ARMD Seedling Semiannual Report


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NARI Seedling Fund – Final Technical Report
Liquefied Bleed for Stability and Efficiency of High Speed Inlets

Project WBS Number: 694478.02.93.02.13.52.22

Investigators: J. David Saunders and Dr. David Davis, NASA Glenn Research Center, Inlet and Nozzle Branch; Dr. Stephen J. Barsi
and Dr. Matthew C. Deans, NASA Glenn Research Center, Propulsion and Propellants Branch; Lois J. Weir and Bobby W. Sanders, TechLand Research, Inc.

Purpose
Through Phase 1 Seedling Fund support, a novel concept entitled, “Liquefied Bleed for Stability and Efficiency of High Speed Inlets” was investigated. The purpose of this 1year effort was to quantify the design parameters and benefits of the liquefied bleed (LB) concept. The potential payoff is enhanced system performance of inlets for highspeed aircraft. Elements of the task include:
1) development of a physical configuration with thermal balance sufficient to cool the bleed air
2) development of a scheme to capture, process, and prevent icing conditions of the liquefied bleed air
3) analysis for first-order effects of cooling on the outflow of the bleed air
4) measurement of the bleed air thermal state (temperature and liquefaction fraction)
5) measurement of the resulting core (propulsive flow path) boundary-layer flow quality
The Phase 1 Seedling Fund effort focused on the first three elements by assembling tools, conducting preliminary analyses, and preparing test plans.

A photograph of a hypersonic research inlet referred to as the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX) is shown in Figure 2. This test article is a fully integrated Turbine-Based Combined-Cycle (TBCC) propulsion system and is currently undergoing testing in the NASA Glenn Research Center (GRC) 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) at conditions approximating Mach 4 flight. Initial results from this test program are reported in References 1 and 2.

Background
Bleed air has traditionally been used for mixedcompression inlets at moderate to high supersonic flight speeds for stability and improved performance (see Figure 1).
Figure 1. Mixed-compression inlet with bleed.
At higher speeds, however, the increased flight enthalpy causes extreme difficulties in using bleed. The bleed air drawn off of the propulsive flow path is at both high temperature and low pressure. Bleed ductwork becomes hot and large, and the bleed air produces increased vehicle drag when exhausted overboard. Therefore, uncooled bleed causes increased drag for supersonic propulsion and cannot effectively be used for hypersonic engine designs.

Figure 2. CCE LIMX installed in the NASA 10x10 SWT.
The series of pipes shown in the upper left of the photo are used to exhaust bleed air from the inlet. Sized to provide effective bleed at Mach 4 conditions, the frontal area of the bleed pipes are comparable to the engine ducts themselves.*
The above example serves to illustrate the difficulty of providing effective uncooled bleed at flight speeds of Mach 4 and higher. However, if the bleed air were to be cooled, the ductwork would become cooler and the size requirement would be reduced. Cooling of the bleed air would be provided by the cryogenic fuels that are required
*Note that this research inlet model did not have the bleed ducting sized to minimize frontal-area drag as would be done in a flight inlet design. Nevertheless, the photograph does qualitatively indicate the impact of bleed ducting.
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by high-speed vehicles. The current research effort is focused on investigating the feasibility and effect of intensely cooling the bleed air to the point of liquefaction.
To date, prior studies have focused on quantifying the relationship between bleed flow rate and bleed duct pressure (Reference 3). A small subset of this research investigated the effect of wall cooling on boundary-layer development (Reference 4). The wall-cooling studies were not coupled to bleed flow extraction nor were the systemlevel benefits of liquefied bleed examined. The potential benefits of the cooling concepts include greatly reduced bleed duct sizing, reduced shock-wave/boundary-layer interaction extent, and oxidant storage for other phases of a flight mission. The concept will enable greatly improved performance and robustness of airbreathing propulsion and thereby enable new design options for high-speed flight.
Approach
For the Phase 1 effort, the technical approach addressed three key areas. First, the physical configuration and thermal balance requirements were analyzed. One of the issues considered was whether the air should be liquefied as it passes through the perforated bleed plate, or should it pass through a conventional bleed plate before entering a separate heat exchanger? The former may result in a more compact, lighter system, but the latter allows for the use of conventional bleed plates whose performance is well characterized, but may require their own cooling system at high Mach numbers. To keep the Phase 1 effort simple, the initial experiments were planned with a decoupled bleed plate and heat exchanger. Because of the slightly different (~57 °F) condensation temperatures for nitrogen and oxygen, the potential for the generation of liquid oxygen and the safety implications was also considered.
The second key area in Phase 1 was to conduct a simplified system study to understand the impact of a liquid bleed cryogenic subsystem on a high-speed vehicle. Estimates of weight and other cost/benefit trades were incorporated into a candidate vehicle concept.
Finally, in order to verify the liquid bleed concept, two proof-of-concept (POC) experiments were planned as part of the Phase 1 effort. These tests are anticipated to be conducted as part of a Phase 2 effort. The first experiment is a small-scale test with a limited number of bleed holes. This test is planned to be conducted in a cryogenics lab at NASA GRC where the infrastructure for handling cryogenic hydrogen and compatible systems is already in place. A small axisymmetric (3.0-in.-diameter) Mach 3 SWT has been designed and partially fabricated to support this test. The second test is a larger scale experiment that includes a complete bleed region with an impinging reflected shock wave as shown in Figure 3.

