3.1. NOx Reduction Potential
While the concepts with conventional combustion chambers show no advancements in NOx and soot emissions, the advanced combustion technology shows significant reduction. At global fleet level, the NOx emissions could be reduced by 91% for the kerosene case and by 99.8% for the hydrogen concept compared with the reference D261. This strong reduction is mostly due to three features. Firstly, the combustion is operated at very lean conditions with the help of a wider stable operating range of the FLOX-based combustion technology compared to conventional swirl-stabilized combustion. Secondly, the authors assume that fully prevaporized fuel injection can be handled at an entry into service of 2040. Therefore, near stochiometric conditions around spray droplets are not present in the combustion, which helps to decrease the NOx emissions. Thirdly, due to the high momentum jets, the FLOX-based combustion achieves more homogeneous temperature profiles, reducing high temperature peaks and strong NOx sources, respectively.
Assumptions on the combustion’s lean blow off (LBO
) limits at different input fuels were derived from atmospheric combustion experiments. Measurements of different FLOX-based combustors utilizing various gaseous fuels have shown that the global air-fuel equivalence ratio
for relevant air preheat temperatures [9
]. Therefore, the
was chosen at 2.6 using prevaporized Jet A-1. Using pure hydrogen, the stable operating range is much wider. Recent unpublished experiments of a FLOX-based combustor using pure hydrogen have shown that the air–fuel equivalence ratio at LBO
is >5. In this study, the
was conservatively chosen to 3.6 using pure hydrogen.
The air split S
is a free parameter in the reactor network. In order to select an optimized air split for the Jet A-1 and the hydrogen fuel case, characteristic turbo fan load points were chosen for a trade-off study. As described in Section 2.2.2
the combustion chamber inlet conditions were taken from the turbo fan performance simulations (see Section 2.3
). For the trade-off study, five characteristic load points were chosen from the fuel specific performance decks: take-off, end of field, second segment, top of climb, and mid cruise. For these load points NOx
emissions were traded off against combustion stability. As a marker for combustion stability, the calculated CO emissions were used. As a result of this trade-off study, the air split S
was fixed for the detailed emission study. On the basis of this study, the air split at hydrogen fuelled combustion was set to 87% and to 65% for Jet A-1 furled combustion.
In the detailed emission calculations, 86 load points along the mission flight trajectory were selected from each fuel specific engine performance map (hydrogen and Jet A-1). The reactor network simulations were carried out for these load points at a fixed air split. For each of these load points, the emissions (NOx, CO, H2O, and CO2) were simulated for both hydrogen and Jet A-1. The results were provided as a combustion emission map to the climate impact analysis.
With the reactor network, the effect of air–fuel equivalence ratio in the combustor
on the total NOx
formation was investigated for the two fuel scenarios. Figure 7
shows an increase in the NOx
emission indices EI with decreasing
for both fuel options for a constant flight Mach number of 0.8. This trend is related to the formation of thermal NO, which is pronounced by increasing adiabatic flame temperatures with the decrease in
. For identical
, the adiabatic flame temperature of hydrogen flames is higher than the ones for Jet A-1. Nevertheless, due to the reactivity of hydrogen, hydrogen combustion allows a stable combustion at leaner mixing conditions compared to Jet A-1. Therefore, the formation of NOx
is similar, or can even be reduced by substituting kerosene with hydrogen, as shown in Figure 7
. This is achieved by choosing leaner conditions at the primary combustion zone in the hydrogen case and shifting the air split Sd
towards more secondary air. This choice was made well inside the stable combustion regime. Figure 7
also shows a weak dependency of the flight level on the NOx
formation. The increase in the NOx
formation is also caused by increasing combustor inlet temperatures for increased engine ratings, leading to slightly higher thermal NO productions.
