes requisite for AM of landing gear components

3. Powder compositions for candidate materials including aluminum, titanium, steel and stainless steel alloys

4. AM techniques to ensure completed printed parts have the required material properties equivalent to existing material

5. Post processing requirements (Heat treatment, shot peening, High Velocity Oxygen Fuel (HVOF), plating (anodize, Low Hydrogen Embrittlement Zinc Nickel (LHE Zn-Ni), etc.)

6. Non-destructive Inspection/Test (NDI/NDT) and general testing methods to ascertain these techniques are robust to ensure part integrity

7. Qualification requirements

8. Equipment and facility requirements are identified and documented to ensure component integrity.

PHASE I: Research, develop & perform coupon level testing of candidate landing gear material. Phase I final report will provide results of how AM developed coupons compared with the traditional manufactured coupons & ability of the material to be viable for airworthiness certification of a candidate landing gear part.

PHASE II: Based on the Phase I research, work with the 417 SCMS government team to identify a candidate landing gear part for selection & AM fabrication, NDI/NDT evaluation, and Qualification Testing. The contractor shall complete a full-scale component fabrication & first article inspection. Following inspection the part shall be submitted to the required NDI/NDT processes for ver/val. After NDI/NDT the part shall be submitted to full qualification testing to ensure airworthiness compliance.

PHASE III DUAL USE APPLICATIONS: Based on the Phase I research, work with the 417 SCMS government team to identify a candidate landing gear part for selection & AM fabrication, NDI/NDT evaluation, and Qualification Testing. The contractor shall complete a full-scale component fabrication & first article inspection. Following inspection the part shall be submitted to the required NDI/NDT processes for ver/val. After NDI/NDT the part shall be submitted to full qualification testing to ensure airworthiness compliance.


1. Additive manufacturing technologies rapid prototyping to direct digital manufacturing, By: Gibson, I., and D. W. Rosen.



KEYWORDS: Additive manufacturing, Landing Gear, Manufacturing, 3D printing, sustainment, high strength steel, materials science, high strength aluminum, titanium, aerospace structures


TITLE: Real-time, 3-D Model Deformation Measurement Capability for Perforated Wall Wind Tunnels


OBJECTIVE: Develop an accurate, real-time, 3-D capability for the quantitative measurements of position, attitude, and surface deformation of test models in large scale wind tunnels.

DESCRIPTION: Dynamic loading and deformation of wind tunnel models during testing contribute to the overall measurement uncertainty of aerodynamic parameters. Unlike flow-field uncertainties, which are quantified by wind tunnel calibration studies, the uncertainty associated with model attitudes and deformities has to be determined during the test. In the absence of a separate and calibrated diagnostic process, the effects produced by these irregularities of the model cannot be distinguished from the random error of the measured aerodynamic data. An innovative technology solution to accurately measure test article position, attitude, surface deformation and the quantification of fluctuations of these parameters is needed to meet test customer requirements.

Application of a measurement system of this nature is difficult due to test facility structure. To achieve air flow quality requirements, the AEDC 16T wind tunnel has porous walls in the test section that houses the test article and the overall porosity must not be significantly diminished. The optical and lighting components of this technology must be implemented without significant modifications to the porous wall design of AEDC’s 16T wind tunnel. The walls are ¾ inch thick and have many ¾ inch diameter holes drilled through them at a 60 degree angle with respect to the wall normal direction. The holes are spaced approximately 2.7 inches apart in a hexagonal pattern, described in reference 6. This pattern allows up to a 1.5 inch diameter hole to be drilled straight through the wall between any two canted perforation holes, and that straight hole can be used for optical access if necessary. This area of the tunnel sees pressures ranging from 47 to 4000 psf and temperatures from 60 to 140 degF. Any installed equipment would need to be able to withstand these conditions.

Innovative techniques to combine multiple images to produce highly accurate, real-time, 3-D test model position, attitude and surface deformation measurements during static and dynamic test events are needed. The ultimate goal is to develop a prototype system that can be transitioned to a wind tunnel test facility at AEDC, such as the 16T transonic wind tunnel. The prototype system should include optical components, data acquisition system, and software required to satisfy the topic.

PHASE I: Demonstrate feasibility of the technique to measure position, attitude and surface deformations of a representative 3-D model in a laboratory environment.

PHASE II: Develop and demonstrate a prototype system that can make real-time, 3-D measurements of position, attitude, surface deformations and can quantify fluctuations of these parameters in a relevant operational wind tunnel environment.