Figure 3. 1- by 1-Foot SWT with shock and bleed.
This test will be conducted in the NASA GRC 1- by 1-Foot SWT (11 SWT), which has a Mach number capability from M = 1.3 to 6.0 and has been used for bleed research in the past. A cryogenic fluid cooling system would need to be added to the tunnel for the proposed cooled bleed experiment. One of the advantages of this facility is the large bleed plenum (Figure 4), which can easily accommodate the liquid-bleed subsystem.
TUNNEL FLOW
BACKSIDE OF BLEED PLATE BLEED PLENUM CONVENTIONAL BLEED EXHAUST (TO 450 PSIG EJECTOR SYSTEM)
Figure 4. 1x1 SWT bleed plenum.
To address the significant challenge of test operations with cryogenic fluids, we are also investigating alternative facilities at the GRC Research Combustion Laboratory (RCL) as well as facilities outside NASA. As will be discussed, later, all these facility options would require modifications with cost impact.

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NARI Seedling Fund – Final Technical Report
Accomplishments
The Liquefied Bleed project progressed well through the technical feasibility and test planning phases. The following five accomplishments are detailed: 1) Feasibility, including thermal balance was confirmed, 2) Potential test facilities have been identified, 3) Mission analyses are encouraging, 4) Detailed test planning was completed, and 5) Test cost estimates were received.
Status of Research
Semi-monthly meetings were conducted since February 2013. These meetings coordinated the research analyses and plans between the researchers in two groups: supersonic inlets and the cryogenic propellants. Additional meetings occurred to gather cost information based on the test plans. Figure 5 shows a previous small rocket test in the Altitude Combustion Stand (ACS), which is one of the facilities being considered for the small-scale POC test.
Figure 5. Cryogenic rocket testing at the ACS of the RCL.
Using experience from the TBCC mode transition inlet design, a range of bleed flow rates and fluid states were used to size the heat exchanger and facility equipment. From this bleed environment, an analysis was conducted to determine the thermodynamic effect of cryogenic cooling on the air density and liquefaction fraction. In Figure 6,

plots of the bleed air's possible exit conditions as a function of hydrogen exit conditions are shown. In both cases, supercritical hydrogen is entering the heat exchanger at 225 psia and 36 R.
Figure 6. Bleed air exit conditions. At a moderate liquid hydrogen flow rate, more than a 600% increase in bleed air density may be expected. If the hydrogen-to-bleed-air ratio is more aggressive, complete air liquefaction is possible. These results support concept feasibility. Heat exchanger sizing, design, and detailed test planning were based on these analyses. The thermal analysis effort is discussed in further detail in Appendix A. The challenge of cryogenic hydrogen safety has led the research team to baseline some of GRC’s propulsion and cryogenic fluid facilities in our test plans. Two test areas were considered, the ACS and Small Multi-propose Research Facility (SMiRF). Both are capable of testing cryogenic propellants. However, SMiRF is actively working with liquid hydrogen, whereas the fuel system for ACS is currently setup for liquid methane. The liquid hydrogen infrastructure was considered more valuable. Therefore, the SMiRF facility is preferred for the cryogenic testing. As a consequence, the research team will plan for two possible test entries: 1) A POC test in a cryogenic (liquid hydrogen) facility. 2) Bleed testing in a larger wind tunnel, such as the 11 SWT (Refs. 5 and 6). These two entries could be conducted in parallel. For the POC cryotest, the research team elected to develop a small supersonic wind tunnel with a circular cross-section: diameter of 3 in. or a test section area of about 7 in.2 A conceptual sketch of the proposed small supersonic wind tunnel and associated heat exchanger for bleed liquefaction is shown in Figure 7a. This heat exchanger design provides a large enough size to maintain minimum wall thickness for the fins and to install critical instrumentation.
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150 psig