3.3. Aircraft Design
The results at aircraft level are presented in the following section. In total, there are five major concepts. First, the reference aircraft D261 based on the Boeing 767-300 with conventional combustion chambers and fossil Jet A-1. Second, there is the new design D261+, which is an evolutionary advancement of the D261 based on the entry into service (EIS) of 2040. This incorporates an advanced high-bypass-ratio turbofan engine and high aspect ratio carbon fibre reinforced wings. Furthermore, this concept and all further hydro-carbon fuel based ones are evaluated twice, first equipped with an conventional combustion chamber and second with new low-NOx
and low-soot technology. Figure 8
shows the increased span and the reduced wing and horizontal tailplane area.
The next two concepts D261+DropIn and D261+SAF incorporate the same technology bricks but are operated and designed for the drop-in capable fuel and the synthetic one, respectively. The differences in mass, geometry, and performance are rather small. Last, the liquid hydrogen powered aircraft D261+H2 is designed with the same technology assumptions. The big LH2 tanks are integrated in front and behind the cabin, see Figure 9
. The access of the pilots to the cabin is not considered, as this would penalize the aircraft’s performance and there is no official regulation. However, the drawback is the additional lavatory and galley inside the cockpit, which adds weight and needs to be serviced. Furthermore, two additional exits or one exit and one top hatch are needed, as the first two doors of the cabin, which are considered to be in the flight crew area, are not accessible, see CS 25.807 [61
]. However, having a long tunnel between cockpit and cabin should result in the same problem. To better maintain a reasonable fines ratio of the fuselage, the diameter has been increased to accommodate an eight-abreast cabin arrangement. The insulation architecture of the cryogenic tanks is foam with a carbon fibre load carrying inner wall and an kevlar outer wall for protection. This reduces the maximum tank volume and the total vented gaseous hydrogen. The maximum operational pressure is set to 0.25 MPa and additional 5% inner tank volume is required for safety aspects. The delivery of the fuel to the consumer is conducted in the liquid state. The high pressure pumps and the heat exchanger for evaporation are installed near the engines. The total mass of the tank structure, delivery lines, and tank subsystems is roughly 9 tonnes, which leads to a total gravimetric index of about 50%. In addition, the maximum operating altitude and the cruise Mach number are traded for the D261+SAF and D261+H2 concepts to reduce the climate impact of the altitude dependent non-CO2
contributions. The maximum altitude constraint is varied between flight level 290 and 410 whereas the cruise Mach number varies between 0.75 and 0.8. For each of these points, a new design is conducted incorporating new sweep angles and high-lift performance. Figure 8
shows the generic top view of all concepts. All new concepts have an increased wing span of 52 m compared with the reference D261. The fuselage width and length of the LH2 aircraft design is increased due to the integration of the big storage tanks which is also visible in Figure 8
The masses of all concepts for the MTOM and maximum payload case are shown in Figure 10
. The masses of all new kerosene based concepts are strongly reduced compared to the reference D261. This is due to the new materials and manufacturing methods as well as the increased efficiency and the correlated lower fuel mass and structural snowball effects. Additionally, the maximum payload has been reduced for all new concepts. Interestingly, the MTOM of the D261+H2 decreases whereas the operational empty mass (OEM) increases. Especially the increased maximum landing mass (MLM) and the decreased MTOM explains the relaxed low-speed thrust requirements, as the wing area is sized for MLM conditions. The structural LH2 storage mass as well as the subsystems and delivery lines are accounted as systems which is why this portion strongly increase in Figure 10
shows several major results at aircraft level. It can be seen that the reference area decreases for the 2040 kerosene concepts and strongly increases again for the LH2 one due to the increased MLM. The thrust loading for take-off and top of climb (TOC) conditions based on an equivalent thrust at sea level (SL) and MTOM shows that the high speed requirements are more sizing for the LH2 concept. That means that the optimal flight altitude is more a trade between power installed and aerodynamic performance, which is why LH2 aircraft concepts tend to fly at slightly lower altitudes than the aerodynamic optimum. Another advantage is that the engine can be down-rated for low speed conditions, which has a positive effect at maintenance and allows shorter take-off field lengths (TOFL). The explained reduced difference between MTOM and MLM is also the reason why the wing loading based on MTOM is lower for the D261+H2 than for the 2040 kerosene concepts. The lift to drag ratio increases for the new 2040 kerosene concepts compared to the reference, mainly due to the increased span and reduced averaged cruise mass, or in other words due to the reduced lift per wing span. The aerodynamic performance decreases again for the D261+H2 due to the wider and longer fuselage to accommodate the LH2 storage tanks. The maximum high lift coefficient CL
is slightly smaller for the D261+H2 as the wider fuselage reduces the useful wing area to install high lift devices. The block energy of the new kerosene based concepts for 2040 decreases by roughly 39%, which mainly results from the increased overall propulsion efficiency. The additional structural mass, together with the reduced aerodynamic efficiency, shown in Table 5
in terms of lift to drag ration, increases the block-energy of the D261+H2. However, the increased engine efficiency of roughly 7%, explained in Section 3.2
, dampens these effects, leading to only an additional 3% block-energy at design range and payload compared with the D261+SAF concept. For smaller ranges, this difference increases due to the higher structural mass fraction of the D261+H2. For a 800 nm mission, the difference is already 6%.