PHASE III DUAL USE APPLICATIONS: Military Applications:  Applicable to subsonic, supersonic & hypersonic wind tunnels where position, attitude and/or surface deformation measurements are needed.
Commercial Applications: Applicable to ground based testing of commercial flight vehicles to include aircraft & missiles.


1. E. T. Schairer, et al., “Model Deformation Measurements of Sonic boom Models in the NASA Ames 9- by 7-Ft Supersonic Wind Tunnel,” AIAA Paper 2015-1913, 53rd AIAA Aerospace Sciences Meeting, January 5-9, 2015, Orlando FL.

2. Y. Le Sant, A. Durand, and M-C. Merienne, “Image Processing Tools Used for PSP and Model Deformation Measurements,” AIAA Paper 2005-5007, 35th AIAA Fluid Dynamics Conference and Exhibit, June 6-9, 2005, Toronto, Ontario, Canada.

3. J. S. Masters and K. E. Tatum, “Unstructured Mesh manipulation in Response to Surface Deformation, AIAA Journal, Vol. 54, pp 331-342 (2016).

4. T. W. Fahringer and B. S. Thurow, “3D Particle Position Reconstruction Accuracy in Plenoptic PIV,” AIAA Paper 2014-0398, 52nd AIAA Aerospace Sciences Meeting, January 13-17, 2014, National Harbor, MD.

5. B.C. Winkelmann, et al., “Application of Stereoscopic Imaging in Aerospace Ground Testing,” AIAA Paper 2014-2519, 30th AIAA Aerodynamic Measurement Technology and Ground Testing Conference, June 16-20, 2014, Atlanta, GA.

KEYWORDS: three-dimensional imaging, high-speed imaging, wind tunnel testing, transonic wind tunnels, test article attitude, surface deformation, 3-D test model position, dynamic loads, wind tunnel model position measurement, stereoscopic imaging


TITLE: Reduction/Elimination of Unsteady Aerodynamic Loads on Model and Balance Support Systems in Large Scale Wind Tunnels


OBJECTIVE: Develop a technology for effective damping or elimination of large oscillations in wind tunnel model balance support systems.

DESCRIPTION: Modern wind tunnel test articles involve models and balance support systems for force and moment measurements. Models are typically installed onto a compact internally mounted balance and supported by the sting at the other end. Models sizes can range between a small weapon such as a 5% scale 250 pound bomb to a 10% scale model aircraft ranging between 200 and 800 pounds. Loads on the larger models are limited to the balance capacity which can average around 3,000 lb of force, 10,000 in-lb of moment. Conditions in the wind tunnel are below 4,000 psf total pressure and less than 140 degrees F. Unsteady flow generated on models can result in vibration and oscillation of the model support system. Such motions, whether generated by the aerodynamics alone and/or in coupled with support system structure, result in contaminated data due to the unexpected motion of the model support system. Certain small to moderate unsteady data can be post processed to improve accuracy of such data. Under certain conditions a model may experience unexpected unsteady loads which exceed safe operational limits of the model support systems’ components. Such conditions, often due to the presence of oscillatory dynamics, may result in significant and expensive damage or loss of components and waste of valuable test time. An example includes rolling moment snaps from abrupt wing stall. Current vibration reduction and isolation techniques involve tuned vibration dampers, viscoelastic material, and active damping with piezoceramic material. While these techniques can provide favorable results in certain situations, their use is typically limited to specific models, configurations and test conditions.

The overall goal of this topic is to develop innovative and practical technology solutions to reduce/limit the amplitudes of oscillations experienced by a balance, using adaptive and or active control techniques. The proposed technology must be feasible for various model and balance support systems and effectively eliminate the potential for serious damage to the systems. It also must be easily adaptable to different model sizes, configurations and test conditions without extensive preparation and build up.

PHASE I: Demonstrate feasibility of an instrumented model support system for damping or eliminating large oscillations in a relevant laboratory environment.

PHASE II: Develop an interactive (adaptive and/or automated) prototype, including mechanical and control software and demonstrate the system in an AEDC large scale wind tunnel or other representative operational environment.

PHASE III DUAL USE APPLICATIONS: Military Applications:  This technology is applicable to wind tunnel testing of legacy and next generation air platforms.

Commercial Applications: This technology is applicable to a large array of Commercial/University wind tunnel facilities across the U.S.


1. Reduction of Wind Tunnel Model Vibration by Means of a Tuned Damped Vibration Absorber Installed in a Model; Igoe, W.B. and Capone, F.T.; NASA-TMX-1606, July 1968.

2. Passive and Active Damping Augmentation Systems in the Fields of Structural Dynamics and Acoustics; Freymann, R”; AIAA-1989-1196.