Plenum Tank (Pre-Existing)

Test Section Inflow

Bleed Air Outlet

M=3 C-D Nozzle

Bleed Ring

Heat Exchanger

Liquid Air Level Probe

Liquid Air Capture Chamber

Exhaust to ATM

LH2 Outlet
LH2 Inlet Test Section Outflow

a) Sketch of the Liquefied Bleed experiment in SMiRF.

accommodates a larger fuel fraction. For the preliminary analysis, the liquefied bleed air was assumed to be liquefied but not stored for later use. Other simplifying assumptions such as addressing thermal balance issues for the dual-mode ramjet or scramjet (DMRJ) flow path and optimal sizing of the TBCC aircraft can be investigated in the future. For the current analysis, baseline aircraft had a very large TOGW of 2106 lbs. Figure 8 shows a preliminary result. More details are given in Appendix B.

b) Bleed exhaust plenum with heat exchanger installed in the 11 SWT.
Figure 7. Heat exchanger installations in the two proposed test facilities for liquefied bleed.
Figure 7b depicts a cross section of the bleed exhaust plenum with a notional design of a heat exchanger installed in the 11 SWT. The sidewall of the wind tunnel test section is represented in blue at the bottom of the figure. The light blue piece is the (existing) tunnel sidewall bleed insert plate. Bleed airflow is removed from the freestream test section flow through the bleed insert plate. A plenum is mounted on top of the bleed pocket, capturing the removed airflow and delivering it into the heat exchanger. The notional heat exchanger is a basic “shell and tube two-pass straight-tube” heat exchanger.
For mission analysis, a FORTRAN-based simulation of a representative TBCC mission has been developed in order to capture the first-order effects of a LB system. The spaceaccess mission that was modeled simulates an airbreathing TBCC-propelled first stage that could boost a second stage to Mach 7. The effect of a liquefied bleed system was assessed by assigning appropriate performance, weight fraction factors, and assumptions to find the resulting minimum take-off gross weight, TOGW. The liquefied bleed benefits are increased supersonic inlet recovery and reduced bleed drag and volume. The volume change
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Figure 8. Results from mission analysis showing a 23% reduction in TOGW due to liquefied bleed.
This plot indicates that liquefied bleed could reduce TOGW by up to 23%. This benefit was accentuated by a pinch point at the turbine-to-ramjet mode transition Mach number. In fact, without bleed, positive acceleration could not be maintained. Higher fidelity studies with a more optimal vehicle could change the magnitude of the result, but the suggested benefit of liquefied bleed is significant. During the first 6 months, the efforts were focused using a facility certified for liquid hydrogen experiments. Making a small supersonic wind tunnel was found to be easier than certifying a facility for liquid hydrogen safety. The GRC RCL provides a safe test environment for use of liquid hydrogen. Use of the 11 SWT will still be considered but for non-chilled and LN2-cooled bleed experiments. Two cryogenic test cells were being considered: ACS (Figure 5) and SMiRF (Figure 9).
Figure 9. Exterior view of SMiRF, the Small Multipurpose Research Facility.