The D261+SAF and the D261+H2 concepts are varied in maximum flight level constraints and cruise Mach number. For each flight level, a new concept is designed to meet the same requirements. The following Figure 11
, Figure 12
, Figure 13
and Figure 14
show vehicle specific dependencies of the the cruise Mach number and the maximum allowable flight level variation for a representative 1500 NM mission. In general, it can be noted that the lower the flight level constraint for one cruise Mach number, the worse the overall performance. The main reason for this is the overall propulsion performance and the aerodynamic behaviour. The Lift-over-Drag ratio (LoD) decreases with decreasing flight levels as the aircraft is forced to fly at lower than optimal lift coefficients. As the first order effects for the aerodynamic optimum flight level are wing loading and speed, this effect can be mitigated by also reducing the flight level, as seen in Figure 12
. The wing loading in cruise can be influenced by the high lift capabilities in landing configuration. For high maximum lift coefficients, the wing can be designed smaller leading to a higher overall wing loading. However, the maximum fuel capacity for the kerosene concepts has to be considered and becomes limiting for more aggressive high lift coefficients. The overall propulsion efficiency decreases with decreasing flight level constraints as the thrust matching between low speed and high speed requirements is not optimal. The low speed is much more sizeable which leads to an off design operation in cruise. While the engine is operated at about 83% of the maximum thrust at cruise rating for the FL 410 and 0.8 Mach number case, allowing barely the required rate of climb of 300 ft/min, the FL 290 and 0.8 Mach number case is operated at around 54%. However, the turbofan is designed for the flight level 410 and Mach 0.8 case and is scaled for the other concepts with different constraint settings. Therefore, the effects shown in Figure 13
can potentially be weakened. Another promising solution could be to boost the turbofan in low speed conditions and design it especially for the specific cruise requirements. The Block-Energy (BE) is shown in Figure 14
. Both concepts show rather similar behaviour. By simply reducing the flight level, the aircraft is forced to fly at non ideal aerodynamic and propulsion conditions leading to strong block energy increases of more than 20%. This can partially be compensated by also reducing the cruise Mach number.
3.4. Global Emissions and Climate Impact Results
The overall results in terms of emissions and climate impact of the different concepts evaluated at global fleet level are shown in Figure 15
and Figure 16
. Furthermore, the global distribution of emissions is visualized in Appendix A
. The grey cells in Figure 15
and Figure 16
indicate the changes per step. The reference aircraft D261 is used to normalize the global emissions and the Average Temperature Response (ATR). The ATR is calculated for a time frame of 100 years and 32 years of operation, starting in 2040. For the D261, the impact of CiC is responsible for 53% and the total NOx
impact, consisting of increased global ozone (O3
) and increased depletion rates of methane (CH4
) for 23%, the carbon dioxide for 19%, and water vapour for 5%. The D261+ with conventional technology advancements for 2040 shows benefits in CO2
and water vapour which correlates with the improved fuel consumption. However, the increased temperatures in the conventional combustion chamber cause higher NOx
emissions, which compensates the benefit of reduced fuel consumption. Additionally, the contrail impact could potentially increase due to the correlation with the overall propulsion efficiency of the turbofan engines. Nevertheless, this has to be interpreted with care as this correlation is based just on few reference points (see Section 2.6
). Despite the strongly reduced fuel consumption, the potential climate impact reduction is not significant and could even be worse, indicated by the uncertainties. For the advanced combustion chamber, the climate impact is reduced by 65%. In addition to the benefits in CO2
O, the advanced combustion chamber (low NOx
and low soot) reduces the climate impact of CiC and NOx
due the reduced NOx
and particles emission of 91% and 90%, respectively.