3. The Use of Friction Springs for Damping Model Vibrations; Fuykschot, P.H.; STA Proceedings, April 1999.

4. Active Damping of Sting Vibrations in Transonic Wind Tunnel Testing; S. Balakrishna, D. H. Butler, R. White, and W. A. Kilgore; AIAA- 2008-0840.

5. Design and Performance of an Active Damper for NASA Common Research Model; S. Balakrishna, D. H. Butler, Michael J. Acheson, & E. Richard White; AIAA-2011-953.

KEYWORDS: Model active dampening, passive dampening, force balance protection, sting dampening, model dynamics, unsteady loads, wind tunnel testing


TITLE: Miniature Video Camera Technology for Embedded Probes

TECHNOLOGY AREA(S): Battlespace, Electronics, Sensors

OBJECTIVE: Development of a miniature video camera technology with the appropriate form factor and optical performance for embedding into small water-cooled probes for evaluating and monitoring turbine engine augmentor performance.

DESCRIPTION: AEDC develops and implements multi-purpose probes that are inserted into the hot, high speed, exhaust streams of turbine engines in order to measure pressure, temperature, velocity, species concentrations, and other performance characteristics. In particular, development of viewing probes have allowed unprecedented optical access to the turbine engine augmentors, but current needs far exceed the capability of available camera technologies. Camera technologies are unavailable that meet the requirements for size, form factor, optical quality and frame rates for embedding into probes used for augmentor performance evaluation, health monitoring and diagnosis. Spectral ranges have been limited to the visible, but short and long wavelength infrared images are needed, but not with a single camera technology.Camera technologies are required that meet the specifications in the table below:

Design Characteristics: Threshold(T), Objective(O), Stretch Goals(S)
Number of pixels: 640 X 480(T), 1280 X 960(O), 2048 X 2048(S)
Frame Rate: 30 fps interlaced(T), 100 fps non-interlaced(O), 1000 fps non-interlaced(S)
Form factor: Cylindrical(T), Cylindrical(O), Cylindrical(S)
Outer Diameter: 0.25 inches(T), 0.125 inches(O), 0.10 inches(S)
Length: 1.65 inches(T), 1.0 inches(O), 0.875 inches(S)
Spectral Range: Visible (T), Visible & near IR(O), Visible to LWIR(S)
Performance Life (at relevant conditions): 1000 hours(T), 1500 hours(O), 2000 hours(S)

Innovative development techniques will be required to achieve the required form factor, size and camera performance. The innovative camera technologies will replace existing small diameter NTSC cameras. The camera technology must have automatic gain control and adjustable shutter speeds. Since the cameras will be embedded into probes inserted into the harsh environment of augmented engine exhaust flows, the camera must be rugged and robust (vibration insensitive for temperatures up to 195 ºF). Manual adjustments over shutter speeds and gains must also be available. The camera systems must have options for multiple lens with image steering capability up to 90 degrees.

PHASE I: Develop a proof-of-principal camera technology design capable of satisfying all objective requirements. A laboratory demonstration proving feasibility that the prototype components meet the vibrational and temperature requirements for extended periods of time (2 hours) is desirable.

PHASE II: Develop a prototype camera system satisfying all objective requirements and/or stretch goals and demonstrate camera performance life of 100 hours without failure in a relevant environment.

PHASE III DUAL USE APPLICATIONS: Formalize the production process and design the appropriate machinery/infrastructure to support full-scale commercial production. 
Military Applications: Monitoring of turbine engine augmentor operation. Gas turbine combustor monitoring. Rocket engine combustor monitoring. Video surveillance.
Commercial Applications: Gas turbine combustor monitoring. Rocket engine combustor monitoring. Video surveillance.


1. Hiers, R. S. Jr. and Hiers, R. S. III, “Development of High Temperature Image Probes for Viewing Turbine Engine Augmentors,” AIAA-2002-2912, 22nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference, St. Louis, Missouri, 24 - 27 June 2002.

2. Hiers, R. S. III and Hiers, R. S. Jr., “Development of Exit-Plane Probes for Turbine Engine Condition Monitoring,” AIAA-2002-4304, 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Indianapolis, Indiana, 7-10 July 2002.

3. Beitel, G. R., Jalbert, P. A., Plemmons, D., Hiers, R. S., and Catalano, D. R., “Development of Embedded Diagnostics for Internal Flow-Field Measurements in Gas Turbine Engines,” AIAA 2004-6865, AIAA/USAF Developmental Test & Evaluation Summit, Woodla

KEYWORDS: viewing probes, turbine engines, diagnostics, augmentors, miniature camera, Visible, LWIR


TITLE: Non-Intrusive, Time-Resolved Turbulence Measurements in Large-Scale Hypersonic Wind Tunnels


OBJECTIVE: Develop and validate a non-intrusive, time-resolved instrument for turbulence measurements in large hypersonic ground test facilities.