NARI Seedling Fund – Final Technical Report Under Phase 1 seedling funding, some of the hardware for the liquefied bleed testing were fabricated. The initial hardware can be set up for a simple supersonic flow test without the heat exchanger. This quick test would assess the flow quality of the Mach 3 nozzle/diffuser design. Figure 10 shows this hardware in relation to the design schematic.
Figure 10. Mach 3 nozzle / diffuser hardware fabricated under the Phase 1 Seedling Fund. Requirement documents for both tests have been submitted to the GRC facilities personnel for feedback and planning purposes (References 5 and 6). As previously mentioned, the SMiRF facility was chosen to develop rough-order-ofmagnitude costs due to its liquid hydrogen capabilities. The promise for additional support outside of the Seedling Fund is a long-term proposition. High-speed aeronautics was unable to fund GRC for hypersonics for the previous fiscal year, (FY13). The Air Force and NASA are collaborating by funding the next phase of TBCC testing for the current fiscal year. Liquefied bleed technology was discussed with the Air Force Research Lab as a possibility for the far-term plans of TBCC/CCE project.
Future Plans
The rough-order-of-magnitude cost information derived from the test requirements for two distinct tests was ascertained. Each of the two tests would require about twice the available resources through the Phase 2 Seedling Fund. Therefore, the Liquid Bleed (LB) effort will not be proposed for immediate continuation. At a reduced staffing level, the next steps will be: investigation of ways to reduce costs, evaluation of test facilities, investigation of alternate fund sources, and re-scoping of test objectives. The future work will lead to a robust future proposal to continue the LB investigation.

Current TRL: The Technology Readiness Level (TRL)
= 2: The technology concept and application are being formulated. POC testing has been identified, designed, and planned.
Applicable NASA Programs/Projects
Benefits of the Liquefied Bleed technology would be applicable to the Fundamental Aeronautics Program/HighSpeed Project. A direct application would be for hypersonics research, a project that is currently on hold. A long-term application of the technology may also be for a future fleet of supersonic aircraft that have transitioned to cryogenic methane or hydrogen to reduce carbon emissions.
References
1. Foster, L.E.; Saunders, J.D., Jr.; Sanders, B.W.; Weir, L.J.: "Highlights from a Mach 4 Experimental Demonstration of Inlet Mode Transition for Turbine-Based Combined Cycle Hypersonic Propulsion." 48th Joint Propulsion Conference and Exhibit; 30 Jul. 1 Aug. 2012; Atlanta, GA, NASA/TM-2012-217724, December 2012.
2. Saunders, J.D.; Foster, L.E.; Sanders, B.W.; Weir, L.J. "Demonstration and Performance of Inlet Mode Transition for Turbine-Based Combined Cycle Hypersonic Propulsion." JANNAF Propulsion Meeting/33rd Airbreathing Propulsion/Joint Subcommittee Meeting; 3-7 Dec. 2012; Monterey, CA; Limited by International Traffic in Arms Regulations (ITAR); NASA/TM-2013217839, May 2013.
3. McLafferty, G.: Pressure Losses and Flow Coefficients of Slanted Perforations Discharging from within a Simulated Supersonic Inlet. UTRC R-0920-1, Dec. 1958.
4. Schlichting, H., (Kestin, J. transl.): Boundary-Layer Theory, Seventh Edition, Chapter XXIII. McGraw-Hill Book Company, 1979.
5. Davis, D.O; “Liquefied Bleed for Stability and Efficiency of High Speed Inlets: Small-Scale Liquid-Bleed (SSLB) Test Requirements” , Revision 0, Jan. 13, 2014.
6. TechLand Research Inc.; “Liquefied Bleed for Stability and Efficiency of High Speed Inlets: Bleed Cooling Test in the NASA GRC 1x1 SWT Test Requirements”, Revision 0, Dec. 2013.
7. Cassidy, M.D.; “Performance Sensitivities of a High Altitude Mach 5 Penetrator Aircraft Concept”, NASA CR-3932, 1985.
8. Cubbison, R.W., and Barnett, D.O.; “Performance Characteristics of a Wing-Body Combination with a Two-Dimensional ExternalInternal-Compression Inlet at Mach 3.5 and 3.0”, NASA TM-X-256, July 1960.
9. Hehs, E.: “Super Hustler, FISH, Kingfish, And Beyond: Part 4: Beyond Kingfish”, Posted 9 March 2012, http://www.codeonemagazine.com/article.html?item_id=92 , Accessed April 14, 2014.
10. Hill, Phillip G.; and Petersen, Carl R.; Mechanics and Thermodynamics of Propulsion, Addison-Wesley, 1970.
Patents
The technology developed during the Phase 1 funding was proposed for patent activity through a new technology disclosure form.