The next two concepts, D261+DropIn and D261+SAF in Figure 15
and Figure 16
, have the same technology level as the D261+ but are powered with drop-in capable kerosene, considering today’s standards, and with 100% synthetic kerosene produced by PtL. Whereas the redesign has little effect at vehicle performance level, the change in climate impact is much higher. The CO2
emissions and impact is strongly reduced for the D261+DropIn and vanishes for the D261+SAF. This is due to the assumption of carbon neutral fuel production incorporating Direct Air Capturing (DAC) as carbon source. For the conventional combustion chamber, the contrail impact is lowered due to the reduced soot emissions of the drop-in and PtL fuel, see Table 1
. The gradual reduction for the D261+DropIn and D261+SAF with advanced combustion chambers results not from the reduced contrail effect, as the soot emissions are already at the lowest provable level where the particles in the environment are dominating the properties of contrails, but from the CO2
neutral production with DAC. All concepts described thus far are operated at performance optimized flight levels. The next concept, displayed in Figure 15
and Figure 16
, reduces the maximum flight level as well as the cruise Mach number to 290 and 0.75, respectively. This results in 5% higher direct operating costs (DOC) compared to the D261+SAF at optimum flight levels, mainly caused by higher crew costs, capital expenditures, and fuel consumption. As the CO2
impact is eliminated within this scenario, the reduced aircraft performance is merely visible. However, the altitude reduction strongly affects the non-CO2
effects, which leads to a reduction of 56% compared to the non restricted case and 69% compared to the reference scenario with conventional combustion. For the advanced combustion chamber, the reduction compared to the reference reaches 93%.
The uncertainties of all low NOx
and low soot combustion are rather high, as the effect of reduced soot emissions at contrail impact is not yet understood profoundly. For conventional combustors, a major part of volatile particle (vPM) emissions condenses on non volatile particle (nvPM) emission. This effect and its impact on contrail impact is well understood. In case of a low NOx
and low soot combustor, the level and role of vPM on contrail formation has not been examined and is not clear. This is reflected by an additional uncertainty of the low NOx
and low soot combustor model results in Figure 16
The last two columns in Figure 15
and Figure 16
are LH2 concepts, the first with no flight level restrictions and the second with the flight level restriction of 290 and a cruise Mach number of 0.75, which causes the same 5% increase in DOC as in the D261+SAF case. The total NOx
emissions and impact vanishes due to the advanced combustion, described in Section 3.1
. Due to the strongly increasing water vapour emissions for LH2 concepts, the water vapour climate impact also increases by 40% compared with the unrestricted D261+SAF. However, due to the low overall impact of water vapour, this strong increase merely affects the overall impact. Furthermore, it can almost be eliminated for the restricted D261+H2 case, shown in the last columns of Figure 16
. The LH2 contrail uncertainties are relatively high as the reference simulations for the climate impact correlation for LH2 are to few to allow a more specific conclusion. However, this confidence band shows the potential range and allows a first comparison with synthetic kerosene concepts.
Summarized, the unrestricted synthetic kerosene with conventional combustion and LH2 concepts result in a 31% and 75% climate impact reduction compared to the reference, respectively. The flight level restricted concepts achieve a 69% and 85% reduction for the synthetic kerosene with conventional combustion and LH2 scenarios, respectively. The four kerosene concepts for 2040 with advanced combustion show a significant reduction in total NOx as well as contrail impact. This leads to a reduction of 65% for the D261+ and 77% for the D261+SAF. The flight level restricted scenario even leads to a 93% reduction in the climate impact compared to the reference.