DESCRIPTION: The non-intrusive measurement of turbulence is a critical capability missing in Hypersonic T&E facilities. Such facilities can measure forces, moments, and surface quantities to support the validation of computations, but do not provide a complete understanding of the hypersonic flow physics needed to improve modeling. Moreover, the flow characterization of any wind tunnel involves knowledge of its freestream turbulence, which is composed of vorticity (incompressible turbulence), entropy (total temperature) and acoustic (tunnel noise) fluctuations. To date, the characterization of free stream fluctuations (tunnel noise) in large scale hypersonic wind tunnels has been limited to high frequency Pitot pressure measurements. Such measurements are performed behind a normal shock which distort the spectral content and the magnitude of the fluctuations as a function of the probe tip geometry (1). Recent boundary layer stability measurements and analysis performed at AEDC Tunnel 9 (2) have shown that 2nd mode transition can be predicted using Mack’s amplitude method. The accuracy of this method is dependent on the accuracy of the tunnel noise measurements at high frequencies.

A measurement technique that can discriminate between the disturbance modes (vorticity, entropy and acoustic) is highly desirable. Particle Image Velocimetry (PIV) and Laser Doppler Velocimetry (LDV) are thwarted by the high stagnation temperature of the flow and the impracticality of seeding the core flow. Approaches such as Rayleigh scattering (3), interferometry (4), and Schlieren methods (5) show promise, but have yet to be successfully applied for turbulence measurements in large scale hypersonic T&E facilities. The goal of this effort is to develop and demonstrate an instrument suitable for non-intrusive turbulence measurements in large-scale hypersonic wind tunnels. An instrument suite would also be acceptable, considering the multimodal nature of the freestream disturbances in hypersonic wind tunnels.

Typical test conditions for the USAF AEDC Hypervelocity Wind Tunnel 9 are: Mach number 8 to 14, run time 0.25 to 5 seconds, stagnation temperature up to 1850 K, static temperature 40 to 200 K, static pressure 3.5 to 10,000 Pa, static density 0.0003 to 0.56 kg/m3 , and velocity 1370 to 2070 m/s. Test section optical access is typically gained through two thick (50 mm) BK7 glass windows located 1.5 to 2 m apart, but smaller diameter (thinner) inserts with other optical materials can be implemented. The test gas is nitrogen. Hypersonic wind tunnel facilities produce low-frequency vibrations during testing which must be tolerated by the proposed instrument. Measurements should resolve turbulent frequencies up to 1 MHz to provide useful data for boundary layer transition. While this solicitation does not require “global” data, multiple point measurements are required to obtain turbulence length scale (without needing Taylor’s hypothesis) and convection velocity magnitudes and directions. Several (~20) simultaneous point measurements are needed without traversing the instrument, in order to characterize turbulence across a boundary layer in a single run. The measurement volume for each point should be less than 1 mm in size in all directions and the distance between measurement points should be as low as 1-mm. Finally, it is very important that the instrument be able to reject the sidewall boundary-layer located over the test cell windows.

PHASE I: Demonstrate the feasibility of a non-intrusive instrument for high-speed flow turbulence measurements, preferably in a small-scale supersonic or hypersonic wind tunnel or similar environment.

PHASE II: Develop a non-intrusive, real-time turbulence prototype measurement system and demonstrate the prototype in large scale hypersonic wind tunnel.

PHASE III DUAL USE APPLICATIONS: The measurement technology can be marketed for non-intrusive turbulence measurements in high-speed wind tunnels and other high-speed flow applications.


1. R. S. Chaudhry, G. V. Candler. Recovery of Freestream Acoustic Disturbances from Stagnation Pressure Spectrum in Hypersonic Flow, AIAA Paper 2016-2059.

2. E.C. Marineau, Prediction Methodology for 2nd Mode Dominated Boundary Layer Transition in Hypersonic Wind Tunnels, AIAA Paper 2016-0597.

3. A. F. Mielke, R. G. Seasholtz, K. A. Elam, J. Panda. Time-average measurement of velocity, density, temperature, and turbulence velocity fluctuations using Rayleigh and Mie scattering. Experiments in Fluids 39 (2): 441-454, 2005.

4. N. J. Parziale, J. E. Shepherd, H. G. Hornung. Free-stream density perturbations in a reflected-shock tunnel, Experiments in Fluids 55(2): 1-10, 2014.