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Nomenclature

11 SWT 1010 SWT Ableed Ainf Aref, Ao Aspill ACS BL CCE LIMX
Cd Cd0 (or Cd0 )
Cl DMRJ DWF f FWF FY HEX Isp L/D LB LH2 LN2 LOX

1- by 1-Foot Supersonic Wind Tunnel 10- by 10-Foot Supersonic Wind Tunnel Area of bleed porous holes Area of capture stream tube Inlet reference area Area of spillage air stream tube Altitude Combustion Stand baseline Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment drag coefficient
zero-lift drag coefficient
lift coefficient dual-mode ramjet dry weight fraction function fuel weight fraction fiscal year heat exchanger specific impulse lift-to-drag ratio liquefied bleed liquid hydrogen liquid nitrogen liquid oxygen

lsf M M0, M Mloc Mstage m N NIST
NB Pbleed Pt0 (or Pt0 )
Pt1 (or Pt1 )
Pt2 (or Pt2 )
POC PWF q RCL SMiRF SWF T m a Tbleed TBCC TOGW TRL TSTO

linear scale factor Mach number Freestream Mach number Local Mach number stage Mach number mass flow rate number of turbine engines National Institute for Standards and Technology No bleed bleed air plenum pressure Freestream total pressure
Inlet entrance total pressure
Inlet exit total pressure
proof of concept payload weight fraction flight dynamic pressure Research Combustion Laboratory Small Multi-propose Research Facility structure weight fraction specific thrust, thrust per inlet airflow
bleed air temperature Turbine-Based Combined-Cycle take-off gross weight Technology Readiness Level Two-Stage-To-Orbit

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NARI Seedling Fund – Final Technical Report
Appendix A. Thermal Balance Analysis

Introduction
One of the first steps to understand the feasibility of the liquefied bleed (LB) concept is to analyze the thermal balance between the cooling capacity of the cryogenic fuel and the bleed flow in the inlet. A series of assumptions were incorporated into the thermodynamic analysis. Results are presented for a typical bleed region showing the amount of density increase as a function of fuel exit temperature.

m , tunnel (lbm/s) 1.904 1.904

m , bleed (lbm/s) 0.2855 0.2855

m , hydrogen (lbm/s) 0.0231 0.0555

Pbleed (psia)
1.137 1.137

Tbleed (R)
530 530

Table A.1. Conditions analyzed for a LB system.

Approach
The baseline liquefied bleed system was chosen based on the R1 bleed region in the CCE LIMX inlet, References 1 and 2. This forward bleed region has particularly low bleed plenum pressure. R1 bleed would both benefit greatly from a LB system and provide the greatest challenge to thermal balance because of low pressure within which the heat exchanger would operate. The analysis assumed that the fuel was liquid hydrogen. The hydrogen from the fuel tank would enter the LB heat exchanger at 30 R. This inflow fuel temperature maximizes the cooling capacity. In a real system, the hydrogen would likely enter the heat exchanger at a higher temperature, which would reduce the cooling effect on the bleed air. Future refinement would vary this temperature to understand the upper bound on the inlet temperature.
For the thermal balance analysis, the hydrogen exit temperature is varied from this inlet temperature of 30 to 540 R. At each exit temperature, the hydrogen enthalpy change is computed. Then, the bleed air enthalpy change is computed using an energy balance. Given the final enthalpy of the air and the pressure, the temperature and density can be evaluated.
These thermodynamic properties are evaluated by interpolating NIST (National Institute for Standards and Technology) tables. Dry air density and enthalpy tables were created from the NIST database. Also, para-hydrogen enthalpy tables were created from the NIST database. Hydrogen is assumed to be an equilibrium mixture of parahydrogen and ortho-hydrogen with different thermodynamic and transport properties. [Equilibrium hydrogen is >99.9% para at 36 R and ~25% para at 540 R. Without a catalyst, para-ortho conversion is slow.]
Conditions for the R1 bleed are given in Table A.1. In the first row, the hydrogen flow used to cool the bleed air was set to 41.67% of stoichiometrically-available hydrogen. The second row is set to use 100% of stoichiometricallyavailable hydrogen for bleed air flow cooling.