5. S. Garg, G. S. Settles. Measurements of a supersonic turbulent boundary layer by focusing schlieren deflectometry, Experiments in Fluids 25 (3): 254-264, 1998.

KEYWORDS: turbulence, hypersonic, instrumentation, non-intrusive, seedless, wind-tunnel


TITLE: Advanced, Low-Erosion Electrode Technologies for Arcjet Testing

TECHNOLOGY AREA(S): Materials/Processes

The technology within this topic is restricted under the International Traffic in Arms Regulation (ITAR), 22 CFR Parts 120-130, which controls the export and import of defense-related material and services, including export of sensitive technical data, or the Export Administration Regulation (EAR), 15 CFR Parts 730-774, which controls dual use items. Offerors must disclose any proposed use of foreign nationals (FNs), their country(ies) of origin, the type of visa or work permit possessed, and the statement of work (SOW) tasks intended for accomplishment by the FN(s) in accordance with section 5.4.c.(8) of the solicitation and within the AF Component-specific instructions. Offerors are advised foreign nationals proposed to perform on this topic may be restricted due to the technical data under US Export Control Laws. Please direct questions to the AF SBIR/STTR Contracting Officer, Ms. Gail Nyikon,

OBJECTIVE: Develop advanced, cost-effective, arc-heater electrode materials and demonstrate significant reduction in erosion rates under typical operating conditions as compared to traditional electrode materials.

DESCRIPTION: Development of advanced hypersonic systems often requires arc-heater-based ground test facilities such as those at the AEDC, which can provide mission-representative high-enthalpy test environments. Development plans for the AEDC arc heater facilities call for expansion of the test environment envelope to broaden mission support, thereby placing considerably higher demands on arc heater subsystems and components, particularly the electrodes which deliver current to and from the high pressure air flow. The very high current density in the electrode arc attachment region, driven by both high arc currents and arc constriction associated with high pressure flow, inevitably causes erosion of electrode material, contaminating the flow and creating conditions conducive to arcing across insulator surfaces. Over the 35+ year history of AEDC arc heater operations, many improvements in electrode design, material selection, cooling techniques, etc., have been implemented to control flow contamination levels and maximize heater component reliability. However, the planned expansion of the test environment envelope will require increases in arc current, flow pressure, and run time, representing a significant increase in the severity of the electrode operating environment.

Individual arc heater runs, operating at thousands of amperes and up to 120 atmospheres of chamber pressure, can last up to several minutes. Because of this, the arc heater electrodes are nominally replaced after about 10 minutes of accumulated run time. Desired operating conditions include significant increases in run times, current, and pressure, which will significantly increase the severity of the thermal environment at the electrode-arc attachment region and likely reduce electrode lifetime to unacceptable levels. Forced electrode cooling and other mitigation measures have been optimized in terms of benefit to the arc attachment region, suggesting that new electrode materials, capable of handling higher current densities while retaining thermal and electrical conductivity similar to copper, be investigated as a possible remedy. Candidate materials/electrodes will be tested at typical arc heater operating conditions to evaluate ampacity, mass loss rate, and other performance parameters as compared to legacy materials/electrodes. After successful demonstration of improvement at typical operating conditions, candidate electrodes will be operated at higher currents, pressures, and run times to explore their expanded operational limits.

PHASE I: Research candidate composites and materials which lend themselves to production for use as electrodes and/or electrode liners. Produce test samples of selected materials and demonstrate ampacity and thermal/electrical conductivity which meet or exceed that of zirconium copper.

PHASE II: Develop prototype electrodes from candidate material(s) that meet/exceed specifications in description.  Dimensions of the electrodes used in the AEDC arc heaters do not exceed 8" in diameter and 6" in length. Demonstrate reduction in mass loss rate as compared to that of zirconium copper and the ability to survive the harsh environment in the AEDC Arc Heater H1 or H3.

PHASE III DUAL USE APPLICATIONS: Military Applications: Electrode Liners for AEDC's H1, H3, and MPAH (Mid Pressure Arc Heater), lighter electronics for any military device, possible EMP resistant devices.
Commercial Applications: Electrical applications involving contacts/electrodes. Conductors in any small electrical device carrying relatively large currents.


1. Subramaniam, Chandramouli, Takeo Yamada, Kazufumi Kobashi, Atsuko Sekiguchi, Don N. Futaba, Motoo Yumura, and Kenji Hata. "One Hundred Fold Increase in Current Carrying Capacity in a Carbon Nanotube–copper Composite." Nature Communications Nat Comms 4 (2013): n. pag. Web. 7 Mar. 2016.

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