Results Figure A.1 shows an example of the data compiled from the NIST database for dry air. Note that the triple point pressure for air is 0.91 psia and the triple point temperature is 107.6 R. So the data in the figure are just about the triple point pressure. Below triple point conditions, the phase transition will be from gas directly to solid.
Figure A.1: Dry air thermodynamic properties. Using the dry air and hydrogen properties, the state of the bleed air exiting a heat exchanger can be calculated. In Figures A.2 and A.3, the results of the energy balance calculation are shown. For this analysis, the heat exchanger is assumed to be 100% efficient. Also, the hydrogen entering the heat exchanger at 225 psia is assumed to be supercritical.

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Figure A.2. Results of the thermal energy balance through a liquefying bleed heat exchanger. Air densification as a function of hydrogen exit temperature.
At the nominal hydrogen flow rate (41.67% of the stoichiometric flow), allowing the hydrogen temperature to increase to room temperature results in more than a 600% densification. At the greater flow rate, (100% of the stoichiometric flow), the bleed air can become fully liquefied.

Figure A.4. Results of varying the heat exchanger efficiency on the thermal energy balance. Air density as a function of hydrogen exit temperature.
Heat exchanger efficiency has a significant effect on densification. At 50% efficiency, even at higher hydrogen flow rate, exit air remains in the two-phase region.
Crystallization of water vapor and carbon dioxide is at least one source of inefficiency. Figure A.5 shows the phase diagram for carbon dioxide. At 1 psia, carbon dioxide will start crystallizing at ~300 R and water vapor will start crystallizing at ~452 R. Frost build-up on the heat exchanger increases the thermal resistance, resulting in less heat transferred from the bleed air to the hydrogen.

Figure A.3. Results of the thermal energy balance. Air exit temperature as a function of
hydrogen exit temperature.
From this nominal analysis, several variations were investigated. These included: variable heat exchanger efficiency, inflow hydrogen pressure, and inflow hydrogen temperature.
The effect of heat exchanger efficiency was investigated by reducing the efficiency by 50%. Results are presented in Figure A.4. Again, the hydrogen is assumed to enter the heat exchanger with supercritical properties at 225 psia.

Figure A.5. Results of varying the inflow hydrogen temperature on the thermal energy balance. Air density as a function of hydrogen
exit temperature.
The effect of inflow hydrogen pressure was investigated by reducing the pressure from 225 to 40 psia. Results are presented in Figure A.6.

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NARI Seedling Fund – Final Technical Report
Figure A.6. Results of varying the inflow hydrogen pressure to 40 psi on the thermal energy balance. Air density as a function of
hydrogen exit temperature.

Summary
The data indicate that under ideal conditions, bleed air will begin to liquefy.
For the nominal case (0.0231 lbm/s hydrogen), the air exiting the plenum would be a two-phase mixture of gas and liquid resulting in >600% increase in the air density.
The results are a strong function of the heat exchanger efficiency, and more detailed analysis and testing is required to quantify the inefficiencies. Ice frost from water vapor and carbon dioxide may be one source of this inefficiency.
For this bleed air condition (forward ‘R1’ bleed from the CCE LIMX test) significant densification can occur using the cryogenic fuel. Therefore significant reductions in the bleed ducting volume and bleed drag are feasible.

Note that the hydrogen enters as a subcritical fluid at 40 psia. This figure, when compared to Figure A.1, shows similar behavior as the supercritical case. This result is not surprising since hydrogen enthalpy is a weak function of pressure.
A final variation was examined by varying inflow hydrogen temperature (Figure A.7).

Figure A.7. Results of varying the inflow hydrogen temperature on the thermal energy balance. Air density as a function of hydrogen
exit temperature.
Hydrogen will warm as it flows from the supply tank and undergoes compression by the pump. The heat exchanger efficiency is kept at 100% for this analysis. Note that only a small resulting region exists where 100% of the bleed flow is liquefied. There is a large region of conditions where two-phase flow exists. As with the other trends, the density increase is not as large as when it is fully liquefied, but there is a significant increase.
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Appendix B. Mission Analysis Study
Approach
The approach to conducting a mission analysis to assess the liquefied bleed technology was broken down into the following steps:
1. Select mission and vehicle concept 2. Pick relevant assumptions 3. Develop a trajectory analysis tool 4. Conduct trade studies
Mission and Vehicle Concept Selection
The mission selected for this study was the Two-Stage-ToOrbit (TSTO) vehicle to deliver payload to low-earth orbit. This mission is similar to the direction taken by the Fundamental Aeronautics Program/Hypersonic Project and its predecessor project whose funding was terminated in 2012. These projects all used a TSTO vehicle in their goal vision. NASA-funded efforts have generally used hydrogen as a fuel because this allows wider Mach operating ranges than less volatile fuels. The liquid bleed technology applies to the first stage of the TSTO vehicle which uses two airbreathing engines in an ‘over/under’ configuration.
For take-off to Mach 4 conditions, advanced hightemperature-capable turbine engines would provide the accelerating thrust. Beyond Mach 4, propulsive thrust would transition to a DMRJ. The Mach 4 speed is termed the transition Mach number. The turbine engines are located above the DMRJ duct in a typical vehicle configuration leading to the ‘over/under’ propulsion scheme depicted in Figure B.1. Because this propulsion system uses turbine engines, it is referred as Turbine-BasedCombined-Cycle (TBCC) propulsion. The transitioning inlet system has been tested and discussed in References 1 and 2. Vehicle system studies which helped to guide this mission analysis are documented in References 7 to 9.
Figure B.1. A Turbine-Based-Combined-Cycle Vehicle with an ‘over/under’ engine configuration.
The second stage of the vehicle is not addressed explicitly in this study. It could be a pure rocket or a hybrid scramjet rocket with liquid oxygen (LOX) augmentation. This stage would contain the orbital payload and accelerate from the staging Mach number of 7 to orbital speed. The liquid bleed technology may be applicable to a scramjet-based second stage. However, its impact on the overall vehicle is thought to be greater for the first stage. Consequently, the impact of liquid bleed technology on the first stage is the focus of this study.
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This study’s analyses were held at a low-fidelity level, consistent with this effort’s scope and low TRL of the LB concept.
Assumptions
Funding levels limited the amount of detail that could be included in the analysis. Absolute predictions of mission performance were not the goal. Rather, a reasonable baseline was simulated from which relevant sensitivities could be understood. The sensitivity of the baseline mission to liquid bleed technology was the goal.
Many simplifications were assumed. The assumptions made for the analysis of the selected TSTO mission to a staging Mach number are listed below.
1. Baseline Vehicle Design 2. Vehicle Geometry 3. CCE LIMX flow splits and performance 4. Liquefied bleed densification 5. Effects due to liquefied bleed 6. Vehicle Dry Weight Fraction 7. Trajectory dynamic pressure 8. Simplified aerodynamics 9. Simplified engine performance 10. Trim drag
They will be discussed individually below. Many of these assumptions are not simple constants or functions. For these assumptions, details of their calculations are modeled within the MSExcel spreadsheet, “Access2Space_v5.xlsx.”
1. Baseline Vehicle Design This simplified analysis utilizes weight fraction analysis as outlined in chapter 10 of Reference 10. Weight fractions are defined for the payload, structure, dry, and fuel (PWF, SWF, DWF and FWF, respectively). The dry weight is the sum of the payload and structure. For the baseline vehicle, the payload was set to 8%, (PWF = 0.08). For the first stage, this ‘payload’ is the entire second stage. The structural weight fraction was optimistically set to 42%, (SWF = 0.42), so the dry weight fraction is 50%, (DWF = 50%). Note that the SWF include the ‘payload’ of the second stage. The remaining weight, 50% of the TOGW, is available for fuel, (FWF = 0.5), or 1 million pounds for our baseline vehicle.
The design parameters for the vehicle were chosen based on experience with an eye to avoid overly constrained or nonlinear areas of the mission space. In other words, the vehicle design is intentionally non-optimal to allow easier sensitivity analyses. Some of specific elements are listed below:
 0.5 is the baseline dry mass ratio or dry weight fraction, (DWF)
 Second stage (payload) mass ratio is part of the first stage’s DWF

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ARMD Seedling Semiannual Report